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Pilot Training ManualPilot Training Manual

King Air 200

Citation XLS

Revision 0

Revision 0

Printed in the United Stated of America

King Air 200 aircraft materials used in this publicationhave been reproduced with permission of

Raytheon Aircraft Corporation.

Copyright © 2011, CAE, Inc.All rights reserved.

NOTICE: This King Air 200 Pilot Training Manual is to be used for aircraft familiarization and training purposes only. It is not to be used as, nor considered a substitute for, the manu-facturer’s Pilot or Maintenance Manual.

Introduction

King Air 200 1-1December 2011

For Training Purposes Only

Welcome to CAE

Welcome to CAE!

Our goal is a basic one: to enhance your safety, proficiency and professionalism within the aviation community. All of us at CAE know that the success of our company depends upon our commitment to your needs. We strive for excellence by focusing on our service to you.

We urge you to participate actively in all training activities. Through your involvement, interaction, and practice, the full value of your training will be transferred to the operational environment. As you apply the techniques presented through CAE training, they will become “second nature” to you.

Thank you for choosing CAE. We recognize that you have a choice of training sources. We trust you will find us committed to providing responsive, service-oriented training of the highest quality.

Our best wishes are with you for a most successful and rewarding training experience.

The Staff of CAE!

King Air 2001-2December 2011


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Introduction



Using this ManualThis manual is a stand-alone document appropriate for various levels of training. Its purpose is to serve as an informational resource and study aid.

The Quick Reference chapter provides limitations, memory items from procedural checklists, and other data for quick review.The Operating Procedures chapter contains sub-chapters that provide a pictorial preflight inspection of the aircraft, normal procedures in an expanded format, standard operating procedures, maneuvers, and other information for day-to-day operations.The Flight Planning chapter covers weight and balance and performance; a sample problem is included.The Systems section is subdivided by aircraft system. Each system chapter contains a discussion of components, preflight and servicing procedures, and abnormal and emergency procedures. At the beginning of the Systems chapter, a list of systems is cross referenced to ATA codes to facilitate further self study, if desired, with the manufacturer’s manuals.Three graphics are used in this manual to direct your attention to a specific location in the cockpit. A shaded area locates various instruments or switches shown in the adjacent photographs.

The graphic at right (top) represents the overhead panel.

The graphic at right (middle) represents the cockpit forward panels.

The graphic at right (below) represents the center pedestal and side panels.




Quick Reference 2



ContentsQuick ReferenceGeneral Limitations

Airstair Door ...............................................................................................2-7Authorized Operations ..............................................................................2-7Baggage Limits ..........................................................................................2-7Certification Status ....................................................................................2-8Cargo ..........................................................................................................2-8Cargo Door .................................................................................................2-8Emergency Exit ..........................................................................................2-8Occupancy Limits ......................................................................................2-8Passenger Seating ....................................................................................2-9

Couch/Passenger Seats ........................................................................2-9Aft-Facing Seats ....................................................................................2-9Lateral-Tracking Seats (if installed) .......................................................2-9Passenger Shoulder Harness ................................................................2-9

Maneuvers ..................................................................................................2-9Minimum Flight Crew ................................................................................2-9Structural Limitations ...............................................................................2-9

King Air 200 Series ................................................................................2-9Windows and Windshield ........................................................................2-10

Fuselage Side Window .........................................................................2-10Crack in Side Window or Windshield ....................................................2-10

Operational LimitsCabin Pressurization Limit ......................................................................2-11Cabin Pressurization Controller ..............................................................2-11Crosswind/Tailwind Components ...........................................................2-11Generator Limits .......................................................................................2-12Starter Limitations ....................................................................................2-13OAT Limits .................................................................................................2-13Airspeed Limitations ................................................................................2-14All Models (at 12,500 lbs) .........................................................................2-16Static Wicks (King Air 200) ......................................................................2-17

Towing ..................................................................................................2-17

King Air 2002-2 For Training Purposes OnlyDecember 2011

Weight Limitations. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-17Center of Gravity Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-17Flight Load Factor Limits (200) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-17Mean Aerodynamic Chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-18Flight in Icing Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-18

Systems LimitationsAutopilot – King Air 200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-19

FAR Part 91 Operations ....................................................................2-19FAR Part 135 Operations ..................................................................2-19

Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-19Approved Fuel Anti-Icing Additive .....................................................2-19Fuel Limits ........................................................................................2-19King Air B200 Series .........................................................................2-20

Fuel Biocide Additive . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-20Approved Engine Fuels ....................................................................2-21Emergency Engine Fuels ..................................................................2-21Limitations on the Use of Aviation Gasoline .....................................2-21Auxiliary Fuel ....................................................................................2-21Fuel Crossfeed .................................................................................2-21Fuel Gauges in the Yellow Arc ..........................................................2-22Fuel Imbalance Between Wings .......................................................2-22Operating with Low Fuel Pressure ....................................................2-22Ice and Rain Protection Systems ......................................................2-22Optional Brake Deice System ...........................................................2-22Pneumatic Deice Boots ....................................................................2-22Ice Vanes (Inertial Separator System) ..............................................2-23Powerplant ........................................................................................2-23Adjustments ......................................................................................2-24Adjustments ......................................................................................2-25Adjustments ......................................................................................2-26Adjustments ......................................................................................2-27

Engine Operating Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-28Oil Specifications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-28Oxygen . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-31Propellers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2-31

King Air 200/B200 .............................................................................2-31Instrument Markings

Cabin Pressure Differential Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . 2-34King Air 200 – before BB-195 ...........................................................2-34King Air 200 – BB-195 and subsequent; BL-1 and subsequent .......2-34

Quick Reference



King Air B200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-34Pneumatic Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-34Fuel Quantity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-35Vacuum/Gyro Suction Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-35King Air 200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-35

King Air 200 (alternate gauge) and King Air B200 . . . . . . . . . . . . . . .2-35Propeller Deice Ammeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-36

System Data SummariesElectrical Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-37

DC System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-37AC System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-38

Fire Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-38Flight Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-39

Flap System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-39Fuel Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-39

Main Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-39Auxiliary Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-40Ice and Rain Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-40

Surface Deice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-40Prop Heat System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-40Anti-Ice Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-41

Brake Deice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-41Pitot Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-41Stall Warning Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-41Fuel Vent Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-41Windshield Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-41Ice Vanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-41

Landing Gear Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-42Electro-Mechanical Landing Gear System . . . . . . . . . . . . . . . . . . . .2-42Hydraulic Landing Gear System . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-42Brake System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-43

Pneumatic Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-43Bleed Air System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-43

Air Conditioning/Heating System . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-44Pressurization System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-44

Powerplant System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2-45



Quick Reference



Quick ReferenceThis chapter contains the aircraft’s operating limits and requirements as well as system-by-system charts summarizing power sources, distribution, and monitors. All limitations are printed in gray.This chapter serves as a convenient reference.



Quick Reference



General Limitations

Airstair Door

WRRIRG Only a crew member should operate the door.

Do not open or check security by moving door handle while aircraft is pressurized and/or in flight.Handle is in locked position when arm is around plunger.

See AFM Supplements section for limitations with the airstair door removed.

NTTE: Only one person should be on the airstair door stairway at any one time (300 lbs maximum).

Authorized Operations

WRRIRG No aircraft is certified for flight into known severe icing.

Day and Night VFR Day and Night IFR Known icing conditions FAR Part 91 operations when all pertinent information and

performance considerations are complied with. FAR Part 135 operations when all pertinent information and

performance considerations are complied with.

Baggage Limits

WRRIRG Do not carry children in the baggage compartment unless secured in a seat.

Maximum Weight in Baggage CompartmentPrior to BB-1052, BB-1091, and BL-58:With Fold-up Seats . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .370 LBSWithout Fold-up Seats . . . . . . . . . . . . . . . . . . . . . . . . . . . . .410 LBS


BB-1052, BB-1091 and subsequent; BL-58 and subsequent; prior aircraft with Beech Kit #101-5068-1 installed:With Fold-up Seats . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .510 LBSWithout Fold-up Seats . . . . . . . . . . . . . . . . . . . . . . . . . . . . .550 LBS

Certification Status Normal Category, FAR Part 23

Cargo

WRRIRG Unless authorized by applicable Department of Transportation regulations, do not carry hazardous material anywhere in the aircraft.

All cargo shall be properly secured by an FAA-approved cargo restraint system.

Cargo must be arranged to permit free access to all exits and emergency exits.

Cargo Door 200C/B200C – Do not open or check security by moving door handle

while aircraft is pressurized and/or in flight. The cabin flooring section withstands loads of 200 pounds per square

foot supported on the seat tracks. Floor areas where seat tracks are not present (walkways and aft baggage/utility area) supports 100 pounds per square foot loads.

NTTE: Prior to first flight of the day, check cabin door/cargo door annunciator circuitry in accordance with Cabin/Cargo Door Annunciator Circuitry Check in the AFM.

Emergency Exit

WRRIRG Only a crew member should operate the door.

Emergency exit must be unlocked before takeoff.

Occupancy LimitsFAR Part 91 Operations (maximum) . . . . . . . . . . . . . 15 Including CrewFAR Part 135 Operations(maximum) . . . . . . . 9 Passengers Plus Crew

Quick Reference



Passenger Seating

Couch/Passenger Seats Do not occupy couch as chaise lounge during takeoff and landing. Maximum weight of drawer contents is 30 lbs (13.6 kg) per drawer. The headrest should be positioned properly for the occupant.

Aft-Facing Seats Only aft-facing seats (placarded as such on the leg crossmember)

are authorized in the aft-facing position. The seatback of each occupied aft-facing seat must be in the fully

raised position and the headrest in the full-up position for takeoff and landing.

Lateral-Tracking Seats (if installed) Seat must be in outboard position for takeoff and landing.

Passenger Shoulder Harness Shoulder harness must be worn during takeoff and landing with

seat in outboard position, seat back upright, and headrest fully extended.

Maneuvers The Beechcraft Super King Air B200 and B200C are Normal Category

Airplanes. Acrobatic maneuvers, including spins, are prohibited.

Minimum Flight Crew One pilot.

Structural Limitations

King Air 200 SeriesS/Ns BB-2, BB-6 through BB-733, BB-735 through BB-792, BB-794 through BB-828, BB-830 through BB-853, BB-871 through BB-873, BB-892, BB-893, BB-895, BB-912, BL-1 through BL-36Maximum Cabin Pressure Differential . . . . . . . . . . . . . . . . . . . . .6.1 PSICabin Door Forward and Aft Side Latches (or bayonets) (4) Safelife (200 only) . . . . . . . . . . . . . . . . . . . 6,000 HRSCabin Door Upper Latch Hooks (2) and Attaching Hardware (200 only) . . . . . . . . . . . . . . . . . . . . . 12,000 HRS


Cabin Door Cam-Lock Actuator Cable Safelife (200C only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9,000 HRSWing and Associated Structure Fatigue Safelife . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20,000 HRSWindshield Frame Screws . . . . . . . . . . . . . . . . . . . . . . . . . . 12,000 HRSAll Wing Attach Bolts, Nuts, Barrel Nut Assemblies

Steel Components . . . . . REPLACE EVERY 6 CALENDAR YEARS OF INSTALLED BOLT AND NUT TIME

Inconel Components . . REPLACE EVERY 15 CALENDAR YEARS OF INSTALLED BOLT AND NUT TIME

Refer to the Beechcraft Structural Inspection and Repair Manual and the Super King Air 200 Series Maintenance Manual for inspection and replacement procedures.

Windows and WindshieldKing Air 200/King Air B200, BB-1158, BB-1167, BB-1193 and after; B6-73 and after; B7-31 and after; BN-5 and after

Fuselage Side Window If cracking, chipping, or stress crazing that can be felt with a fingernail

occurs in either ply of the exterior window, replace the window according to instructions in the Maintenance Manual.

If the window cannot be replaced prior to the next flight, pressurized flight is prohibited. Install the following placards to conduct unpressurized flight. – Install the following placard in clear view of the pilot:

PRESSURIZED FLIGHT IS PROHIBITED DUE TO A DAMAGED WINDOW. CONDUCT FLIGHT WITH THE CABIN PRESS SWITCH IN THE DUMP POSITION.

– Install the following placard next to the pressurization control:UNPRESSURIZED FLIGHT ONLY PERMITTED.

If a crack exists in both the inner and outer plies of the exterior window, replace the window prior to further flight unless an appropriate “Ferry Permit” is obtained through the proper authority.

Crack in Side Window or WindshieldIf it has been determined that a crack has developed in any side window or windshield: maintain altitude at 25,000 ft or less reset pressurization controller:

– windshield – 4.0 PSI – side window – minimum practical

Quick Reference



Operational LimitsMaximum Operating Pressure Altitude

Normal Operations – King Air 200 Prior to BB-54, except 38, 39, 44 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31,000 FTNormal Operations – King Air 200 BB-38, 42, 44, 54 and subsequent*; BL-1 and subsequent . . . . . . . . . . 35,000 FT

*Also includes earlier aircraft with Beech kit Nos. 101-5007-1 and 101-5008-1 in compliance with Beechcraft Service Instruction No. 0776-341.

Normal Operations – King Air B200 . . . . . . . . . . . . . . . . . . . 35,000 FTKing Air 200 with Aviation Gasoline:

Both Standby Boost Pumps Operative . . . . . . . . . . . . . . 31,000 FTEither Standby Boost Pump Inoperative . . . . . . . . . . . . . 20,000 FTClimbs without Crossfeed Capability . . . . . . . . . . . . . . . . 20,000 FT

Yaw Damper System Inoperative . . . . . . . . . . . . . . . . . . . . . 17,000 FTNo Restriction with aft strakes installed.

VMCA Demonstration Minimum . . . . . . . . . . . . . . . . . . . . . .5,000 FT AGL

Cabin Pressurization LimitKing Air 200 – Maximum CabinPressure Differential . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .6.1 PSIKing Air B200 – Maximum CabinPressure Differential . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .6.6 PSI

Cabin Pressurization Controller Depressurize cabin before landing. Refer to the Operating Handbook

chart to determine the correct pressure altitude setting.

Crosswind/Tailwind ComponentsCrosswind (maximum demonstrated) . . . . . . . . . . . . . . . . . . . . . .25 KtsTailwind (takeoff/landing [maximum charted]) . . . . . . . . . . . . . . . .10 Kts

External Power Unit 28 to 28.5 VDC output 400A continuous 1,000A surge

All B200s and subsequent 28 to 28.4 VDC output 300A continuous 1,000A surge


Generator LimitsMaximum sustained generator load is limited as follows:

In Flight – Sea Level to 31,000 ft altitude . . . . . . . . . . . . 1.00/100%In Flight – Above 31,000 ft altitude . . . . . . . . . . . . . . . . . . 0.88/88%On Ground . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.85/85%

During ground operation, observe the limitations shown in Tables 2-1, 2-2, and 2-3.

Generator Load Minimum Gas Generator RPM – N1%Without A/C With A/C

(right engine only)King Air 200

0 to 0.70 0.70 to 0.75 0.75 to 0.80 0.80 to 0.85

52 55 60 65

60 60 60 65

King Air 2000 to 0.75 0.75 to 0.80 0.80 to 0.85

56 60 65

62 62 65

BB1439 and subsequent0 to 0.75 0.75 to 0.80 0.80 to 0.85

61 61 65

62 62 65

Table 2-1: King Air 200 and B200 Generator Limits

Generator Load Minimum Gas Generator RPM N1%Without A/C With A/C

(right engine only)In Flight (all altitudes)

0 to 0.75 IDLE 62 IDLE0.75 to 1.00 63 68 85

Table 2-2: 300A Lear-Siegler Starter-Generator 23085-001 Limits

Type of Operation Minimum N1% RPM Max Generator Load % of LoadGround (from sea level to 5,000 ft1) 52

55 65 70

50 66 90

100Flight2 75 100

Table 2-3: 300A Lear-Siegler Starter-Generator 23069-016 Limits1 Sea level to 5,000 ft., observe engine ITT limits when operating at low N1. Decrease high ITT by

reducing accessory load and/or increasing N1 speed.2 This flight operation is for airspeeds of 116 KIAS and higher. Observe engine ITT limits.

Quick Reference



Starter Limitations Standard Start Cycle:

– 40 seconds ON/60 seconds OFF – 40 seconds ON/60 seconds OFF – 40 seconds ON then 30 minutes OFF

300-Amp Lear-Siegler (Optional) Standard Start Cycle:

– 30 seconds ON/3 minutes OFF – 30 seconds ON/30 minutes OFF

For engine wash: – 30 seconds ON/15 minutes OFF

For engine soak: – 30 seconds ON/10 minutes OFF – 30 seconds ON/10 minutes OFF – 30 seconds ON/30 minutes OFF

OAT LimitsSea Level to 25,000 ft Pressure Altitude . . . . . . . . . . . . MAX ISA +37°CAbove 25,000 ft Pressure Altitude . . . . . . . . . . . . . . . . MAX ISA + 31°CAll Altitudes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .MIN -53.9°C


Airspeed LimitationsAirspeed limitations are detailed in Tables 2-4 and 2-5.

SPEED KCAS KIAS REMARKSVA Maneuvering Speed (12,500 lbs)

182 181 Do not make full or abrupt control move ments above this speed.

VFE, Maximum FlapExtension/Extended:

Approach – 40% Full Down – 100%

200 144

200 146

Do not extend flaps or operate with flaps in prescribed position above these speeds.

VLO, MaximumLanding Gear Operating:

Extension Retraction

182 164

181 163

Do not extend or retract landing gear above the given speed.

VLE, Maximum Landing Gear Extended

182 181 Do not exceed this speed with the landing gear extended.

VMCA, Minimum Control Airspeed

91 86 This is the lowest airspeed at which the aircraft is directionally controllable with one engine at takeoff power when the other engine suddenly becomes inoperative with propeller windmilling.

VMO Maximum Operating:BB-2 to 198 without Beech Kit 101-5033-1MMO, Maximum Operating:

270

0.48 M

269

Do not exceed these airspeeds or Mach numbers in any operation.

VMO Maximum Operating:BB-199 and subs., BL-1 and subs.; prior S/Ns with Beech Kit 101-5033-1MMO, Maximum Operating:

260

0.52 M

259

Do not exceed these airspeeds or Mach numbers in any operation.

Table 2-4: King Air 200 Speed Limitations

Quick Reference



SPEED KCAS KIAS REMARKSVA Maneuvering Speed (12,500 lbs)

182 181 Do not make full or abrupt control move ments above this speed.

VFE, Maximum FlapExtension/Extended:

Approach – 40% Full Down – 100%

200 155

200 157

Do not extend flaps or operate with flaps in prescribed position above these speeds.

VLO, MaximumLanding Gear Operating:

Extension Retraction

182 164

181 163

Do not extend or retract landing gear above the given speed.

VLE, Maximum Landing Gear Extended

182 181 Do not exceed this speed with the landing gear extended.

VMCA, Minimum Control AirspeedHartzell PropellersMcCauley Propellers

9291

8686

This is the lowest airspeed at which the aircraft is directionally controllable with one engine at takeoff power when the other engine suddenly becomes inoperative with propeller windmilling.

VMO Maximum Operating:

MMO, Maximum Operating:

260

0.52 M

259 Do not exceed these airspeeds or Mach numbers in any operation.

Table 2-5: King Air B200/B200C Speed Limitations1 BB-2, BB-6 through BB-733, BB-735 through BB-792, BB-794 through BB-828, BB-830, etc.,

(200s). Airspeed indicator marked in CAS values.

2 B200 indicators marked in IAS values.


All Models (at 12,500 lbs)VMCG, Ground Minimum Control . . . . . . . . . . . . . . . . . . . . . . . . . 84 KIASVMCA, Air Minimum Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86 KIASTakeoff (Flaps 0%/Flaps 40%):

V1/VR(rotation) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95 KIAS/94 KIAS50 ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 121 KIAS/106 KIASV2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 121 KIAS/106 KIAS

VSSE, Intentional One-Engine Inoperative . . . . . . . . . . . . . . . . 104 KIASVY, Two-Engine Best Rate-of-Climb . . . . . . . . . . . . . . . . . . . . . 125 KIASVYSE, One-Engine Inoperative Best Rate-of-Climb. . . . . . . . . . 121 KIASVX, Two-Engine Best of Angle-of-Climb . . . . . . . . . . . . . . . . . . 100 KIASVXSE, One-Engine Inoperative Best Angle-of-Climb . . . . . . . . . 115 KIASMaximum Glide Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135 KIASTurbulent Air Penetration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 170 KIASBalked Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100 KIASCruise Climb:

Sea Level to 10,000 ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160 KIAS10,000 to 20,000 ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140 KIAS20,000 to 25,000 ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 130 KIAS25,000 to 35,000 ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 120 KIAS

Landing Approach:Flaps 100% . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103 KIASFlaps 0% . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132 KIAS

Icing Conditions (minimum) . . . . . . . . . . . . . . . . . . . . . . . . . . . 140 KIASEffective Windshield Deicing (maximum) . . . . . . . . . . . . . . . . 226 KIASEmergency Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 181 KIASStall Speeds – Power Idle, 0° Angle-of-Bank:

100% Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .75 KIAS (VSO)40% Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85 KIAS0% Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99 KIAS (VS1)

Airstart (minimum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140 KIASAutopilot Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . VMO/MMO

Flight with Cabin Entrance Door Removed . . . . . . . . . . . . . . . 205 KIAS

NTTE: Exceeding the nosewheel deflection limit markings during towing operations damages the nose strut/linkage. Nosewheel deflection of approximately 10° or more with the rudder gust lock installed damages the nosewheel steering linkage.

Quick Reference



Static Wicks (King Air 200) One wick may be missing or broken from:

1. Each wing (includes aileron)2. Each side of horizontal stabilizer3. Vertical stabilizer

Maximum of three wicks may be missing. Due to varying configurations, consult your MEL for number and

position of static wicks.

Towing Do not tow the aircraft with rudder gust lock installed. Do not tow the aircraft if one or more landing gear struts are

deflated.

Weight LimitationsMaximum Ramp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,590 LBSMaximum TakeoffAll Except FAR Part 135 Operations . . . . . . . 12,500 LBS

FAR Part 135 Operations . . . . . . SEE MAXIMUM ENROUTE WEIGHT CHART IN AFM

Maximum Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12,500 LBSMaximum Zero Fuel:

King Air 200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10,400 LBSKing Air B200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11,000 LBS

Center of Gravity Limits The reference datum is 83.5 inches (212.09 cm) forward of the

center of the front jack point. Aft limit – 196.4 inches (498.8 cm) aft of datum at all weights. Forward at 12,500 lbs (5,669.9 kg) – 185.0 inches (469.9 cm) aft of

datum with straight line variation to 181.0 inches (459.7 cm) aft of datum at 11,279 lbs (5,116.06 kg).

Forward at 11,279 lbs (5,116.06 kg) or less – 181.0 inches (459.7 cm) aft of datum.

Flight Load Factor Limits (200)Flaps Up . . . . . . . . . . . . . . . . . . .3.17 POSITIVE Gs

1.27 NEGATIVE Gs Flaps Down . . . . . . . . . . . . . . . .2.00 POSITIVE Gs 1.27 NEGATIVE Gs, (0.0 G [B200])


Mean Aerodynamic ChordMAC Length. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70.41 INCHESLeading Edge of MAC . . . . . . . . . . . . 171.23 INCHES AFT OF DATUM

Flight in Icing ConditionsSevere icing may result from environmental conditions outside of those for which the airplane is certificated. Flight in freezing rain, freezing drizzle, or mixed icing conditions (supercooled liquid water and ice crystals) may result in ice build-up on protected surfaces exceeding the capability of the ice protection system, or may result in ice forming aft of the protected surfaces. This ice may not be shed using the ice protection systems, and may seriously degrade the performance and controllability of the airplane. During flight, severe icing conditions that exceed those for which the

airplane is certificated shall be determined by the following visual cues. If one or more of these visual cues exists, immediately request priority handling from Air Traffic Control to facilitate a route or an altitude change to exit the icing conditions. – Unusually extensive ice accumulation on the airframe and

windshield in areas not normally observed to collect ice. – Accumulation of ice on the upper surface of the wing, aft of the

protected area. – Accumulation of ice on the engine nacelles and propeller spinners

farther aft than normally observed. Since the autopilot, when installed and operating, may mask tactile

cues that indicate adverse changes in handling characteristics, use of the autopilot is prohibited when any of the visual cues specified above exist, or when unusual lateral trim requirements or autopilot trim warnings are encountered while the airplane is in icing conditions.

All wing icing inspection lights must be operative prior to flight into known or forecast icing conditions at night.

Quick Reference



Systems Limitations

Autopilot – King Air 200

FAR Part 91 Nperations Refer to the applicable FAA Approved Flight Manual Supplement in

the AFM Supplements Section.

FAR Part 135 Nperations Refer to the applicable FAA Approved Flight Manual Supplement in

the AFM Supplements section for your particular autopilot installation except for Minimum Altitude, which is established by FAR Part 135.93.

Fuel System

Approved Fuel Anti-Icing Additive

WAUIOR Anti-icing additive must be properly blended with the fuel to avoid deterioration of the fuel cells. The additive concentration by volume shall be a minimum of 0.06% and a maximum of 0.15%. Approved procedure for adding anti-icing concentrate is contained in AFM Section IV, Normal Procedures.

Use anti-icing additive conforming to Specification MIL-I-27686.

Fuel LimitsKing Air 200 SeriesMinimum Temperature LimitEngine oil is used to heat the fuel on entering the fuel control. Since no temperature measurement is available for the fuel at this point, it must be assumed to be the same as the OAT. Operations with Commercial Grade fuels are prohibited below the OAT indicated below unless approved anti-icing fuel additives are used. Military Grade fuels have anti-icing additives blended in the fuel at the refinery, and no further treatment is necessary. Operations with Military Grade fuels below the temperatures indicated are prohibited.A minimum oil temperature of 55°C is recommended for fuel heater operation at takeoff power.


2-1

King Air B200 Series

COMMERCIAL GRADES MILITARY GRADESJET A: -40°C JP-4: -58°C

JET A-1: -47°C JP-5: -46°C JET B: -50°C JP-8: -50°C

Fuel Biocide AdditiveFuel biocide-fungicide BIOBOR JF in concentrations of 135 PPM or 270 PPM may be used in the fuel. BIOBOR JF may be used as the only fuel additive, or it may be used with the anti-icing additive conforming to MIL-I-27686 specification. Used together, the additives have no detrimental effect on the fuel system components.Refer to the King Air 200 or Super King Air 200 Series Maintenance Manual and to the latest Pratt and Whitney Canada Engine S/N.3044 for concentrations to use and for procedures, recommendations, and limitations pertaining to the use of biocidal/ fungicidal additives in turbine fuels.

Quick Reference



Approved Tngine Fuels

WAUIOR JP-4 fuel per MILT-5624 has anti-icing additive per MIL-I-27686 blended at the refinery, and no further treatment is necessary. Some fuel suppliers blend anti-icing additive in their storage tanks. Prior to refueling, check with the fuel supplier to determine whether or not the fuel has been blended. To assure proper concentration by volume of fuel on board, blend only enough additive for the unblended fuel.

Commercial Grades. . . . . . . . . . . . . . . . . . . . . . . . . Jet A, Jet A-1, Jet BMilitary Grades . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .JP-4, JP-5, JP-8

Tmergency Tngine FuelsCommercial Aviation Gasoline Grades . . . . . . . . . . . . . 80 RED, 91/98,

100LL*, 100 GREEN, 115/145 PURPLE

Military Aviation Gasoline Grades . . . . . . . . . . . . . . . . . . . . 80/87 RED,100/130 GREEN, 115/145 PURPLE

*In some countries, this fuel is colored green and designated “100L.”

Limitations on the Use of Aviation Gasoline

WRRIRG JP-4 fuel per MILT-5624 has anti-icing additive per MIL-I-27686 blended at the refinery, and no further treatment is necessary. Some fuel suppliers blend anti-icing additive in their storage tanks. Prior to refueling, check with the fuel supplier to determine whether or not the fuel has been blended. To assure proper concentration by volume of fuel on board, blend only enough additive for the unblended fuel.

Operation is limited to 150 hours between engine overhauls. Operation is limited to 20,000 ft pressure altitude (FL 200) or below

if either standby pump is inoperative. Crossfeed capability is required for climbs above 20,000 ft pressure

altitude (FL 200). Operation above 31,000 ft (FL 310) is prohibited.

Auxiliary Fuel Do not put any fuel into the auxiliary tanks unless the main tanks are full.

Fuel Crossfeed Crossfeeding of fuel is permitted only when one engine is inoperative.


Fuel Gauges in the Yellow Arc Do not takeoff if fuel quantity gauges indicate in the yellow arc or

indicate less than 265 lbs (120.2 kg) of fuel in each main tank system.

Fuel Imbalance Between WingsMaximum Allowable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1,000 LBS

Nperating with Low Fuel Pressure Operation of either engine with its corresponding fuel pressure

(L/R FUEL PRESS annunciator) illuminated is limited to 10 hours before overhaul or replacement of the engine-driven fuel pump. Windmilling time need not be charged against this time limit.

Ice and Rain Protection SystemsSustained Icing Conditions Airspeed . . . . . . . . . . . . . 140 Kts MINIMUM On King Air B200 S/Ns BB-743, 793, 829, 854 to 870, 874 to 891,

894, 896 to 911, 913 to 1438, 1440 to 1443; BL-37 to 138, sustained flight in icing conditions is prohibited with flaps extended. This does not include approach and landing, if needed.

NTTE: The rudder boost system may not operate when the brake deice system is in use.

Nptional Brake Deice System Do not operate system above 15°C ambient temperature. Do not operate system longer than 10 minutes (one deice timer cycle)

with the landing gear retracted. If operation does not automatically terminate approximately 10 minutes after gear retraction, manually select the system off.

Maintain 85% N1 or higher during periods of simultaneous brake deice and wing boot operation. If inadequate pneumatic pressure is developed for proper wing boot inflation, select brake deice system off.

Both sources of instrument bleed air must be in operation. Select brake deice system off during single engine operation.

Pneumatic Deice BootsMinimum Ambient Temperature for Operation of Deicing Boots . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -40°C

Quick Reference



Ice Vanes (Inertial Separator System) The ice vanes shall be extended for operations in ambient temperature of

+5°C or below when flight free of visible moisture cannot be assured. On King Air 200 aircraft, the ice vanes shall be retracted for all

operations in ambient temperatures of +15°C or above. On King Air B200 aircraft, ICE VANES LEFT and RIGHT shall be

extended or ENGINE ANTI-ICE LEFT and RIGHT shall be ON for operation in ambient temperatures of +5°C or below when flight free of visible moisture cannot be assured.

On King Air B200 aircraft, ICE VANES LEFT and RIGHT shall be retracted or ENGINE ANTI-ICE LEFT and RIGHT shall be OFF for all takeoff and flight operations in ambient temperatures of above +15°C.

On King Air 200/B200 S/Ns prior to BB-1439; prior to BL-138, once the manual override system is activated (i.e., anytime the ICE VANE EMERGENCY MANUAL EXTENSION handle has been pulled out), do not attempt to operate the ice vanes electrically until the override assembly inside the engine cowling has been properly reset on the ground. Even after the manual extension handle has been pushed back in, the manual override system is still engaged. Ice vanes should be extended for all ground operations for all B200 models. It is also recommended for all 200s.

Procedures for mechanically resetting the Ice Vane EMERGENCY MANUAL EXTENSION system can be found in the AFM, Chapter 8, Servicing and Handling Under Miscellaneous Maintenance.

Powerplant Number of Engines – 2 Engine Manufacturer – Pratt & Whitney of Canada

(Longueuil, Quebec, Canada) Engine Model Number PT6A-41 (King Air 200) or PT6A-42

(King Air B200) Do not lift power levers in flight. Lifting the power levers in flight

or moving the power levers in flight below the flight idle position could result in nose down pitch and a descent rate leading to aircraft damage and injury to personnel.

Overtemperature limits and actions required are presented in Figures 2-2, 2-3, 2-4, and 2-5.

Overtorque limit and actions required are shown on Figure 2-5.


2-2

Adjustments

AREA A 1. DETERMINE AND CORRECT CAUSE OF OVERTEMPERATURE.

2. RECORD IN ENGINE LOG BOOK.AREA B PERFORM HOT SECTION INSPECTION.AREA C RETURN ENGINE TO OVERHAUL.

NTTE: Interturbine temperatures shown make no allowance for instrument errors.

PT6A-41 Overtemperature Limits – All Conditions Except Starting

Quick Reference



2-3

Adjustments

AREA A NO ACTION REQUIRED.AREA B 1. DETERMINE AND CORRECT CAUSE OF

OVERTEMPERATURE.2. RECORD IN ENGINE LOG BOOK.

AREA C PERFORM HOT SECTION INSPECTION.AREA D 1. RETURN ENGINE TO AN APPROVED

OVERHAUL FACILITY2. FOR AIRLINES UTILIZING THE MODULAR

PROGRAMA. DO NOT SECTION INSPECTIONB. RETURN POWER SECTION TO OVERHAULC. RETURN FUEL CONTROL UNIT TO

OVERHAULD. INSPECT COMPRESSORE. INSPECT PNEUMATIC LINE (P3) AND

REPLACE P3 AIR FILTER


PT6A-40, -42 AND -42A Overtemperature Limits – All Conditions Except Starting


Adjustments

AREA A 1. DETERMINE AND CORRECT CAUSE OF OVERTEMPERATURE.

2. VISUAL INSPECT THROUGH EXHAUST DUCT.

3. RECORD IN ENGINE LOG BOOK.AREA B PERFORM HOT SECTION INSPECTION.AREA C SHIP ENGINE TO OVERHAUL.


PT6A-38, -40, -41, -42, and -42A Overtemperature Limits – Starting Conditions Only

2-4

Quick Reference



Adjustments

AREA A NO ACTION REQUIREDAREA B DETERMINE AND CORRECT CAUSE OF

OVERTORQUE.RECORD IN ENGINE LOG BOOK.

AREA C RETURN ENGINE TO AN APPROVED OVERHAUL FACILITY.

PT6A-40, -41, -42 AND -42A Overtorque Limits – All Conditions

2-5


Engine Operating Limits The following limitations presented in Tables 2-6, 2-7, and 2-8

shall be observed. Each column presents limitations. The limits represented do not necessarily occur simultaneously. Refer to Pratt & Whitney Engine Maintenance Manual for specific actions required if limits are exceeded.

Oil Specifications Any oil specified by brand name in the latest revision of Pratt &

Whitney SB 3001 is approved for use in the PT6A-41 (King Air 200) or the PT6A-42 (King Air B200) engine.

Operating Condition SHP Torque (ft-lbs)1

Max Observed ITT (°C)

N1 RPM N1 % Prop RPM N2

Oil Press (PSI)2

Oil Temp °C

Starting – – 10003 – – – – -40 (min)Low Idle – – 6604 19,500 52 (min) – 60 (min) -40 (min)High Idle – – – – 5 – – -40 (min)Takeoff9 850 2230 750 38,100 101.5 2000 105 to 135 10 to 99Max Continuous and Cruise 850 22306 750 38,100 101.5 2000 105 to 135 10 to 99Cruise Climb and Rec Cruise 850 22306 725 38,100 101.5 2000 105 to 135 0 to 99Max Reverse7 – – 750 – 88 1900 105 to 135 0 to 99Transient – 27503 850 38,5008 102.68 22003 – 0 to 1049

Table 2-6: King Air 200 Engine Operating Limits (PT6A-41)1 Torque limit applies within range of 1,600 to 2,000 propeller RPM (N2). Below 1,600 RPM, torque limited to 1,100 ft-lbs.2 When gas generator speeds are above 27,000 RPM (72% N1) and oil temperatures are between 60 and 71°C, normal oil pressure are:

100 to 135 PSI below 21,000 ft and 85 to 135 PSI at 21,000 ft and above

During extremely cold starts, oil pressure may reach 200 PSI. Oil pressure between 60 and 85 PSI is undesirable; it should be tolerated only for the completion of the fight, and then only at a reduced power setting not exceeding 1,100 ft-lbs torque. Oil pressure below 60 PSI is unsafe; it requires that either the engine be shut down, or that a landing be made as soon as possible with minimum power to sustain flight. Fluctuations of ±10 PSI are acceptable.

3 These values are time limited to 5 seconds.4 High ITT at ground idle may be corrected by reducing accessory load and/or increasing N1 RPM.5 At approximately 70% N1.6 Cruise torque values vary with altitude and temperature.7 This operation is time limited to one minute.8 These values are time limited to 10 seconds.9 These values are time limited to 5 minutes.

Quick Reference






Oil Press (PSI)2

Oil Temp °C3, 4

Starting – – 10005 – – – – -40 (min)Low Idle – – 7506 21,000 56 (min) – 60 (min) -40 to 99High Idle – – – – 7 – – -40 to 99Takeoff 850 2230 800 38,100 101.5 2000 100 to 135 0 to 99Max Continuous and Cruise 850 22308 800 38,100 101.5 2000 100 to 135 0 to 99Cruise Climb and Rec Cruise 850 22308 770 38,100 101.5 2000 100 to 135 0 to 99Max Reverse9 – – 750 – 88 1900 100 to 135 0 to 99Transient – 27505 850 38,50010 102.610 22005 2002 0 to 10411

Table 2-7: King Air B200 Engine Operating Limits (PT6A-42); S/Ns BB-743, 793, 829, 854 to 870, 874 to 891, 894, 896 to 911, 913 to 1438, 1440 to 1443; BL-37 to 138

1 Torque limit applies within range of 1,600 to 2,000 propeller RPM (N2). Below 1,600 RPM, torque limited to 1,100 ft-lbs.2 When gas generator speeds are above 27,000 RPM (72% N1) and oil temperatures are between 60 and 71°C, normal oil pressure are:

100 to 135 PSI below 21,000 ft and 85 to 135 PSI at 21,000 ft and above

During extremely cold starts, oil pressure may reach 200 PSI. Oil pressure between 60 and 85 PSI is undesirable; it should be tolerated only for the completion of the fight, and then only at a reduced power setting not exceeding 1100 ft-lbs torque. Oil pressure below 60 PSI is unsafe; it requires that either the engine be shut down, or that a landing be made at the nearest suitable airport with minimum power to sustain flight. Fluctuations of ±10 PSI are acceptable.

3 A minimum oil temperature of 55°C is recommended for fuel heater operation at takeoff power.4 Oil temperature limits are -40°C and 99°C. However, temperature of up to 104°C are permitted for a maximum time of 10 minutes.5 These values are time limited to 5 seconds.6 High ITT at ground idle may be corrected by reducing accessory load or increasing N1 RPM.7 At approximately 70% N1.8 Cruise torque values vary with altitude and temperature.9 This operation is time limited to one minute.10 These values are time limited to 10 seconds.11 Values above 99°C are time limited to 10 minutes.





Oil Press (PSI)2

Oil Temp °C3, 4

Starting – – 10005 – – – – -40 (min)Low Idle – – 7506 22,875 61 (min) 12 60 (min) -40 to 99High Idle – – – – 7 – – -40 to 99Takeoff6 850 2230 800 38,100 101.5 2000 100 to 135 0 to 99Max Continuous and Cruise 850 22308 770 38,100 101.5 2000 100 to 135 0 to 99Cruise Climb and Rec Cruise 850 22308 770 38,100 101.5 2000 100 to 135 0 to 99Max Reverse9 – – 750 – 88 1900 100 to 135 0 to 99Transient – 27505 850 38,50010 102.610 22005 2002 0 to 10411

Table 2-8: King Air B200 Engine Operating Limits (PT6A-42); S/Ns BB-1439, BB-1444 and subsequent except BB-1463; BL-139 and subsequent; BW-1 and subsequent

1 Torque limit applies within range of 1,600 to 2,000 propeller RPM (N2). Below 1,600 RPM, torque limited to 1,100 ft-lbs.2 When gas generator speeds are above 27,000 RPM (72% N1) and oil temperatures are between 60 and 71°C, normal oil pressure are:

Below 21,000 ft 100 to 135 PSI; 21,000 ft and above 85 to 135 PSI

During extremely cold starts, oil pressure may reach 200 PSI. Oil pressure between 60 and 85 PSI is undesirable; it should be tolerated only for the completion of the fight, and then only at a reduced power setting not exceeding 1,100 ft-lbs torque. Oil pressure below 60 PSI is unsafe; it requires that either the engine be shut down, or that a landing be made at the nearest suitable airport with minimum power to sustain flight. Fluctuations of ±10 PSI are acceptable.

3 A minimum oil temperature of 55°C is recommended for fuel heater operation at takeoff power.4 Oil temperature limits are -40°C and 99°C. However, temperature of up to 104°C are permitted for a maximum time of 10 minutes.5 These values are time limited to 5 seconds.6 High ITT at ground idle may be corrected by reducing accessory load or increasing N1 RPM.7 At approximately 70% N1.8 Cruise torque values vary with altitude and temperature.9 This operation is time limited to one minute.10 These values are time limited to 10 seconds.11 Values above 99°C are time limited to 5 minutes.12 1,100 RPM for McCauley propeller and 1,180 RPM for Hartzell propeller.

Quick Reference



Oxygen If the oxygen system pressure drops to less than 50 PSI, system-

purge is required.

Propellers

King Air 200/B200Propeller Rotational Speed LimitsTransients not exceeding 5-seconds . . . . . . . . . . . . . . . . . . .2,200 RPMReverse . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1,900 RPMAll other conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .2,000 RPM

Minimum Idle SpeedHartzell Propellers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1,180 RPMMcCauley Propellers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1,100 RPM

Propeller Rotational Overspeed Limits The maximum propeller overspeed limit is 2,200 RPM and time-

limited to five-seconds. Sustained propeller overspeeds faster than 2,000 RPM indicate failure of the primary governor. Flight may be continued at propeller overspeeds up to 2,080 RPM provided torque is limited to 1,800 foot-pounds.

Sustained propeller overspeeds faster than 2,080 RPM are unapproved. *

*2,120 RPM BB1444 and after.



Quick Reference



Instrument MarkingsRefer to Tables 2-9 and 2-10 for airspeed indicator markings.

Marking KCAS Value or Range

KIAS Value or Range

Significance

Red Line 91 86 VMCA or Air Minimum Control SpeedWhite Arc 80 to 144 75 to 146 Full-Flap Operating Range

Wide White Arc 80 to 102 75 to 99 Lower Limit is Stalling Speed (VSO) at maximum weight with full flaps (100%) and idle power.

Narrow White Arc 102 to 144 99 to 146 Lower Limit is Stalling Speed (VS) at maximum weight with flaps up (0%) and idle power. Upper limit is maximum speed permissible with flaps extended beyond Approach (more than 40%).

White Triangle 200 200 Maximum Flaps-to/at-Approach (40%) Speed

Blue LIne 122 121 One Engine Inoperative Best Rate-of-Climb Speed

Red/White Hash-Marked Pointer2

270 269 Maximum speed for any operationor value equal to 0.48M,

whichever is lowerRed/White Hash-Marked Pointer3

260 259 Maximum speed for any operationor value equal to 0.48M,

whichever is lower

Table 2-9: King Air 200 Airspeed Indicator Markings1

1 The airspeed indicator is marked in CAS values.2 BB-2 through 198 except those modified by Beech Kit 101-5033-1.3 BB-199 and subsequent, BL-1 and subsequent, any earlier aircraft modified by Beech Kit

101-5033-1.


Marking KCAS Value or Range

KIAS Value or Range

Significance

Red Line 91 86 VMCA or Air Minimum Control SpeedWhite Arc 80 to 155 75 to 157 Full-Flap Operating Range

Wide White Arc 80 to 102 75 to 99 Lower Limit is Stalling Speed (VSO) at maximum weight with full flaps (100%) and idle power.

Narrow White Arc 102 to 155 99 to 157 Lower Limit is Stalling Speed (VS) at maximum weight with flaps up (0%) and idle power. Upper limit is maximum speed permissible with flaps extended beyond Approach (more than 40%).

White Triangle 200 200 Maximum Flaps-to/at-Approach (40%) Speed

Blue LIne 122 121 One Engine Inoperative Best Rate-of-Climb Speed

Red/White Hash-Marked Pointer

260 259 Maximum speed for any operationor value equal to 0.52M, whichever is lower

Table 2-10: King Air B200 Airspeed Indicator Markings1

1 The airspeed indicator is marked in IAS values.

Cabin Pressure Differential Gauge

King Air 200 – before BB-195Green Arc (approved operating range) . . . . . . . . . . . . . . . 0 TO 6.0 PSIRed Arc (unapproved operating range) . . . 6.0 PSI TO END OF SCALE

King Air 200 – BB-195 and subsequent; BL-1 and subsequentGreen Arc (approved operating range) . . . . . . . . . . . . . . . 0 TO 6.1 PSIRed Arc (unapproved operating range) . . . 6.1 PSI TO END OF SCALE

King Air B200Green Arc (approved operating range) . . . . . . . . . . . . . . . 0 TO 6.6 PSIRed Arc (unapproved operating range) . . . 6.6 PSI TO END OF SCALE

Pneumatic GaugeGreen Arc (normal operating range) . . . . . . . . . . . . . . . . . 12 TO 20 PSIRed Line (maximum operating limit) . . . . . . . . . . . . . . . . . . . . . . 20 PSI

Quick Reference



Fuel QuantityYellow Arc (No Takeoff Range) . . . . . . . . . . . . . . . . . . . . . 0TO265 LBS

Vacuum/Gyro Suction Gauge

King Air 200Narrow Green Arc (normal from 35,000 to 15,000 ft MSL) . . . . . . . . . . . .3.0 TO 4.3 IN HGWide Green Arc (normal from 15,000 ft to sea level) . . . . . . . . . . . . . . .4.3 TO 5.9 IN HG

King Air 200 (alternate gauge) and King Air B200Narrow Green Arc (normal from 35,000 to 15,000 ft MSL) . . . . . . . . . . . .2.8 TO 4.3 IN HGWide Green Arc (normal from 15,000 ft to sea level) . . . . . . . . . . . . . . .4.3 TO 5.9 IN HG35K Marked on Face of Gauge at . . . . . . . . . . . . . . . . . . . . . .3.0 IN HG15K Marked on Face of Gauge at . . . . . . . . . . . . . . . . . . . . . .4.3 IN HG


Propeller Deice AmmeterGreen Arc (normal operating range) . . . . . . . . . . . . . . . . 14 TO 18 AMPBB-1444 and subsequent . . . . . . . . . . . . . . . . . . . . . . . . 18 TO 24 AMP

Powerplant Instrument MarkingsInstrument Red Line

Minimum LimitYellow Arc

Caution RangeGreen Arc

Normal OperatingRed Line Maximum Limit

6

Interstage Turbine Temperature 1 – – 400 to 750°C 2

400 to 800°C 3

750°C 2

800°C 3

Torquemeter – – 400 to 2230 ft-lbs 2230 ft-lbsPropeller Tachometer – – 1600 to 2000 RPM 2000 RPMGas Generator Tachometer – – 61 to 101.5% 4 101.5%Oil Temperature – – 10 to 99°C 99°COil Pressure 5 60 PSI 60 to 100 PSI 4 105 to 135 PSI 2

100 to 135 PSI 3

85 to 135 PSI 4

200 PSI135 PSI 4

Table 2-11: Powerplant Instrument Markings1 Starting limit (dashed red line): 1000°C.2 King Air 200.3 King Air B200 S/Ns BB-743 to 1443 with exceptions; BL-37 to 138.4 King Air B200 S/Ns BB-1439, 1444 through 1485 except 1463 and 1484; BL-139 and BL-140.5 A dual-band yellow/green arc extends from 85 to 100 psi, indicating the extended range of normal oil pressure for operation at, or above, 21,000 ft. A red diamond at 200 psi indicates upper transient limit.

6 Red line maximum limits are maximum continuous or cruise values. Transients may occur at higher values.

Quick Reference



System Data Summaries

Electrical Systems

DC System

Power Sources BatteryTwo starter/generators 250A (STD)External power unit

Control Battery switchStart switchGenerator switch

Two-position on BB-88 and priorThree-position on BB-89 and subsequent

Distribution Hot Battery busBattery relayMain Battery busLeft and right start relaysIsolation busLeft and right Generator busesNos. 1 through 4 Dual-Fed busesAvionics buses No. 1, No. 2, and optional No. 3

Monitoring L/R DC voltmeterL/R GEN annunciatorsBattery charge annunciatorEXT POWER annunciatorVolt/loadmeter

Protection Voltage regulator Generator paralleling Reverse current sensing and control Over-voltage protection Over-excitation protection Under-excitation protection GPU reverse polarity sensing Generator buses

325A isolation limiters CBs and current limiters

Dual-Fed buses60A current limiters and/or 50A CBs and diodes (70A)

Hot Battery BusFuses or CBs


Electrical Systems (continued)

AC SystemPower Sources No. 1 Inverter

No. 2 InverterControl INVERTER NO. 1/OFF/NO. 2 switchDistribution Generator buses

Inverters No. 1 and No. 226 VAC bus115 VAC avionics

Monitoring INVERTER annunciatorVolt/Frequency meter

Protection DC to inverter50A or 60A current limiters

Inverter outputFuses and CBs

Fire Protection

Power Sources No. 1 Dual-Fed bus – fire detectionHot Battery bus – optional fire extinguishingPortable hand fire extinguishers

Distribution Extinguisher bottle to corresponding engine (no crossfiring)Control TEST SWITCH FIRE DET (& EXT, if installed)

ENG FIRE-PUSH TO EXT (L/R) lens/switch (if installed)Monitor ENG FIRE (L/R) annunciators

Red ENG FIRE-PUSH TO EXT (L/R) lens (if installed) Amber D lens to confirm electrical wiring continuity (if installed)Green OK lens to test system (if installed)MASTER WARNING (L/R) flashers

Protection FIRE DET CB (5A)FIRE extinguisher fuse (prior to BB-1096) (if installed)FIRE EXTINGUISHER CB (5A)

(BB-1096 and subsequent) (if installed)

Quick Reference



Flight Controls

Flap SystemPower Sources Electric motor and control No. 3 Dual-Fed busControl Flap handleMonitor Flap position indicatorProtection Power and Control CBs

Split flap protectionLimit switches

Fuel Systems

Main Fuel SystemPower Sources Hot Battery bus (BB-1097, 1095 and prior) or

No. 3 Dual-Fed busLeft standby pump

Hot Battery bus (BB-1097, 1095 and prior) orNo. 4 Dual-Fed bus

Right standby pumpNo. 4 Dual-Fed bus

Crossfeed valveHot Battery bus and/or No. 3 and No. 4 Dual-Fed buses Firewall shutoff valves (L/R)

Distribution Wing tanks (gravity feed) to nacelle tank Nacelle tank to engine

Control SwitchesSTANDBY PUMP CROSSFEED FIREWALL SHUTOFF VALVES

Monitor Main fuel gauges Fuel flow indicator Annunciators

FUEL CROSSFEED FUEL PRESS CROSSFEED (closes motive flow valve on receiving side, opens crossfeed valve, turns on standby boost pump on feeding side, and illuminates crossfeed annunciator)

Protection Circuit breakers Check valves Fuses Fuel drain system Fuel filters (pressure switches) Vent systems Oil/Fuel heat exchanger


Auxiliary Fuel System

Power Sources Motive FlowDistribution Auxiliary (center) tank (automatic transfer to nacelle tank with

AUX TRANSFER switch in AUTO)Control Switches

AUX TRANSFER OVERRIDE-AUTO(opens motive flow valve)

Monitor Aux fuel gaugesNO TRANSFER lights

Protection Circuit breakersFuses

Ice and Rain Systems

Surface Deice SystemPower Sources Bleed Air

No. 1 Dual-Fed busDistribution Wing leading edge boots

Horizontal stabilizer leading edge bootsControl DEICE CYCLE switch

SINGLE – inflation/deflation of wing boots then horizontal stabilizer bootsMANUAL – inflation of all boots simultaneously

Monitor Visual monitoring for wingPneumatic gauges

Protection Circuit breakers

Prop Heat System

Power Sources No. 1 Dual-Fed bus (auto)No. 3 and 4 Dual-Fed bus (manual)

Distribution Heated boot for each propeller bladeControl Switches

PROP AUTOPROP MANUALINNER/OUTER (200 only)

Monitor Prop ammeterLoadmeters

Protection Circuit breakers/Circuit breaker switch (AUTO)

Quick Reference



Anti-Ice Systems

Brake Deice SystemSources Engine P3 bleed airControl BRAKE DEICE switchMonitor BRAKE DEICE ON annunciatorProtection 10-minute timer

Pitot HeatPower Source Dual-Fed buses Nos. 1/2Control PITOT circuit breaker switchesProtection Circuit breaker switch

Stall Warning HeatPower Source Dual-Fed bus No. 2Control STALL WARN circuit breaker switches

Landing gear safety switchProtection Circuit breaker switch

Fuel Vent HeatPower Source Dual-Fed buses Nos. 1/2Control FUEL VENT circuit breaker switches

Windshield HeatPower Source L/R GEN busControl WSHLD ANTI-ICE switchesProtection Circuit breaker (5A)

Temperature sensing element Temperature controller 50A current limiters

Ice VanesPower Source Dual-Fed buses Nos. 1/2Control ICE VANE switches

VANE MANUAL PULL handleMonitor ICE VANE amber and green annunciatorsProtection Circuit breakers

Manual override system


Landing Gear Systems

Tlectro-Mechanical Landing Gear SystemPower Source No. 2 Dual-Fed bus

Landing gear control relayRight Generator bus

28 VDC split-field 11/2 HP motorControl LDG GEAR CONTROL handle

EMERGENCY ENGAGE handleMonitor Gear handle light

Gear warning horn Gear DOWN position lights

Protection Landing gear relay (5A) Circuit breaker (60A) Right main gear squat switch Emergency engage handle Limit switches Dynamic brake relay Solenoid-operated down-lock hook (landing gear handle)

Hydraulic Landing Gear SystemPower Sources Right Generator bus

No. 2 Dual-Fed busLanding gear control power relay

Electric Motor DrivenHydraulic Pump (Power Pack)

Distribution Landing gearControl LDG GEAR CONTROL handle

Pressure switch Down-lock switches (3) Time delay module

Monitor HYD FLUID LOW annunciator Accumulator precharge direct reading gauge

Protection Circuit breakersLANDING GEAR RELAY (5A) Landing gear powerpack (60A)

Pressure switch Thermal relief valve Down-lock switches (3) Internal nose gear mechanical lock Squat switches (L/R) Low fluid level sensor Time delay module Solenoid-operated down-lock hook (landing gear handle)

Quick Reference



Landing Gear Systems (continued)

Brake SystemPower Source Hydraulic pressureDistribution Master cylinders

Parking brake valvesControl Brake pedals

PARKING BRAKE handle Emergency braking: reverse propeller for taxiing or slowing

(-3° blade angle zero thrust) top of red and white strips on throttle quadrant

Shuttle valves S/N prior to BB-666: valve adjacent to each set of pedals permit changing braking action from one to the other

S/N BB-666 and subsequent: dual brakes plumbed in series to allow either set of pedals to perform

Pneumatic Systems

Bleed Air SystemPower Sources Bleed air (each engine – station P3)Distribution Prior to 18 PSI regulator:

Brake deice Rudder boost ΔP switch

After 18 PSI regulator Bleed air warning system Rudder boost servos Flight hour meter Door seal (if installed) Vacuum Deice boots

Control BLEED AIR VALVE switchesMonitor PNEUMATIC PRESSURE gauge

GYRO SUCTION gaugeProtection BL AIR FAIL annunciators

Bleed air shutoff valves Relief valves Check valves


Air Conditioning/Heating System

Power Sources Engine-bleed air – heating Right engine – Freon system

Distribution Engine compressor bleed air Environmental control unit Cabin Cockpit

Control Swit ches BLEED AIR VALVES MANUAL TEMP VENT BLOWER AFT BLOWER RADIANT HEAT (prior to BB-1439)ELEC HEAT (BB-1439 and subsequent)

CABIN TEMP MODE selector CABIN TEMP control selector Right engine RPM above 60%

Monitor CABIN AIR gauge Thermostat AIR CND N1 LOW annunciator

Protection High and low pressure switches N1 speed switch47 PSI pressure switch

Pneumatic Systems (continued)

Pressurization SystemPower Sources Bleed air (each engine – station P3)Distribution Flow control unit

Air-to-air heat exchangers Cockpit Cabin

Control Swit ches BLEED AIR VALVES CABIN PRESS

Pressurization controllerMonitor Cabin altitude and differential pressure gauge

Cabin VSIProtection BL AIR FAIL annunciators

Bleed air shutoff valves CABIN ALT annunciator Squat switch Outflow/safety valves

(negative/maximum differential relief)Passenger oxygen mask auto deployment system

Quick Reference



Powerplant System

Power Sources Rev erse flow, free turbine engines Pratt and Whitney PT6A-41 (200)Pratt and Whitney PT6A-42 (B200)

Distribution Air from inlet screen to: Axial-flow compressor Centrifugal-flow compressor section Annular combustion chamber

Hot, high-pressure gas from combustion chamber to: Single-stage, axial-flow turbine (to drive compressor and accessory section) Two-stage, axial-flow turbine (to drive power turbine shaft)

Pow er turbine shaft drives Propeller Reduction gearbox

Control Lev ers Power Propeller Condition (fuel control unit)

Swit ches IGNI TION AND ENGINE START

(ON/OFF/STARTER ONLY) (L/R) ENG AUTO IGNITION (ARM/OFF) (L/R) ICE VANE (EXTEND/RETRACT) (L/R) PROP GOV (TEST/OFF)

Monitor Engi ne Operation ITT Torque Prop RPM N1 RPM

Indic ators FUEL FLOW OIL TEMP OIL PRESS

Warn ing annunciators ENG FIRE L/R FUEL PRESS L/R OIL PRESS L/R CHIP DETECT L/R

Cauti on/advisory annunciators ICE VANE L/R (amber) ICE VANE EXT L/R (green) AUTOFEATHER L/R IGNITION ON (L/R)

Protection Engi ne operating parameters (overspeed, overtemperature, overtorque)

N1 governor (overspeed)Torque limiter (overtorque) Magnetic chip detector (oil contamination warning) Oil -to-fuel heat exchanger

(fuel warming – see Ice and Rain Protection)Fuel shutoff valves



Operating Procedures 3



This section presents four individual elements of flight operations: preflight inspection, expanded normal procedures, Standard Operating Procedures (SOP), and maneuvers. Although they are addressed individually in this manual, their smooth integration is critical to ensuring safe, efficient operations.The Preflight Inspection chapter illustrates a step-by-step exterior inspection of the aircraft. Preflight cockpit and cabin checks are also discussed. The Expanded Normal Procedures chapter presents checklists for normal phases of flight. Each item, when appropriate, is expanded to include limitations, cautions, warnings, and light indications.The Standard Operating Procedures chapter details Pilot Flying/Pilot Not Flying callouts and verbal or physical responses.The Maneuvers chapter pictorially illustrates normal and emergency profiles. Additionally, written descriptions are included for most phases of flight with one or both engines operating.




Preflight Inspection 3A

King Air 200 3A-1December 2011


ContentsPreflight InspectionChecklist Usage

Figure: Cockpit Flow Pattern (Left Seat) ..................................... 3A-6Figure: Cockpit Flow Pattern (Right Seat) .................................. 3A-6

Cockpit InspectionExternal Inspection

Figure: Preflight Inspection Walkaround Path ............................ 3A-13A Left Wing ............................................................................................ 3A-15B Left Engine ........................................................................................ 3A-21C Nose ................................................................................................... 3A-25D Right Engine ...................................................................................... 3A-29E Right Wing ......................................................................................... 3A-33F Tail....................................................................................................... 3A-39

Cabin Inspection

King Air 2003A-2December 2011


ThFs pair FntrntFonally lrft blank.

Preflight Inspection



Preflight InspectionAn essential part of the preparation made before any flight is the preflight inspection during which a crewmember verifies the aircraft’s physical readiness. After a thorough initial preflight, subsequent same-day inspections are abbreviated.No detail is overlooked during the first preflight of the day. To ensure safety, correct abnormal conditions (e.g., low tire pressure) as well as minor discrepancies prior to flight.Begin the preflight inspection inside the aircraft to verify the initial cockpit setup and essential functions. Follow with the exterior inspection, which begins just outside the aircraft door, proceeds clockwise around the aircraft, and ends at the aircraft door. Finally, return to the aircraft interior to check passenger and crew compartments for readiness.







Checklist UsageTasks are executed in one of two ways: as a sequence that uses the layout of the cockpit controls and

indicators as cues (i.e., “flow pattern”) as a sequence of tasks organized by event rather than panel location

(e.g., After Takeoff, Gear – UP, Flaps – UP).Placing items in a flow pattern or series provides organization and serves as a memory aid.A challenge-response review of the checklist follows execution of the tasks; the Pilot Not Flying (PNF) calls the item, and the appropriate pilot responds by verifying its condition (e.g., “Propeller Anti-Ice” [challenge] – “ON” [response]).Two elements are inherent in the execution of normal procedures: use of either the cockpit layout or event cues to prompt the correct

switch and/or control positions followed by the normal checklist as a done list

use of normal checklists as “done” lists.



Cockpit Flow Pattern (Left Seat)

Cockpit Flow Pattern (Right Seat)




Cockpit InspectionWARIRN

Beards and mustaches should be carefully trimmed so that they will not interfere with proper oxygen mask sealing. The fit of the oxygen mask should be checked on the ground for proper sealing. Hats and “earmuff” type headsets must be removed prior to donning masks. Headsets and eyeglasses worn by crew members may interfere with quick-donning capabilities.

CWUUIIR The elevator trim system must not be manually, electrically, or by action of the autopilot forced past the limits indicated on the elevator trim indicator scale.

Control Locks. . . . . . . . . . . . . . . . . . . . . . . . . . . . REMOVED/STOWEDTrim Tabs (Elevator/Aileron/Rudder) . . . . . . . . . . . . . . . . . . . . .0 UNITSFuel Control Panel CBs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INRight Side Panel CBs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INLanding Gear Switch Handle . . . . . . . . . . . . . . . . . . . . . . . . . . . . DOWNLanding Gear Control CB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INElectrical Switches. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFF/AUTOOxygen System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKED

Passenger Manual Override . . . . . . . . . . . . . . . . . . . . . .PUSH OFFOxygen System Ready . . . . . . . . . . . . . . . . . . . . . . . . . . . PULL ONCrew Diluter Demand Masks . . . . . . . . . . . . . . . . . . . . .DON MASK/

CHECK/STOWDon oxygen mask. Check fit and operation. After performing oxygen mask check, stow the mask so that it is available for immediate use.

Oxygen Duration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DETERMINERefer to Tables 3A-1 and 3A-2, following page.

NTTe: A bottle of 1,850 PSIG at 15°C is fully charged (100% of capacity). Read directly from the table. Read the oxygen presure from the gauge. Read IOAT (with battery ON). Determine the percent of usable capacity from the following graph (e.g., 1,100 PSI at 0°C equals 57%). Compute the oxygen duration in minutes from the table by multiplying the full bottle duration by the percent of usable capacity (e.g., pilot and copilot with masks set at 100% plus 6 passengers equals 10 people using oxygen).



Stated Cylinder

Size (cu ft)

** Number of People Using1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16** 17**

Duration in Minutes22 151 75 50 37 30 25 21 18 16 15 13 12 11 10 10 * *

49 or 50 334 167 111 83 66 55 47 41 37 33 30 27 25 23 22 20 1966 445 222 148 111 89 74 63 55 49 44 40 37 34 31 29 27 26

76 or 77 514 257 171 128 102 85 73 64 57 51 46 42 39 36 34 32 30115 772 386 257 193 157 128 110 96 85 77 70 64 59 55 51 48 45

Table 3A-1: King Air 200 Oxygen Duration with Full Bottle (100% Capacity)

Stated Cylinder

Size (cu ft)

** Number of People Using1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16** 17**

Duration in Minutes22 144 72 48 36 28 24 20 18 16 14 13 12 11 10 * * *50 317 158 105 79 63 52 45 39 35 31 28 26 24 22 21 19 1870 488 244 162 122 97 81 69 61 54 48 44 40 37 34 32 30 28115 732 366 244 183 146 122 104 91 81 73 66 61 56 52 48 45 43

Table 3A-2: King Air B200 Oxygen Duration with Full Bottle (100% Capacity)* Will not meet oxygen requirements.

** For oxygen duration computations, count each diluter-demand crew mask in use as 2 (e.g., with 4 passengers and a crew of 2, enter the table at 8 people using).




Hot Battery Bus/Fuel Panel Check . . . . . . . . . . . . . . . . . PERFORMEDCircuit Breakers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INFirewall Fuel Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CLOSE

Listen for operation.Quiet Ramp:

Standby Pumps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ONListen for operation.

Battery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ONFUEL PRESS annunciators illuminate.

Firewall Fuel Valve (both) . . . . . . . . . . . . . . . . . . . . . . . . .OPENFUEL PRESS annunciators extinguish.

Standby Pumps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFFFUEL PRESS annunciators illuminate.

Noisy Ramp:Standby Pump CBs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PULLFirewall Fuel Valves CBs . . . . . . . . . . . . . . . . . . . . . . . . . . PULLFirewall Fuel Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CLOSEStandby Pumps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ONBattery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON

FUEL PRESS annunciators illuminate.Firewall Fuel Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OPEN

FUEL PRESS annunciators extinguish.Standby Pumps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF

FUEL PRESS annunciators illuminate.Standby Pump CBs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INFirewall Fuel Valves CBs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IN

Crossfeed Switch . . . . . . . . . . . . ALTERNATELY LEFT AND RIGHTFUEL CROSSFEED annunciator illuminates and extinguishes and both FUEL PRESS annunciators extinguish.

Voltmeters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS TOTESTBoth voltmeters should read normal battery voltage of 24V. No voltage on one side indicates current limiter is out; 23V minimum for battery start; 20V minimum for external power

NTTe: On BB-1096, 1098 and subsequent; BL-58 and subsequent, standby pumps are not on the Hot Battery bus.

Flaps . . . . . . . . . . . . . . . . . . . SELECTED UP/INDICATING UPFuel Quantity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKEDBattery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFFParking Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETCockpit Fire Extinguisher . . . . . . CHECK PRESSURE/SECURE







External InspectionUnfold the preflight inspection diagram on the following page for ease of reference. Note that each segment of the preflight inspection is identified by letters A through F. Subsequent pages provide sequenced checklists of each preflight inspection segment. A large locator photo identifies the general location of each inspection. Adjacent photos detail the checklist items. Photos read from left to right.Limitations and specifications are noted if relevant to the checklist.







Preflight Inspection Walkaround Path

C

B

A

D E

F



1

2

3

4




A Left Wing1.Flaps: Both flap panels should be fully retracted. Ensure they are not bent nor distorted. Some movement of the flaps should be possible, and there should be no sign of flap binding. Note especially the condition of rivets on top of the flaps. The inboard flap must be flush with the bottom of the fuselage; the outboard flap trailing edge may vary from 1/4 inch above to 1/4 inch below the inboard flap.2.Fuel Sump aft of wheel well: A total of six fuel sumps need to be drained (Figure 3A-1) on the fuel system. Two sumps are quick drain style and four are flush-mounted. Use a fuel tester probe to push up on the spring-loaded push-to-drain type valve. Check for water and/or sediment. Dispose of samples in a fuel sample container. At this point, check the gravity feed line from wing tanks to nacelle (aft of wheel well).3.Aileron and Trim Tab: The aileron and the trim tab should be in the neutral position as previously set in the cockpit. The aileron can be as much as ± 1/2 inch above or below the flap trailing edge. Check general condition and freedom of movement of the aileron. Check bonding wires for secure attachment. The trim tab should be aligned with the aileron on the inboard side. The hinge should not have excess play.4.Outboard Wing Fuel Sump: Drain the outboard flush mounted wing sump forward of the aileron. Check that inverter cooling louvers are all clear.

AIR INLET OUTBOARD WING

FUEL DRAIN RECESSED VENT

HEAT RAM VENT

GRAVITY FLOW LINE DRAIN

OUTLET STRAINER AND DRAIN

FUEL FILLER CAP

LEADING EDGE FUEL DRAIN

FUEL FILTER DRAIN STRAINER

DRAIN

AUX FILLER CAP

FUEL PUMP DRAIN

Figure 3A-1: Fuel System



5

6

7

8

9

11

12

13

14




A Left Wing (continued)5. Air Inlet: Check the general condition of the air inlet on the aft side of the wing tip. It should be clear and free of obstructions.6. Static Wicks: Check for three static wicks on each aileron and one on each wing tip. Ensure the static wicks are secure and free of damage.

NTTe: All wicks must be installed and in good condition when VLF/Omega equipment is used.

7. Wing Tip/Lights: Check the light assembly for lens cracks, security, and any indication of a burned out bulb.8. Main Fuel Tank: Ensure there are no leaks in the main fuel tank. Check the fuel cap near the wing tip; the locking tab should be flush with the surface of the cap and pointing aft.9. Stall Warning Vane: Check that the stall warning vane on the leading edge of the left outboard wing section moves freely and that the vane and mounting assembly are in good condition.10.Tie Downs and Chocks: Check that tie downs and chocks are removed.11.Outboard Deice Boot: The deice boot should be securely attached to the leading edge. Ensure that it is undamaged. Check for significant scratches or cuts in the boot. Check condition of the stall strip attached to the boot.12. Recessed Fuel Vent/Heated Ram Air Vent: Ensure both vents under the wing are unobstructed and undamaged. The heated ram air vent is aft and outboard from the recessed fuel vent.13. Nacelle Fuel Drains: Drain the flush-mounted nacelle fuel pump, strainer, and fuel filter drains forward of the wheel well. Check condition of samples.14. Landing Gear/Doors: Ensure tires are in good condition, doors are secure, struts are properly inflated, and wheel wells clean and free of fluid leaks. Linkage should be secure with no signs of unusual wear or cracking. Check uplock and downlock microswitches for condition. Brake deice lines, when installed, should be secure and free of fuel, oil, and hydraulic fluid. The deice manifold should be securely attached to the axle and free of damage. Brake lines should be secure and free of fraying or leaking.

Static ick LimitationDue to varying configurations, consult your MEL for number and position of static wicks.



15

16

17




A Left Wing (continued)15. Wing Locker: Door should be properly secured.16.Fire Extinguisher Pressure (if installed): Check that the fire extinguisher cylinder is properly serviced. The cylinder is located aft in the main wheel well. Reference Table 3A-3, Pressure versus Temperature chart.17.Ice Light: Check for condition and security.

Temp °F IndicatedPressure PSI

-40° 190 - 240-20° 220 - 2750° 250 - 315-20° 290 - 365-40° 340 - 42060° 390 - 48080° 455 - 550100° 525 - 635120° 605 - 730140° 700 - 840

Table 3A-3: Pressure Versus Temperature



1

2A

2B

3

4

5

6




B Left Engine1.Engine Oil: The oil filler access door is on the top left side of the engine cowling. Check quantity. A low reading may occur if the engine has been idle and oil is allowed to seep into the scavenge pump reservoir. If this occurs, motor engine two minutes and then re-check quantity. Replace dipstick; make sure caplock flange is turned clockwise to stop position.

NTTe: Service the oil system law Consumable Materials in the

Handling, Service, & Maintenance section (8) of the AFM and P&WC SB 3001.

DO NOT MIX different brands of oil (except as provided in Consumable Materials).

Normal operating range is FULL to 4 U.S.quarts low. Maximum oil consumption is 1 U.S.quart in 10 hours of

operation.

2.Left Cowling: Open cowling door (2A) and check for fuel, oil, and/or bleed air leaks. Check condition of linkages, hoses, and accessories. Ensure nuts or bolts are properly safetied. Check general condition of engine. Close cowling door and ensure it is secure and flush with surface. Check alignment marks on forward (upper) cowling bolts for proper indication (2B).3.Left Exhaust Stack: Check that the exhaust stack is free of cracks and secure. There should be no cracks in any weld of the exhaust pipe. Check condition of scupper inside the stack. Remove intake covers and prop restraints.4.Propeller: Ensure there are no cracks, nicks, corrosion, or other possible damage to the propeller. The blades should not twist and there should be no unusual noise or binding when rotating the propeller. The deice boot should be secure and free of damage. Check prop seals for leakage.5.Engine Air Intake: Ensure that the air intake lip is free of dents, exhaust leaks, and cracks. The engine air intake should be unobstructed. Check that the ice vane in the intake is fully retracted; the bypass door under the cowling should be fully retracted and flush with the cowling surface.6.Right Exhaust Stack: Check that the exhaust stack is free of cracks and secure. There should be no cracks in any weld of the exhaust pipe. Check condition of scupper inside the stack



7A

7B

8

9

10

12

13




B Left Engine (continued)7.Right Cowling: Open cowling door (7A) and check for fuel, oil, and/or bleed air leaks. Check condition of linkages, hoses, and accessories. Ensure nuts or bolts are properly safetied. Ensure generator blast tube air scoop is clear. Check general condition of engine. Close cowling door and ensure it is secure and flush with surface. Check alignment marks on forward (upper) cowling bolts for proper indication (7B).8.Auxiliary Fuel Tank: Inspect the wing in the area of the auxiliary fuel tank, and ensure the fuel cap is secure and locked with the tab pointing aft.9.Inboard Deice Boot: The deice boot should be securely attached to the leading edge. Check for significant scratches or cuts in the boot. Check the stall strip attached to the boot for security.10.Heat Exchanger Inlet: Check the heat exchanger inlet and outlet for cracks or obstructions.11.Hydraulic Access Panel: Open/inspect fluid level and accumulator nitrogen pressure.12.Inboard Fuel Tank Sump: Drain quick-drain valve on the underside of the auxiliary tank near the fuselage.13. Lower Antennas: Check the underside of the aircraft for signs of fuel or other leaks. Check antennas and beacon for security and condition.



1A

1B

2

3

4

5

6




C Nose1.Temperature Probe: Ensure that the security and condition of the temperature probe beneath pilot side window (1A), or on the lower fuselage on S/Ns 1439 and 1444 and subsequent (1B) when digital OAT is installed, is satisfactory.2.Pilot Windshield/Wiper: The windshield should be clean, free from cracks, discoloration, crazing, and excessive delamination. Ensure security and condition of the pilot wiper arm assembly.3.Left Access Panel: Check that the left access panel on nose section is secure.4.Air Conditioner Outlet Duct: Check that the outlet duct is clear.5. Left Pitot Tube: The pitot cover should be removed; ensure the pitot tue is secure and in good condition.6.Nose Gear/Door: Check that the nose gear door hinges are in good condition. Ensure turn limits on the nosewheel strut have not been exceeded. Metal plate should be straight and holes should be circular. Check condition of the uplock and downlock microswitches. Check tire, wheel, and strut linkages.7.Landing/Taxi Lights: Check condition and security of the landing and taxi light.8.Radome: Visually inspect the radome.9.Right Pitot Tube: The pitot cover should be removed; ensure the pitot tube is secure and in good condition.

7

8

9



10

11

12




C Nose (continued)10. Air Conditioning Ram Air Scoop Inlet: Check that the inlet duct is free of obstructions.11. Avionics or Right Access Panel: Check that the right access panel and all fasteners are securely attached.12.Copilot Windshield/Wiper: The windshield should be clean, free from cracks, discoloration, crazing, and excessive delamination. Ensure security and condition of the wiper.



1

2

3

4

5

6




D Right Engine1. Fuel Sump Strainer Drain: Six or seven fuel sumps need to be drained on each wing. Two sumps are quick drain style and four are flush-mounted. Use a fuel tester probe to push up on the spring-loaded push-to-drain type valve. Check for water and/or sediment. Dispose of samples in a fuel sample container. At this point, check the inboard fuel tank sump on the underside of the auxiliary tank near the fuselage.2.Battery Air Inlet: The battery air inlet is on the bottom of the wing between the nacelle and the fuselage. The thermostatically controlled valve should be securely in place, should not bind, and should be in the proper position for battery box temperature (fully open at 70 to 80°F and fully closed at 30°F).3.Inboard Deice Boot: The deice boot should be securely attached to the leading edge. Check for significant scratches or cuts in the boot. Check the stall strip attached to the boot for security.4.Heat Exchanger Inlet: Check the heat exchanger inlet and outlet for cracks or obstructions.5.Battery Air Outlet: Ensure that the battery outlet vent on top of the wing is unobstructed.6.Auxiliary Fuel Tank: Ensure there are no leaks. Check the fuel cap; the locking tab should be flush with the surface of the cap and pointing aft.

AIR INLET

OUTBOARD WING FUEL DRAIN

RECESSED VENT

HEAT RAM VENT

GRAVITY FLOW LINE DRAIN

OUTLET STRAINER AND DRAIN

FUEL FILLER CAP

LEADING EDGE FUEL DRAIN

FUEL FILTER DRAIN

STRAINER DRAIN

AUX FILLER CAP

FUEL PUMP DRAIN

Figure 3A-2: Fuel System



7

8A

8B

9

10

11

12

13A

13B




D Right Engine continued)7. Engine Oil: The oil filler access door is on the top left side of the engine cowling. Check quantity. A low reading may occur if the engine has been idle and oil is allowed to seep into the scavenge pump reservoir. If this occurs, motor engine two minutes and then re-check quantity. Replace dipstick; make sure caplock flange is turned clockwise to stop position.

NTTe: Service the oil system law Consumable Materials in the

Handling, Service, & Maintenance section (8) of the AFM and P&WC SB 3001.

DO NOT MIX different brands of oil (except as provided in Consumable Materials).

Normal operating range is FULL to 4 U.S.quarts low. Maximum oil consumption is 1 quart in 10 hours of operation.

8. Left Cowling: Open cowling door (8A) and check for fuel, oil, and/or bleed air leaks. Check condition of linkages, hoses, and accessories. Check aircraft compressor belts on the right engine accessory case for wear. Ensure nuts or bolts are properly safetied. Check general condition of engine. Close cowling door and ensure it is secure and flush with surface. Check alignment marks on forward (upper) cowling bolts for proper indication (8B).9. Left Exhaust Stack: Check that the exhaust stack is free of cracks and secure. There should be no cracks in any weld of the exhaust pipe. Check condition of scupper inside the stack. Removeinlet cover and prop restraints.10. Propeller: Ensure there are no cracks, nicks, corrosion, or other possible damage to the propeller. The blades should not twist and there should be no unusual noise or binding when rotating the propeller. The deice boot should be secure and free of damage. Check prop seals for no leakage.11. Engine Air Intake: Ensure that the air intake lip is free of dents, exhaust leaks, and cracks. The engine air intake should be unobstructed. Check that the ice vane in the intake are fully retracted; the bypass door under the cowling should be fully retracted and flush with the cowling surface.12. Right Exhaust Stack: Check that the exhaust stack is free of cracks and secure. There should be no cracks in any weld of the exhaust pipe. Check condition of scupper inside the stack.13. Right Cowling: Open cowling door (13A) and check for fuel, oil, and/or bleed air leaks. Check condition of linkages, hoses, and, accessories. Ensure nuts or bolts are properly safetied. Ensure generator blast tube air scoop is clear. Check general condition of engine. Close cowling door and ensure it is secure and flush with surface. Check alignment marks on forward (upper) cowling bolts for proper indication (13B).



1

2

3

4

5

6

7




E Right Wing1. Ice Light: Check for condition and security.2. Firewall Fuel Filter Sump: Drain the flush-mounted firewall fuel filter drain; it is about even with the bottom of the firewall. Check condition of sample.3. Fuel Pump and Strainer Drains: Drain the flush-mounted nacelle fuel pump and strainer drains forward of the wheel well. Check condition of samples.4. Landing Gear/Doors: Ensure tires are in good condition, doors are secure, struts are properly inflated, and wheel wells clean and free of fluid leaks. Linkage should be secure with no signs of unusual wear or cracking. Check uplock and downlock microswitches for condition. Brake deice lines, when installed, should be secure and free of fuel, oil, and hydraulic fluid. The deice manifold should be securely attached to the axle and free of damage. Brake lines should be secure and free of fraying or leaking.5. Fire Extinguisher Pressure: Check that the fire extinguisher cylinder is properly serviced. The cylinder is located aft in the main wheel well. Reference Table 3A-4, Pressure versus Temperature chart.6. Recessed Fuel Vent/Heated Ram Air Vent: Ensure both vents under the wing are unobstructed and undamaged. The heated ram air vent is aft and outboard from the recessed fuel vent.7. Wing Fuel Sump: Drain the quick-drain wing fuel sump near the leading edge outboard of the nacelle; this sump drains the leading edge tanks.

Temp °F IndicatedPressure PSI

-40° 190 - 240-20° 220 - 2750° 250 - 315-20° 290 - 365-40° 340 - 42060° 390 - 48080° 455 - 550100° 525 - 635120° 605 - 730140° 700 - 840

Table 3A-4: Pressure Versus Temperature



8

10

11

12

13

14

15

4 16




E Right Wing (continued)8. Ground Power Unit Access: Ensure the pins for the ground power unit access are clean and in good condition. The door on the underside of the wing outboard of the engine must be properly closed and secured.9. Tie Downs and Chocks: Check that tie downs and chocks are removed.10. Outboard Deice Boot: The deice boot should be securely attached to the leading edge. Ensure that it is undamaged. Check for significant scratches or cuts in the boot. Check condition of the stall strip attached to the boot.11. Main Fuel Tank: Ensure there are no leaks in the main fuel tank. Check the fuel cap near the wing tip; the locking tab should be flush with the surface of the cap and pointing aft.12. Wing Tip/Lights: Check the light assembly for lens cracks, security, and any indication of a burned out bulb.13. Air Inlet: Check the general condition of the air inlet on the aft side of the wing tip. It should be clear and free of obstructions.14. Static Wicks: Check for three static wicks on each aileron and one on each wing tip. Ensure the static wicks are secure and free of damage.


15. Aileron: The alieron should not be below the outboard flap. It can be as much as 1/2 inch above to 1/2 inch below the flap trailing edge. Check bonding wires for secure attachment. The hinge should not have excess play.16. Outboard Wing Fuel Sump: Drain the outboard flushmounted wing sump forward of the aileron.

Static ick LimitationDue to varying configurations, consult your MEL for allowable missing wicks.



17

2 18




E Right Wing (continued)17. Flaps: Both flap panels should be fully retracted. Ensure they are not bent nor distorted. Some movement of the flaps should be possible, and there should be no sign of flap binding. Note especially the condition of rivets on top of the flaps. The inboard flap must be flush with the bottom of the fuselage; the outboard flap trailing edge may vary from 1/2 inch above to 1/2 inch below the inboard flap.18. Fuel Sump aft of wheel well: Drain the gravity feed line from wing tanks to nacelle (aft of wheel well).



1

2

3

4

5

7

8




F Tail1. Cabin Windows/Emergency Escape Hatch: Visually check the cabin windows. Inspect the emergency exit, ensure it is secure and flush with the fuselage.2. Oxygen Door: Ensure the oxygen door is properly closed and flush with fuselage.3. Right Static Ports: Check static ports are clear.4. Emergency Locator Transmitter: The Emergency Locator Transmitter (ELT) is aft of the oxygen door and static ports on the fuselage. Push the door in and ensure the switch is armed Figure 4, page 3A-38). Check the ELT antenna for condition and security.5.Ventral Fin Water Drain: Ensure the ventral fin water drain is unobstructed.6. Tie Downs and Chocks: Check that tie downs and chocks are removed.7. Lower Antennas/Beacon: Check the underside of the aircraft for leaks. Check antennas and beacon for security and condition.8. Tailcone Access Panels: Ensure that the tailcone access door and inspection panels are secure.

ANTENNA (TYPICAL)

ELT ACCESS DOOR (TYPICAL)

ON

OFF

AR

RES

ET

ELT ACCESS DOOR

ELT TRANSMITTER

ARM - OFF - ON

EMERGENCY LOCATOR TRANSMITTER SWITCH

ARM OFF ON


Figure 3A-3: Emergency Locator Transmitter



9

10

11

12

13

14

15

16




F Tail (continued)9. Empennage: Visually inspect the empennage.10. Right Stabilizer Deice Boot: The deice boot should be secure and in good condition.11. Static Wicks: Check for two static wicks on the elevator: one on the rudder and one on the vertical stabilizer.


12. Control Surfaces/Trim Tabs: Ensure trim tabs are in line with control surfaces. Check drain holes on underside of elevator. The elevator trim tab neutral position (previously set in the cockpit) can be checked by observing that the trailing edge of the trim tab aligns with the trailing edge of the elevator when the elevator is resting against the downstops. Any difference is out of tolerance.13. Navigation/Strobe Light/Antennas: Visually inspect the navigation and strobe lights.14. Left Stabilizer Deice Boot: The deice boot should be secure and in good condition.15. Left Static Ports: Check static ports for condition.16.Left Side Door/Windows: Check door hinge and latches for security, cracks, and alignments. Check door seal for deterioration and cleanliness. Check cable(s) for fraying. The steps should have no play.

Static ick LimitationDue to varying configurations, consult your MEL for number and position of static wicks.







Cabin InspectionWARIRN

Only a crewmember should close and lock the door.

Cabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOCKED/SECUREOn King Air 200C/B200C aircraft, prior to first flight of day, check cabin/cargo door annunciator circuitry in accordance with Cabin/Cargo Annunciator Check in the Pilot’s Operating Manual/AFM.

Toilet Knife Valve (Monogram Electric Toilet) . . . . . . . . . . . . . . . .OPENLoad and Baggage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SECUREWeight and Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKEDCabin Fire Extinguisher . . . . . . . . . . . . . . . . . . . . .CHARGED/SECURECabin Seats and Belts . . . . . . . . . . . . . . .SECURE/GOOD CONDITIONWindows . . . . . . . . . . . . . . . . . . . . . . . . . . CLEAN/GOOD CONDITIONPassenger Oxygen Mask Compartments . . . . . . . . . . . . . . . CHECKEDDoors . . . . . . . . . . . . . . . . . . . . . . . . . . . . VERIFY CLOSED/LATCHEDEmergency Exit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SECURE

Interior lock must be in the unlocked position to permit access from outside the aircraft in an emergency.

Passenger Briefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CONDUCT




Expanded Normal Procedures 3B

King Air 200 3B-1December 2011


ContentsExpanded Normal ProceduresNormal Procedures

Before Starting Engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3B-5Starting Engines – Battery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3B-8Starting Engines – External Power . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-10Engine Clearing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-13Before Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-13Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-15Runup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-15Before Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-17Before Takeoff – Final Items While Taking Runway . . . . . . . . . . . . . 3B-19Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-19Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-19At 10,000 Ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-20Transition Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-20Cruise . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-20Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-20Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-21Before Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-21Final Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-22Maximum Reverse Thrust Landing . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-22Balked Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-22After Clearing Runway . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-23Traffic Pattern Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-23

After Landing/Taxiback .....................................................................3B-23Before Takeoff ..................................................................................3B-23Takeoff ..............................................................................................3B-23Climb .................................................................................................3B-24Approach ..........................................................................................3B-24Before Landing .................................................................................3B-24After Landing (After Clearing Runway) .............................................3B-24

Shutdown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-25Engine Clearing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-26

King Air 2003B-2December 2011


Parking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-26Towing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-27

Figure: Aircraft Turning Radius . . . . . . . . . . . . . . . . . . . . . . . . 3B-28Mooring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-29Cabin Door Annunciator Circuitry Check (B200) . . . . . . . . . . . . . . 3B-29Cabin/Cargo Door Annunciator Circuitry Check (B200C) . . . . . . . 3B-29Heating/Cooling System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-30Radiant Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-30Defroster Air . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-30Ni-Cad Battery Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-31Sample Passenger Briefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-32

Cold Weather OperationPreflight Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-33Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-34Takeoff and Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-34Icing – AD 96-09-13 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-34

Procedures for Exiting the Severe Icing Environment . . . . . . . . . . 3B-35Icing Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-36

In Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-37Before Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-38

Hot Weather/Desert OperationsPreflight Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-39Engine Start . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-39Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-39Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-39Shutdown/Postflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-39

Expanded Normal Procedures



Expanded Normal ProceduresThis chapter outlines and expands normal operating procedures and includes applicable cautions and warnings . Also presented are parking, mooring, and towing/taxiing procedures as well as cold weather operations .







Normal Procedures

Before Starting EnginesPassenger Briefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

FAR 91 requires that the pilot-in-command or a crewmember brief the passengers on relevant safety items (e .g ., seatbelts, door operation, emergency exits, etc .) .An exception to the oral briefing rule is if the pilot-in-command determines passengers are familiar with the briefing content. A printed card with required information should be available to each passenger to supplement the oral briefing. See page 3B-31 for sample briefing.

Seats/Seatbelts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .LOCKED/ONFasten seat belts and shoulder harnesses; position seat backs upright . Ensure lateral-tracking seats are in the outboard position .

Control Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REMOVE/STOWOxygen System Preflight Inspection . . . . . . . . . . . . . CONFIRM COMPLETEBrakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET

While applying brakes, push parking brake knob completely in, depress button on end of parking brake knob, and pull completely out .

Figure 3B-1: Control Locks



Pedestal Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET (BOTH)Rudder Boost Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONCabin Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESSPressurization Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SETElectric Elevator Trim Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONEFIS Power (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFEFIS Reversionary Switches (if installed) . . . . . . . . . . . . . . . . . . . .NORMAL

Gear Handle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DOWN/CB INThrottle Quadrant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET

Power Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IDLEProp Control Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FULL FORWARDCondition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL CUTOFF

Cabin Signs . . . . . . . . . . . . . . . . . . . NO SMOKE/FASTEN SEAT BELTS (RH)Environmental Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SET (RH)

Cabin Temp Mode Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFBleed Air Valve Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ENVIR OFFVent Blower Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOAft Blower Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFRadiant Heat Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFF

Lower Panel Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETBeacon/Nav Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AS REQUIREDTurn on beacon light for all engine starts .

Ice Vane T-Handles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INEngine Anti-Ice Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EXTENDMicrophone Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .NORMAL (BOTH)Oxygen Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)

Oxygen Control Handle:Passenger Manual Override . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INSystem Ready . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .PULL ON

Crew Masks (quick-donning) . . . . . . . . . . . . . . . . . . . . CHECK/100% (BOTH)Circuit Breakers (L/R) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IN (BOTH)Pilot’s Static Air . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NORMAL (RH)

NTTe: The engine ice vanes should be extended for all ground operations to minimize ingestion of ground debris . Turn engine anti-ice OFF, when required, to maintain oil temperature within limits (B200 only) .




Hot Battery Bus/Fuel Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKRefer to Preflight section for Hot Battery bus Fuel Panel check procedure.

Battery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON/CHECKBattery Voltage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23 VDC MINIMUMCurrent Limiters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK

Check that voltage is a minimum of 23V for battery start and 20V for external power start . No voltage indicates current limiter is out .

Pilot’s Instrument Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKCompass Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SLAVEDEFIS Power (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK

Test Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . HOLD FOR 5 SECONDSGreen TEST annunciator illuminates for 5 seconds.

Test Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RELEASE TOOFFFuel Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKStall Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TEST (RH)Fire Detectors/Extinguishers . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TEST (RH)Overhead Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETLanding Gear Handle Lights . . . . . . . . . . . . . . . . . . . . .CHECK TOP/BOTTOMHydraulic Fluid Sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .TEST

Fluid sensor annunciator illuminates during test .

Weight and Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKTOLD Card . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE/BUGS SET (BOTH)Emergency Exit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SECURE/UNLOCKBefore Starting Engines Checklist . . . . . . . . . . . . . . . . . . . . . . . COMPLETE



Starting Engines – BatterySee Quick Reference chapter for starter/generator limitations.

CCAUTON If no ITT rise occurs within 10 seconds after moving the condition lever to LOW IDLE, move the condition lever to FUEL CUT-OFF . Allow 60 seconds for fuel to drain and starter to cool . Refer to the Engine Clearing procedure that follows the Starting Engines – External Power procedure .

CCAUTON If ITT appears likely to exceed 1,000°C, move condition lever to FUEL CUTOFF . Leave ignition and engine start switches in ON position . Continue motoring the engine to reduce ITT . Refer to Engine Clearing procedure .

During engine start, crew duties should be defined and organized. The pilot monitors ITT, N1, and 10-seconds time limit for light off; the copilot is responsible for starter time limits and all other indications or abnormalities . He provides verbal confirmation of oil pressure, ignition, and fuel pressure. This allows the pilot to concentrate on the two most important starting parameters: ITT and N1 . In addition, it prevents both pilots from looking at the same gauge at the same time and leaving other indicators unmonitored .

Cabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE/LOCK/LIGHT OUTEnsure handle is locked by attempting to open the door without pressing the release button . Check position of the safety arm and diaphragm plunger under the door step . Ensure the green marks on each of the four locking bolts aligns with black pointer visible in the observation port . On B200C aircraft, verify the orange mark on each of the six rotary cam locks align with the notch on its door frame plate . Verify the CABIN DOOR annunciator is extinguished .

Cargo Door (B200C only) . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE AND LOCKVerify the cargo door is closed and locked by checking that the upper and lower door handles are in the closed and locked position . Attempt to open cargo door latches without releasing safety lock . Verify the lower handle is closed and locked by attempting to rotate the handle without lifting the orange lock hook . Ensure that the orange index mark on each of the four rotary cam locks aligns with its notch on the door frame plate .

Rotating Beacon/Nav Lights . . . . . . . . . . . . . . . . . . . . . . . .ON/AS REQUIREDEngines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . START

Right Ignition and Engine Start Switch . . . . . . . . . . . . . . . . . . . . . . . . . . .ONCheck the following: IGNITION ON annunciator illuminates FUEL PRESS annunciator extinguishes propeller begins rotation oil pressure rises .

Right Condition Lever . . . . . . . . . . . . . . . . . . . . LOW IDLE AT 12%N1 (MIN)Move condition lever to LOW IDLE when N1 indicates 12% or above . Ensure fuel flow is between 135 and 150 pph.




ITT and N1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MONITOR/1,000°C MAXRight Oil Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60 PSI (MIN)

At low idle engine oil pressure should indicate a minimum of 60 PSI.Right Ignition and Engine Start Switch . . . . . . . . . . . .OFF AT 50% N1 (MIN)

NTTe: After aborting start attempt, allow 60 seconds delay for fuel draining, motor the engine for a minimum of 15 seconds, and allow the engine to stop completely before attempting another start .

Right Condition Lever . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HIGH IDLE

CCAUTON Ensure the right generator is off for start . Energizing the left starter with the right generator operating causes damage or failure of the right 325A current limiter . The same is true for the right starter with the left generator operating .

Ensure ITT is stabilized . To avoid excessive ITT, adjust the condition levers to a higher N1 speed (approx . 60% N1) during ground operation in high ambient temperatures, at high elevations, and during periods of high generator load .If an abnormally high ITT occurs, particularly if accompanied by an N1 decrease, turn off the generator and air conditioner compressor before attempting to accelerate the right engine .

Right Generator Switch . . . . . . . . . . . . . . . . . . . . . .HOLD IN RESET 1 SEC/ONThe right generator annunciator extinguishes; loadmeter displays high load due to battery charging . Battery charge annunciator illuminates 6 seconds after a load in excess of 7 Amps .

Right Loadmeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .BELOW 50%Right Generator Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFLeft Ignition and Engine Start Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

Check the following: IGNITION ON annunciator illuminates FUEL PRESS annunciator extinguishes propeller begins rotation oil pressure rises .

Right Generator Switch . . . . . . . . . . . . . . . . . . . . . .HOLD IN RESET 1 SEC/ONSwitch generator to ON position as left N1 accelerates through 12% .

Left Condition Lever . . . . . . . . . . . . . . . . . . . . . . . .LOW IDLE AT 12% N1 (MIN)Move condition lever to LOW IDLE when N1 indicates 12% or above . Ensure fuel flow is between 135 and 150 PPH. BATTERY CHG annunciator illuminates.

ITT and N1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MONITOR/1,000°C MAXLeft Oil Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60 PSI (MIN)

At low idle engine oil pressure should indicate a minimum of 60 PSI.

Left Ignition and Engine Start Switch . . . . . . . . . . . . . . . . . . . . . .OFF AT 50% N1

Ensure all engine gauges indicate normal .



Current Limiters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKVoltmeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27 .5 TO 29V (BOTH)Push for volts .

Left Generator . . . . . . . . . . . . . . . . . . . . . . . . . . . .HOLD IN RESET 1 SEC/ONOn S/N BB-037 and subsequent; BL-001 and subsequent, if the battery charge annunciator does not extinguish within five minutes, check battery condition. Refer to Page 3B-30 for the Ni-Cad Battery Check .

NTTe: On S/Ns BB-1439, 1444 and subsequent except 1463; BL-139 and subsequent, the generator can be left on for the second engine start .

Starting Engines – External PowerSee Limitations chapter for starter/generator limitations.

CCAUTON Do not connect an external power source to the aircraft unless battery indicates a charge of at least 20V . If battery voltage is less than 20V, recharge or replace the battery before reconnecting the ground power unit .

CCAUTON Use only an external power source fitted with an AN-type plug.

During engine start, crew duties should be defined and organized. The pilot monitors ITT, N1, and 10-seconds time limit for light off; the copilot is responsible for starter time limits and all other indications or abnormalities . He provides verbal confirmation of oil pressure, ignition, and fuel pressure. This allows the pilot to concentrate on the two most important starting parameters: ITT and N1 . In addition, it prevents both pilots from looking at the same gauge at the same time and leaving other indicators unmonitored .

Cabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE/LOCK/LIGHT OUTEnsure handle is locked by attempting to open the door without pressing the release button . Check position of the safety arm and diaphragm plunger under the door step . Ensure the green marks on each of the four locking bolts aligns with black pointer visible in the observation port . On B200C aircraft, verify the orange mark on each of the six rotary cam locks align with the notch on its door frame plate . Verify the CABIN DOOR annunciator is extinguished .




Cargo Door (B200C only) . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE AND LOCKVerify the cargo door is closed and locked by checking that the upper and lower door handles are in the closed and locked position . Attempt to open cargo door latches without releasing safety lock . Verify the lower handle is closed and locked by attempting to rotate the handle without lifting the orange lock hook . Ensure that the orange index mark on each of the four rotary cam locks aligns with its notch on the door frame plate .

AVIONICS MASTER PWR Switch . . . . . . . . . . . . . . . . . . . . . . . CONFIRM OFFGenerators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CONFIRM OFF

Ensure both generator switches are in the OFF position .

Battery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONThe battery tends to absorb transients power surges .

Battery Voltage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .MINIMUM 20VGround Power Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFF

Ground power unit should be off before connecting to the aircraft .

Ground Power Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CONNECT/ONGPU Voltage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK

CCAUTON Voltages greater than 30V for extended periods will damage the battery .

Verify GPU output voltage set at 28 .0 to 28 .4 volts (King Air B200) or 28 .0 to 28 .5 volts (King Air 200) .

Parking Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETWheel Chocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REMOVE

NTTe: If battery partially discharged, the BATTERY CHARGE annunciator illuminates approximately 6 seconds after external power is on-line . If the annunciator fails to extinguish within 5 minutes, perform the Ni-Cad Battery Check procedure .

NTTe: If the ground power unit does not have a standard AN plug, check the polarity of the plug . The positive lead from the ground power unit must connect to the center post . The negative lead must connect to the front post and a positive voltage of 24 to 28 VDC must be applied to the small polarizing pin of the airplane’s external power receptacle .

Rotating Beacon/Nav Lights . . . . . . . . . . . . . . . . . . . . . . . .ON/AS REQUIREDRight Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLEAR

Verify area near right propeller is clear before continuing start procedure .

Engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STARTRight Ignition and Start Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

Check the following: IGNITION ON annunciator illuminates



FUEL PRESS annunciator extinguishes propeller begins rotating oil pressure rises .

Right Condition Lever . . . . . . . . . . . . . . . . . . . . LOW IDLE AT 12%N1 (MIN)Move condition lever to LOW IDLE when N1 indicates 12% or above . Ensure fuel flow is between 135 and 150 PPH.

ITT and N1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MONITOR/1,000°C MAX

CCAUTON If no ITT rise occurs within 10 seconds after moving the condition lever to LOW IDLE, move the condition lever to FUEL CUT-OFF . Allow 60 seconds for fuel to drain and starter to cool . Refer to the Engine Clearing procedure that follows the Starting Engines – External Power procedure .

CCAUTON If ITT appears likely to exceed 1,000°C, move condition lever to FUEL CUTOFF . Leave ignition and engine start switches in ON position . Continue motoring the engine to reduce ITT . Refer to Engine Clearing procedure .

Right Oil Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60 PSI (MIN)At low idle engine oil pressure should indicate a minimum of 60 PSI.

Right Ignition and Engine Start Switch . . . . . . . . . . . .OFF AT 50% N1 (MIN)Right Condition Lever . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .LOW IDLELeft Ignition and Engine Start Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

Check the following: IGNITION ON annunciator illuminates FUEL PRESS annunciator extinguishes propeller begins rotation oil pressure rises .

Left Condition Lever . . . . . . . . . . . . . . . . . . . . .LOW IDLE AT 12% N1 (MIN)Move condition lever to LOW IDLE when N1 indicates 12% or above . Ensure fuel flow is between 135 and 150 PPH.

ITT and N1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MONITOR/1,000°C MAXLeft Oil Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60 PSI (MIN)

At low idle engine oil pressure should indicate a minimum of 60 PSI.Left Ignition and Engine Start Switch . . . . . . . . . . . . . . . . . . .OFF AT 50% N1

Ensure all engine gauges indicate normal .

Current Limiters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKRight Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FEATHER

NTTe: After aborting start attempt, allow 60 seconds delay for fuel draining, motor the engine for a minimum of 15 seconds, and allow the engine to stop completely before attempting another start .




Ground Power Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF/DISCONNECTAfter the second engine has been started, disconnect the ground power unit and secure the access door .

Right Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .FULL FORWARDRight Generator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .RESET/ONVoltmeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27 .5 TO 29V (BOTH)

Push for volts to check DC voltage .

Left Generator . . . . . . . . . . . . . . . . . . . . . . . . . . . .HOLD IN RESET 1 SEC/ONOn S/N BB-037 and subsequent; BL-001 and subsequent, the BATTERY CHG annunciator illuminates approximately six seconds after the generator is on line . If the annunciator does not extinguish within five minutes, check battery condition.

Starting Engines – External Power . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

Engine Clearing

CCAUTON Do not exceed starter limits .

Condition Lever . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL CUTOFFIgnition/Engine Start Switch . . . . . . . . . . . . . .STARTER ONLY 15 SECONDS

MINIMUM/40 SECONDS MAXIMUMIgnition/Engine Start Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFEngine Clearing Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

Before TaxiInverters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON/CHECK

Voltage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115 ±3%VFrequency . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 400 ±1% HzInverter to be Used . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

INST INV annunciator extinguishes. 26 VAC may be checked with torque meters .

Volt/Loadmeters . . . . . . . . . . . . . .PARALLEL WITHIN 10% OF EACH OTHERRequired Generator Voltage Output . . . . . . . . . . . . . . . . . .28 .25 ±0 .25 VDC

Avionics Master . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONEFIS Power (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONEFIS Aux Power (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONLights (external and cabin) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredEnvironmental Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SET (RH)

Bleed Air Valves Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPENENVIR OFF for more efficient cooling on the ground.



Cabin Temp Mode Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOVent Blower Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTO

During operation in AUTO, MANUAL HEAT, or MANUAL COLD, the ventilation blower operates in the LOW position . For maximum cooling, select the ventilation blower to HIGH and the aft blower to ON . With the air conditioner on, maintain at least 60% (200) or 62% (B200) N1 speed on the right engine . If below N1 minimum speed, the AIR CND N1 LOW annunciator illuminates and the air conditioner compressor clutch disengages .

Radiant Heat/Aft Blower (if installed) . . . . . . . . . . . . . . . . . . . . As RequiredUse the radiant heat system only in conjunction with manual temperature control mode .

Temperature Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . As RequiredCabin Air Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . As Required

If needed, divert cabin air flow to cockpit.

Coffee Bar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredFlight Instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check

NTTe: If excessive engine ITT occurs during any one of the following conditions, adjust the condition levers for a higher N1 speed .

When high generator loads are required . During operations at high ambient air temperatures . During operations at high field elevations

NTTe: If excessive ITTs are encountered, particularly if accompanied by a decreasing N1, the associated generator should be turned off before attempting to accelerate the engine . If the right ITT is high, also turn off the air conditioner by selecting the CABIN TEMP MODE switch to OFF .

Brake Deice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKCCAUTON

Do not leave brake deice on longer than required to check function of annunciators at ambient temperatures above 15°C .

Bleed Air Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OpenCondition Levers (if brake deice required) . . . . . . . . . . . . . . . . . . HIGH IDLE

You may see a BLEED AIR FAIL light at low idle.Pneumatic Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKBrake Deice (annunciator illuminated) . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONPneumatic Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . OBSERVE DECREASEBrake Deice (annunciator extinguished) . . . . . . . . . . . . . . . . . . . . . . . . .OFFPneumatic Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RECOVEREDCondition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .LOW IDLE




TaxiPropeller Beta Range may be used during taxi with minimum blade erosion up to the point where N1 increases . Exercise care when taxiing on unimproved surfaces . If possible, conduct engine runup on a hard surface, free of sand and gravel, to avoid pitting of propeller blades and aircraft surfaces .

Brakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK (BOTH)Both pilot and copilot apply brakes to test for normal effectiveness .

Instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NO FLAGS/CHECKThe pilot reads his flight instrument indications aloud (e.g., “Zero on the airspeed, pitch and bank, and VSI. Altimeter 29.95; 480 ft; headings three one zero once, twice, three times; turn and bank checks; no flags.”) The copilot responds with “Checked, no flags.”

Before Taxi Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

NTTe: Brake deice control valves may become inoperative if valves are not cycled periodically . One cycle of the valves is required daily regardless of the weather conditions .

RunupNose Wheel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CENTERParking Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETAutopilot (See AFM Autopilot supplement) . . . . . . . . . . . . . . . . . . . . CHECKOverspeed Governors/Rudder Boost . . . . . . . . . . . . . . . . . . . . . . . . . CHECK

Rudder Boost Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONProp Control Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FULL FORWARDCondition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .LOW IDLEProp Test Switch . . . . . . . . . . . . . . . . . . . . . . . . HOLD TOPROP GOV TESTPower Lever . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INCREASE

Check that propeller stabilizes at 1,870 ±40 RPM . Continue to increase until rudder movement is felt . Observe ITT and torque limits . As the rudder boost system operates, observe forward movement of the rudder pedal on the same side as power lever increase .

Power Lever . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IDLESecond Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . REPEAT PROCEDUREProp Test Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RELEASE

Primary Governors . . . . . . . . . . . . . . . . . . . . . . . . . EXERCISE AT 1,800 RPMSet power for 1,800 RPM on both engines; check for proper propeller response . Verify the RPM, with the propeller control positioned at the top of the red hatched area on the throttle quadrant, corresponds with the bottom of the green arc (1,600 RPM) .

Pressurization/Pneumatic/Vacuum . . . . . . . . . . . . . . . . . . .CHECK SET (RH)Bleed Air Valve Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPENCabin Pressure Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET



Adjust cabin altitude selector knob so that CABIN ALT dial indicates an altitude 500 ft below field pressure altitude. Set rate control selector knob index between 9 and 12 o’clock positions .

Pressurization Switch . . . . . . . . . . . . . . . . . . . . . HOLD IN TEST POSITIONContinue holding switch in TEST position.

Cabin Altitude Indicator . . . . . . . . . . . CHECK FOR DESCENT INDICATIONLeft Bleed Air Valve Switch . . . . . . . . . . . . . . . . . . . . . . . INSTR/ENVIR OFF

L BL AIR OFF annunciator illuminates .Pneumatic Pressure Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . .GREEN ARCGyro Suction Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . .WIDE GREEN ARCCabin altitude indicator continues to show descent .Right Bleed Air Valve Switch . . . . . . . . . . . . . . . . . . . . . . INSTR/ENVIR OFF

R BL AIR OFF annunciator illuminates .Pneumatic Pressure Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0Suction Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0BLEED AIR FAIL Annunciators . . . . . . . . . . . . . . . . . . . BOTH ILLUMINATE

Cabin altitude indicator shows climb .Left Bleed Air Valve Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPEN

L BL AIR OFF annunciator extinguishes .Pneumatic Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .GREEN ARCSuction Gauge . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .WIDE GREEN ARCBLEED AIR FAIL Annunciators . . . . . . . . . . . . . . . . . . . BOTH EXTINGUISH

Cabin altitude indicator shows descent .Right Bleed Air Valve Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPEN

R BL AIR OFF annunciator extinguishes .Pressurization Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS POSITIONCabin Pressure Controller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RESET

Set the controller to cruise altitude + 1000 ft or 500 ft above takeoff field pressure altitude – whichever is higher .

Ice Vanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKWCANTNN

Either the main or standby engine anti-ice actuator must be operational on each engine before takeoff .

Observe the following: torque drop and ITT rise in extended position rise to original torque in retracted position visually confirm ice vane bypass door extension and retraction ICE VANE annunciators .

Surface Deice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK/AS REQUIREDPlace DEICE switch in both positions: SINGLE (up)/MANUAL (down). Check deice pressure gauge. Check boots visually for inflation and hold down; inflation is six seconds for wings and four seconds for horizontal stabilizer.




Ice Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK/ON AS REQUIREDCheck and set the following as required by observing loadmeters during operation . pitot heat fuel vent heat stall vane heat windshield heat propeller deice .

NTTe: Loadmeter indications with switch operation. Load fluctuation is easier to observe on single generator .

Autofeather . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK/ARMPower levers . . . . . . . . . . . . . . . . . . . . . . . .APPROX 500 FT-LBS TORQUEAutofeather Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HOLD TOTEST

Both AUTOFEATHER annunciators illuminate .Left Engine Power Lever . . . . . . . . . . . . . . . . . . . . . . . . . .RETARD TO IDLE

At approximately 400 ft-lbs, opposite annunciator extinguishes . At approximately 200 ft-lbs, left annunciator extinguishes . Propeller starts to feather .

Left Engine Power Lever . . . . . . . . . . . . . . . . . . . . . . BACK TO 500 FT-LBSRepeat test for right engine .

Power Levers . . . . . . . . . . . . . . . . . . . . . . . .APPROX 500 FT-LBS TORQUEBoth AUTOFEATHER annunciators illuminate .

Power Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IDLE Both AUTOFEATHER annunciators extinguish .

Autofeather switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ARMRunup Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

NTTe: As propeller feathers, torque increases over 220 ft-lbs . This causes the propeller to cycle out of and then back into feather with associated on/off indications of the AUTOFEATHER annunciator .

Before Takeoff

WCANTNN The electric trim system operates only by movement of pairs of switches . Any movement of the elevator trim wheel with only one switch depressed denotes a system malfunction . If this occurs, turn off the elevator tab control switch; conduct flight only by manual trim wheel operation.

Parking Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET



Avionics/Radar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET (BOTH)Call out the navaid, frequency, and course information for primary and secondary navigation radios . Ensure DME is properly set, then set radar to standby or off, and initialize VLF .

Brake Deice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFEFIS (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .TESTElectric Elevator Trim System . . . . . . . . . . . . . . . . . . . . . . . . CHECK (BOTH)

Control Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONPilot/Copilot Switches . . . . . . . . . . . . . . . . . . . . . . . . . CHECK OPERATION

Check trim switch operation both sides . Check trim operation for proper pilot override response .

Trim Disconnect . . . . . . . . . . . . . . . . . . . . . . . . . . . DEACTIVATES SYSTEMCheck trim disconnect both sides .

Control Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF/THEN ONElevator Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON (RH)Trim Tabs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .FREE/SET

Two to three units of up trim is normally required for approximately neutral control pressure at single-engine climb speeds .

Engine Control Friction Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK/SET (RH)

Flap Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FULL DOWNCheck gauge indication for flaps FULL.

Flap Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . APPROACH/UPCheck gauge indication. Visually confirm flaps up.

Flight Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKCheck for proper direction of movement and freedom of travel .

Autofeather Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ARMManual Feathering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK

Check for positive but even response from both propellers .Fuel Quantity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKEngine Instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKFlight Instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK/SETBleed Air Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPEN (RH)Pressurization Switch/Controller . . . . . . . . . . . . . . . . . . . . . . PRESS/CHECKElectric Heat (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFAnnunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF/CONSIDEREDGenerator Loads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKEngine Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . . RETRACT/IF NOT REQUIREDFlight Director . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETCrew Briefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

See Standard Operating Procedures section for crew briefing.

Before Takeoff Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE




Before Takeoff – Final Items While Taking RunwayTransponder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON/SET (RH)Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON/As RequiredAnti-ice/Deice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

L/R pitot heat, L/R fuel vent heat, stall warning heat, “The Hot Five.”

Auto Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ARML/R IGN ON annunciators illuminate .

Radar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredHeading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK (BOTH)

Compare indicated headings with runway heading .

Before Takeoff – Final Items Checklist . . . . . . . . . . . . . . . . . . . . COMPLETE

TakeoffBrakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . HOLDPower . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET

Must be set by 65 Kts .

IGNITION ON Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF (RH)Both ignition annunciators extinguish at approximately 400 ft-lbs of torque .

Autofeather Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON (RH)Both AUTOFEATHER arm annunciators illuminate above 90% N1 .

Engine instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)Both engines instruments indicate normal .

Brakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RELEASETakeoff Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

NTTe: Increasing airspeed will cause torque and ITT increase .

NTTe: Must use the “Brake Hold” technique to meet charted data for takeoff performance .

ClimbGear . . . . . . . . . . . . . . . . . . . . . . . . . . . . POSITIVE RATE-OF-CLIMB/UP (RH)

All gear and handle lights extinguish .Landing/Taxi Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFFlaps (if extended) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .UP (RH)

Retract flaps above VYSE. Flap gauge indicates flaps up.Yaw Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON (RH)Climb Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SET (RH)

At 400 ft AGL minimum . Propeller RPM at 1,900; power set .



Prop SYNC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONAutofeather . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFWindshield Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredCabin Signs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SEATBELTS (RH)Cabin Pressurization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)

Reset controller if necessary . Adjust rate knob if necessary .

Aft Blower . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF (RH)Engine Instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)Climb Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

At 10,000 FtExterior Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredSeat Belt/No Smoking Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . .As Required

Transition AltitudeAltimeters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET 29 .92Oxygen . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKTransition Altitude Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

Cruise

WCANTNN Do not lift power levers in flight.

Cruise Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETSystems Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE Cabin Electrical Engines Environmental Fuel – monitor auxiliary fuel gauges to ensure fuel is being transferred from

the auxiliary tanks Oxygen

Cruise Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

DescentPressurization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SET (RH)

Set to airport pressure altitude plus 500 ft. Approximately 75% N1 is required to maintain pressurization schedule during descent . Refer to manufacturer’s Pilot Operating Manual or the CAE Operating Handbook for controller settings. Set rate control knob as desired.




Windshield/Anti-ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredSet to NORMAL or HI well before descent into warm, moist air to aid in defogging.

Autofeather . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ARMCabin Signs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . As Required (RH)TOLD Card . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPUTE/BUGS SET (BOTH)Altimeters . . . . . . . . . . . . . . . . . . . . . . . .SET AT TRANSITION LEVEL (BOTH)Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON/AS REQUIRED

Call out all lights that are turned on .

Descent Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

Approach

CCAUTON Avoid propeller operation in the range of 1,750 to 1,850 RPM because it may cause ILS glideslope interference .

Pressurization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)Crew Briefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

See Standard Operating Procedures section for crew briefing.Altimeters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET/CROSS-CHECK (BOTH)Cabin Signs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NO SMOKE/FSB (RH)Fuel Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .APPROACH (RH)Approach Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

Before LandingApproach Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CONFIRMLanding Gear . . . . . . . . . . . . . . . . . . . . . . . . . DOWN/THREE GREEN (BOTH)Landing Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

Under low visibility conditions, do not turn on landing and taxi lights because of light reflections.

Prop Sync . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredRadar . . . . . . . . . . . . . . . . . . . . . . . . . . . . STANDBY/ON AS REQUIRED (RH)Surface Deice (as required) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CYCLEBefore Landing Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED



Final ItemsCCAUTON

To ensure constant reversing characteristics, ensure propeller control is in FULL INCREASE RPM position .

When landing assured:Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DOWN (RH)Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NORMAL APPROACH SPEEDYaw Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF (RH)After Touchdown:Prop Levers . . . . . . . . . . . . . . . . . . . . . . . .FORWARD AT TOUCHDOWN (RH)Power Levers . . . . . . . . . . . . . . LIFT AND SELECT BETA OR GROUND FINEBrakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredFinal Items Checklists . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

Maximum Reverse Thrust LandingRefer to AFM Performance section for applicable landing distance charts .

CCAUTON If possible, move propellers out of reverse at approximately 40 Kts to minimize propeller blade erosion . Exercise care when reversing on runways with loose sand, dust, or snow on the surface . Flying gravel damages propeller blades; dust or snow may impair the pilot’s forward field of vision at low aircraft speeds.

Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .DOWNAirspeed . . . . . . . . . . . . . . . . . . . . . NORMAL LANDING/APPROACH SPEEDYaw Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFCondition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .HIGH IDLE (RH)Propeller Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .FULL FORWARD (RH)Power Levers . . . . . . . . . . . . . . . . . . . LIFT/REVERSE AFTER TOUCHDOWNBrakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredCondition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LOW IDLE (RH)Maximum Reverse Thrust Landing . . . . . . . . . . . . . . . . . . . . . . COMPLETED

Balked LandingPower . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .MAXIMUM ALLOWABLEAirspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ESTABLISH 100 KTS

When clear of obstacles, establish normal climb (VYSE) .

Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .UP (RH)Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . UPBalked Landing Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED




After Clearing RunwayAuto Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFEngine Ice Vane (B200 Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EXTENDLanding/Taxi Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredRecognition/Strobe Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFIce Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFTrim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . UPRadar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STANDBYBrake Deice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CYCLEAfter Clearing Runway Checklist . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE

Traffic Pattern ChecklistAftru LandFni/TaxFbackAuto Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFIce Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . As Required (RH)Trims . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET

Brfour TakroffCrew Briefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETERadios . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETAltitude Alerter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SET (RH)Flight Director . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKFinal Items:

Headings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECKAuto Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ARMIce Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As Required

TakroffIgnition Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF (RH)Autofeather Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON (RH)Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .SET (RH)Engine Instruments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CHECK (RH)



ClFmbGear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .UP (RH)Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .UP (RH)Landing/Taxi Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFYaw Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON (RH)Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET

AppuoachCrew Briefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETEAltimeters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SET/CHECKAutofeather . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ARMFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .APPROACH (RH)

Brfour LandFniLanding Gear . . . . . . . . . . . . . . . . . . . . . . . . . . DOWN (RH)/3 GREEN (BOTH)Landing/Taxi Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONFinal Items:

Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredYaw Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFPropellers . . . . . . . . . . . . . . . . . . . . . . .FORWARD AT TOUCHDOWN (RH)

Aftru LandFni (Aftru ClrauFni Rgnway)Auto-ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFEngine Ice Vanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .EXTENDEDLanding/Taxi Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredRecognition/Strobe Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFAnti-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFTrims . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . RESET (RH)Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .UP (RH)Radar/Transponder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STANDBYAFTER LANDING Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED




Shutdown

CCAUTON Monitor ITT during shutdown . If sustained combustion occurs, perform Engine Clearing procedure checklist . During shutdown, ensure the compressors decelerate freely. Do not close the fuel firewall shutoff valves for normal engine shutdown .

CCAUTON The standby boost pumps and crossfeed are connected to the Hot Battery bus . Failure to turn these switches OFF discharges the battery .

EFIS Power (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFEFIS Aux Power (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFParking Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SETAvionics Master . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFInverter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFAutofeather . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFCoffee Bar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF (RH)Environmental Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF (RH)

Cabin Temp Mode Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFVent Blower Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOAft Blower Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFRadiant Heat Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFElectric Heat (if installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFF

ITT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STABILIZED 1 MINUTE MINIMUMCondition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL CUT-OFF

Monitor ITT . Check for normal spool-down .

Propellers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FEATHEREnsure propeller control operation .Listen for normal spool-down .

Lights/Cabin Sign Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF (RH)Standby Pumps/Crossfeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFBattery Voltage/Current Limiters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK

Voltmeters PRESS FOR VOLTS. No voltage indicates current limiter is out.

Overhead Panel Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFOxygen Control Handle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INControl Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredBattery and Generator Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFTie-downs/Chocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredParking Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As Required

High brake temperatures may damage brakes if parking brake set .



External Covers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredPropeller Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredOil Quantity Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredShutdown Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

Engine Clearing

CCAUTON The parking brake should be left off and wheel chocks installed while the aircraft is unattended . Changes in ambient temperature can cause the brakes to release or exert excessive pressure .

Condition Lever . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CUTOFFIgnition and Engine Start Switch . . . . . . . . . . . . . . . . . . . . . .STARTER ONLY

Monitor ITT .

Ignition and Engine Start Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFEngine Clearing Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

ParkingParking Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PULL OUTWheel Chocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AS REQUIREDAircraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TIE DOWN (IF REQUIRED)Brakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DEPRESS PEDALS

To release the brakes, depress brake pedals and push parking brake control in .

Parking Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED

NTTe: Avoid setting the parking brake when the brakes are hot from severe usage or when moisture conditions and freezing temperatures could form ice locks .




Towing

CCAUTON Towing the aircraft with a tug while control locks are installed can seriously damage the steering linkage .

CCAUTON Do not tow or taxi aircraft if gear shock struts are deflated.

Tow Bar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CONNECTThe tow bar connects to the upper torque knee fitting of the nose strut. Refer to aircraft turning radius figure.

Control Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REMOVEAircraft Steering . . . . . . . . . . . . . BY HAND UNLESS CONNECTED TO TUG

Although the tug controls aircraft steering, someone should be in the pilot seat to operate the brakes in case of an emergency . Do not use propellers or control surfaces to push or move the aircraft .Exceeding the turn limit damages the nose gear and linkage . Maximum nosewheel turn angle is 48° left and right .

Towing Checklist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETED



Aircraft Turning Radius

21' 1'OUTSIDE GEAR 19' 6"

NOSE WHEEL

39' 10"RADIUS FORWING TIP




Mooring

CCAUTON High velocity winds can cause structural damage . When winds above 75 Kts are expected, move the aircraft to a safe area; hangar it if possible .

Wheel Chocks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . INSTALLTie-downs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SECURE

Secure tie-downs to three mooring eyes, one under each wing and one in the ventral fin.

Prop Restraints/Covers/Control Locks . . . . . . . . . . . . . . . . . . . . . . . SECUREThe propeller may windmill even in light winds; a windmilling propeller is a safety hazard . Prolonged windmilling at zero oil pressure can damage bearings .Secure the propeller with one blade down when mooring to allow moisture to drain from spinner .

NTTe: Do not set parking brakes during low temperatures when an accumulation of moisture may cause the brakes to freeze . Do not set brakes when they are hot from severe use .

Cabin Door Annunciator Circuitry Check (B200)Battery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONDoor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OPEN/MECHANISM LOCKED

CABIN DOOR annunciator illuminates .

Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE AND LATCH/NOT LOCKEDCABIN DOOR annunciator remains illuminated .

Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE AND LOCKVerify CABIN DOOR annunciator extinguishes .

Battery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFF

Cabin/Cargo Door Annunciator Circuitry Check (B200C)Battery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFCargo Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE AND LOCKCabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLOSED BUT NOT LOCKED

Verify that CABIN DOOR annunciator illuminates .

Cabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPENVerify CABIN DOOR annunciator extinguishes .

Battery Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON



Cabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPENCheck that CABIN DOOR annunciator illuminates .

Cabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLOSE AND LOCKVerify CABIN DOOR annunciator extinguishes .

Heating/Cooling SystemBleed Air Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPEN OR ENVIR OFF

Use ENVIR OFF for more efficient cooling on the ground.

Cabin Temperature Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOVent Blower . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTORadiant Heat or Aft Blower . . . . . . . . . . . . . . . . . . . . . . . . . . . . AS REQUIRED

Radiant heat should only be used with cabin temperature mode in manual .

Temperature Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AS REQUIREDCabin Air Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AS REQUIRED

NTTe: With the cabin temperature mode switch in AUTO, MAN HEAT, or MAN COOL, the ventilation blower operates at low speed . Placing the vent blower switch in HIGH increases air circulation . To obtain maximum cooling, place the ventilation blower in HIGH and the aft blower (if installed) ON . With air conditioning on, maintain the right engine at 60% N1 or higher . If below minimum N1 speed, the AIR COND N1 LOW annunciator illuminates and the air conditioning clutch disengages . For maximum heating, the ventilation blowers should be in HIGH and the aft blower selected OFF .

Radiant HeatOverhead radiant heat can be used in conjunction with an ground power unit to warm the cabin before engine starting . Radiant heat also provides supplemental in flight heating.

Defroster AirWindshield Defroster Air Control . . . . . . . . . . . . . . . . . . . . . . . . . . . PULL ONPilot, Copilot, and Cabin Air Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFF

Turn these controls OFF if increased defroster airflow is required.




Ni-Cad Battery CheckThe BATTERY CHARGE annunciators illuminate when there is an above normal battery charging current . After engine start, battery charge current is high and the BATTERY CHARGE annunciators illuminate.

Either Generator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .OFFDC Voltage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28 VOLTSBattery Switch . . . . . . . . . . . . . . . .OFF MOMENTARILY/NOTE LOADMETER

DECREASEIf loadmeter decrease exceeds 2 .5%:Battery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .CONTINUE CHARGINGBattery Switch . . . . . . . . . . . . . . . .OFF MOMENTARILY/NOTE LOADMETER

DECREASERepeat every 90 seconds while noting decrease in loadmeter .

BATTERY CHARGE Annunciator . . . . . . . . . . . . . . . . . . . . . . EXTINGUISHESWHEN LOADMETER DECREASE IS LESS THAN 2 .5%



Sample Passenger BriefingWelcome aboard . Our estimated time of arrival at ___________ is ___________We will be climbing to ___________ Enroute, weather is ___________Please ensure your seat belts are fastened during takeoff and landing . Please comply with the No Smoking and Seat Belt signs. We recommend for your comfort and safety that you fasten your seat belt anytime you are seated. Your aircraft has two exits: the main exit to the rear on the left side and a forward emergency exit on the right side . Operating instructions are on the exits as well as on the information card in your seat back pocket .In the unlikely event we should lose pressurization, an oxygen mask will automatically be released from the overhead compartment . Place the mask over your mouth and nose . Breathe normally . Ensure the line attached to your mask has pulled the key that allows oxygen to flow to the mask.Any crew member will be happy to assist you with any questions about the safety equipment as well as the beverage and snacks available for this flight.Thank you for flying with us. Enjoy your flight.




Cold Weather OperationRefer to the respective aircraft’s Maintenance Manual for deicing and anti-icing solutions and procedures .

Preflight InspectionSnow and ice on an airplane can seriously affect performance. Wing contour may be sufficiently altered by snow and ice accumulation where it affects wing lift qualities . Remove snow with a soft brush or mop . Do not remove snow and ice accumulations by chipping or mechanical means . Use of glycol-based deicing fluids is recommended. Deicing/anti-icing fluids conforming to specification MIL-A-8243 are recommended by the airframe manufacturer .In addition to the normal preflight inspection items, the following should be inspected during cold weather operations .

NTTe: Deicing/anti-icing fluid type and concentration and precipitation rate affect the effective treatment time . Refer to the applicable aircraft Maintenance Manual for recommended solutions and procedures for the removal of ice, snow, and frost .

All Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLEAR OF SNOW AND ICEVerify all surfaces are clear of snow and ice . Pay particular attention to the wings, horizontal stabilizer, and vertical stabilizer. Snow and ice accumulation can seriously affect aircraft performance .

Tires and Brakes . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK FOR FREEZE-UPIf tire freeze-up occurs, anti-ice solutions may be used on the tires or brakes . Do not use anti-ice solutions that contain lubricants such as oil . Use of these solutions will decrease brake effectiveness .

Vents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLEARPay particular attention to all vents, exhausts, and other openings . Remove accumulated snow or ice from these areas .

Control Surfaces and Hinges . . . . . . . . . . CLEAR OF SNOW/ FREEDOM OF MOVEMENT

Inspect all control surfaces and hinge areas for accumulated snow or ice . Control surfaces should move freely with no signs of binding . A thorough check of all flight controls should be made for complete freedom of movement.

Propellers and Hubs . . . . . . . . . . . . . . . . . . . . . . . CLEAN/FREE TO ROTATEPropeller blades and hubs should be free of ice . If engine inlet covers were not used during snow and freezing rain conditions, rotate the propellers by hand in the normal direction to ensure freedom of movement before engine start . After engine start exercise the propellers through low and high pitch, beta range, and into reverse range to flush any congealed oil through the system.



TaxiBrake Deice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

If the optional brake deice system is installed, turn it on for taxiing, landing in snow, slush, or freezing rain .

Taxiing . . . . . . . . . . . . . . . . . . . . . . . . . . . . AVOID DEEP SNOW AND SLUSHAvoid taxiing the aircraft in deep snow and slush to prevent forcing snow and slush into the brake assemblies and subsequent brake freezing .

Ground Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SLOWTaxi slowly to compensate for possible decreased braking action . Allow more clearance when maneuvering .

Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .RETRACTTaxi with the flaps retracted to prevent snow and slush thrown up by the wheels entering the flap mechanism. Thrown snow and slush can also damage the flaps’ lower surfaces.

Before Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK FOR HAZARDSEnsure the runway is clear of hazards such as snow drifts, glazed ice, and ruts .

Takeoff and FlightTakeoff Distance . . . . . . . . . . . . . . . . . . . . .ALLOW ADDITIONAL DISTANCE

When using a runway covered with snow or slush, increase takeoff distance .

Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CYCLE AT 500 FT AGLAt 500 ft AGL, cycle the landing gear to dislodge moisture on the landing gear retraction components .

Ice Vanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EXTENDWhen flying through visible moisture during takeoff, extend the inertial ice vanes to prevent engine ice ingestion .

Icing – AD 96-09-13Severe icing may result from environmental conditions outside of those for which the airplane is certificated. Flight in freezing rain, freezing drizzle, or mixed icing conditions (supercooled liquid water and ice crystals) may result in ice build-up on protected surfaces exceeding the capability of the ice protection system, or may result in ice forming aft of the protected surfaces . This ice may not be shed using the ice protection systems, and may seriously degrade the performance and controllability of the airplane .During flight, severe icing conditions that exceed those for which the airplane is certificated shall be determined by the following visual cues. If one or more of these visual cues exists, immediately request priority handling from Air Traffic Control to facilitate a route or an altitude change to exit the icing conditions . Unusually extensive ice accretated on the airframe in areas not normally

observed to collect ice . Accumulation of ice on the upper surface of the wing aft of the protected area .




Accumulation of ice on the propeller spinner farther aft than normally observed .Since the autopilot may mask tactile cues that may indicate adverse changes in handling characteristics, use of the autopilot is prohibited when any of the visual cues specified above exist, or when unusual lateral trim requirements or autopilot trim warnings are encountered while the airplane is in icing conditions .All icing detection lights must be operative prior to flight into icing conditions at night (this supersedes any relief provided by the Master Minimum Equipment List [MMEL]) .The following weather conditions may be conducive to severe in-flight icing: Visible rain at temperatures below 0°C ambient air temperature . Droplets that splash or splatter on impact at temperatures below 0°C ambient

air temperature .

Puocrdgurs fou TxFtFni thr Srvrur IcFni TnvFuonmrntThese procedures are applicable to all flight phases from takeoff to landing. Monitor the ambient air temperature . While severe icing may form at temperatures as cold as -18°C, increasing vigilance is warranted at temperatures around freezing with visible moisture present. If the visual cues specified in the Limitations Section of the AFM for identifying severe icing conditions are observed, accomplish the following: Immediately request priority handling from Air Traffic Control to facilitate a

route or an altitude change to exit the severe icing conditions in order to avoid extended exposure to flight conditions more severe than those for which the airplane is certificated.

Avoid abrupt and excessive maneuvering that may exacerbate control difficulties. Do not engage the autopilot . If the autopilot is engaged, hold the control wheel firmly and disengage the

autopilot . If an unusual roll response or uncommanded roll control movement is

observed, reduce the angle-of-attack . Do not extend flaps during extended operation in icing conditions. Operation

with flaps extended can result in a reduced wing angle-of-attack, with the possibility of ice forming on the upper surface farther aft on the wing than normal, possibly aft of the protected area .

If the flaps are extended, do not retract them until the airframe is clear of ice. Report these weather conditions to Air Traffic Control.



Icing Flight

CCAUTON Due to the distortion of the wing airfoil, stalling speeds should be expected to increase as ice accumulates . For the same reason, stall warning systems are not accurate and should not be relied on . Maintain a comfortable airspeed margin above normal stall speed when ice is on the aircraft . Maintain a minimum of 140 Kts airspeed during operation in sustained icing conditions to prevent ice accumulation on unprotected surfaces . If windshield icing occurs, reduce airspeed to 226 Kts or below .

The Beechcraft King Air 200/B200 is approved for flight in icing conditions as defined by FAR 25, Appendix C. These conditions do not include, nor were tests conducted in, all conditions that may be encountered such as freezing drizzle, mixed conditions, or conditions defined as severe.Some icing conditions not defined in FAR 25 have the potential of producing hazardous ice accumulations that exceed the capabilities of the aircraft’s ice protection capabilities and/or create unacceptable aircraft performance . Flight into icing conditions that lie outside the FAR-defined conditions is not prohibited. However, pilots must be prepared to promptly divert the flight if hazardous ice accumulations occur .Refer to Aircraft Flight Manual (AFM) Section II for icing flight limitations. Also refer to AFM Section IIIA for abnormal procedures involving ice protection equipment and operations outside the FAR 25, Appendix C icing envelope .

Before Takeoff (Runup) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . COMPLETE




In lFiht

CCAUTON If in doubt, extend ice vanes . Engine icing can occur even though no surface icing is present . If freedom from visible moisture cannot be assured and ambient air temperature is +5°C or below, activate engine anti-icing systems . Visible moisture is moisture in any form such as clouds, ice crystals, snow, rain, sleet, hail, or a combination of these . Ice vanes should be retracted at +15°C and above to ensure adequate engine oil cooling . Operation of strobe lights will sometimes show ice crystals not normally visible . Ambient temperatures can be as much as 10 to 15° colder than indicated outside air temperature .

Engine Ice Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONBefore visible moisture is encountered at ambient temperatures of +5°C and below or at night when freedom from visible moisture is not assured at +5°C and below . Operation of strobe lights will sometimes show ice crystals not normally visible .Engine Ice Vanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . EXTEND

L/R ICE VANE EXT green lights illuminate (SNs prior to BB1439 and BL139) .L/R ENG ANTI-ICE green lights illuminate (SNs BB1439 and BL139 and subsequent) .

Engine Ice Protection Operation . . . . . . . . CHECK/ NOTE TORQUE DROPIf either engine’s ice vane does not reach the selected position within 15 seconds, the respective yellow L/R ENG ICE FAIL light (SNs BB1439 and BL139 and subsequent) or the yellow L/R ICE VANE light (SNs prior to BB1439 and BL139) will illuminate . Refer to Engine Ice Vane Failure procedure in the AFM or CAE Operating Handbook .

Electrothermal Propeller Deice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AUTOThe system may be operated continuously in flight and will function automatically until the switch is turned off .Increasing engine RPM briefly relieves propeller imbalance caused by ice accumulation . Repeat as necessary .



Surface Deice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . SINGLE/MANUALWCANTNN

All components of the surface deice system must be monitored during icing flight to ensure the system is functioning normally.

CCAUTON If the deice ammeter does not indicate 14 to 18 Amps (18 to 24 Amps on four-bladed propellers) or the automatic timer fails to switch, refer to the AFM Abnormal Procedures or CAE Operating Handbook .

When 1/2 to 1 inch of ice accumulates:Surface Deice Switch . . . . . . . . . . . . . . . . . . . . . .SINGLE THEN RELEASE

Repeat as necessary .If surface deice switch SINGLE position fails:Surface Deice Switch . . . . . . . . . . . . . . . . . . . . . . . HOLD IN MANUAL FOR

SIX SECONDS THEN RELEASERepeat as necessary .

Windshield Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NORMAL/HIMaximum airspeed for effective windshield anti-icing is 226 Kts .

Left/Right Fuel Vent Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONLeft/Right Pitot Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONStall Warning Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ONIce Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As RequiredAlternate Static Air Source . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .As Required

Brfour LandFni

CCAUTON Operation of the surface deice system in ambient temperatures below -40°C can cause permanent damage to the deice boots .

Surface Deice Switch . . . . . . . . . . . . . . . . . . . . . . . SINGLE THEN RELEASEBefore landing approach cycle the wing deice boots to shed as much residual ice as possible regardless of the amount of ice remaining on the boots. Stall speeds can be expected to increase if ice is not shed from the deice boots .

Approach Speeds . . . . . . . . . . . . . INCREASES IF RESIDUAL ICE REMAINSLanding Distance . . . . . . . . . . . . . . . . . . INCREASES IF APPROACH SPEED

INCREASES




Hot Weather/Desert OperationsObserve aircraft performance limitations computed from Section V of the AFM . Temperature affects engine power, braking, takeoff distance, and climb performance . In areas of high humidity, non-metallic materials are subject to moisture absorption and increase the weight of the aircraft . In very dry areas, protect the aircraft from dust and sand .

Preflight InspectionDuring the inspection, be particularly conscious of dust and sand accumulation on components lubricated with oily or greasy lubricants .

Protective Covers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REMOVELanding Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CLEAN DUST/DIRT

Clean dust and dirt from the landing gear shock struts . Check gear doors, position switches and squat switches for condition and operation . Check tires and struts for proper inflation.

Engine Inlet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REMOVE DUST/SAND

Engine StartDuring engine starts at high outside temperatures, engine ITT is higher than normal but should remain within limits .To avoid excessive ITT, adjust the condition lever to a higher N1 speed (approximately 60%) during ground operations in high ambient temperatures, or at high elevations, and during periods of high generator load . Even higher N1 speeds may be necessary on the right engine if the air conditioner compressor is operating .If an abnormally high ITT is encountered, particularly if accompanied by an N1 decrease, the associated generator (and air conditioner compressor on the right engine) should be turned off before attempting to accelerate the engine .

TaxiIf the airport surfaces are sandy or dusty, avoid the exhaust wake and propwash of other aircraft .

TakeoffEnsure takeoff performance is adequate for the conditions and runway length .

Shutdown/Postflight Install all aircraft protective covers .

Do not allow sand or dust to enter fuel tanks while refueling .

Do not leave reflective objects in the cockpit on the glareshield. Reflected

heat can distort the windshield optical properties .




Standard Operating Procedures

King Air 200 3C-1December 2011


3CContentsStandard Operating ProceduresGeneral Information

Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-5Flow Patterns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-5Checklists . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-5

Omission of Checklists . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-5Challenge/No Response . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-6

Abnormal/Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-6Time Critical Situations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-6Aborted Takeoffs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-7Critical Malfunctions in Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-7Non-Critical Malfunctions in Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-7Radio Tuning and Communication . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-7Altitude Assignment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-8Pre-Departure Briefings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-8Advising of Aircraft Configuration Change . . . . . . . . . . . . . . . . . . . . . 3C-8Transitioning from Instrument to Visual Conditions . . . . . . . . . . . . . . 3C-8

Phase of Flight SOPHolding Short . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3C-9Takeoff Roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-10Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-11Cruise . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-13Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-14Precision Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-16Precision Missed Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-20Precision Approach Deviations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-22Non Precision Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-23Non-Precision Approach Deviations . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-27Visual Traffic Patterns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-28Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3C-30

King Air 2003C-2December 2011






Standard Operating ProceduresCAE strongly supports the premise that the disciplined use of well-developed Standard Operating Procedures (SOP) is central to safe, professional aircraft operations, espe-cially in multi-crew, complex, or high performance aircraft.If your flight department has an SOP, we encourage you to use it during your training. If your flight department does not already have one, we welcome your use of the CAE SOP.Corporate pilots carefully developed this SOP. A product of their experience, it is the way CAE conducts its flight operations.The procedures described herein are specific to the King Air 200 and apply to specified phases of flight. The flight crew member designated for each step accomplishes it as indicated.







General Information

DefinitionsLH/RH – Pilot Station. Designation of seat position for accomplishing a given task because of proximity to the respective control/indicator. Regardless of PF or PM role, the pilot in that seat performs tasks and responds to checklist challenges accordingly.PF – Pilot Flying. The pilot responsible for controlling the flight of the aircraft.PIC – Pilot-in-Command. The pilot responsible for the operation and safety of an aircraft during flight time.PM – Pilot Monitoring. The pilot who is not controlling the flight of the aircraft.

Flow PatternsFlow patterns are an integral part of the SOP. Accomplish the cockpit setup for each phase of flight with a flow pattern, then refer to the checklist to verify the setup. Use normal checklists as “done lists” instead of “do lists.”Flow patterns are disciplined procedures; they require pilots to understand the aircraft systems/controls and to methodically accomplish the flow pattern.

ChecklistsUse a challenge-response method to execute any checklist. After the PF initiates the checklist, the PM challenges by reading the checklist item aloud. The PF is responsible for verifying that the items designated as PF or his seat position (i.e., LH or RH) are accomplished and for responding orally to the challenge. Items designated on the checklist as PM or by his seat position are the PM’s responsibility. The PM confirms the accomplishment of the item, then responds orally to his own challenge. In all cases, the response by either pilot is confirmed by the other and any disagreement is resolved prior to continuing the checklist.After the completion of any checklist, the PM states “___ checklist is complete.” This allows the PF to maintain situational awareness during checklist phases and prompts the PF to continue to the next checklist, if required.Effective checklists are pertinent and concise. Use them the way they are written: verbatim, smartly, and professionally.

Omission of ChecklistsWhile the PF is responsible for initiating checklists, the PM should ask the PF whether a checklist should be started if, in his opinion, a checklist is overlooked. As an expression of good crew resource management, such prompting is appropriate for any flight situation: training, operations, or checkrides.



Challenge/No ResponseIf the PM observes and challenges a flight deviation or critical situation, the PF should respond immediately. If the PF does not respond by oral communication or action, the PM must issue a second challenge that is loud and clear. If the PF does not respond after the second challenge, the PM must ensure the safety of the aircraft. The PM must announce that he is assuming control and then take the necessary actions to return the aircraft to a safe operating envelope.

Abnormal/Emergency ProceduresWhen any crewmember recognizes an abnormal or emergency condition, the PIC designates who controls the aircraft, who performs the tasks, and any items to be monitored. Following these designations, the PIC calls for the appropriate checklist. The crewmember designated on the checklist accomplishes the checklist items with the appropriate challenge/response.The pilot designated to fly the aircraft (i.e., PF) does not perform tasks that compromise this primary responsibility, regardless of whether he uses the autopilot or flies manually.Both pilots must be able to respond to an emergency situation that requires immediate corrective action without reference to a checklist. The elements of an emergency procedure that must be performed without reference to the appropriate checklist are called memory or recall items. Accomplish all other abnormal and emergency procedures while referring to the printed checklist.Accomplishing abnormal and emergency checklists differs from accomplishing normal procedure checklists in that the pilot reading the checklist states both the challenge and the response when challenging each item.When a checklist procedure calls for the movement or manipulation of controls or switches critical to safety of flight (e.g., throttles, engine fire switches, fire bottle discharge switches), the pilot performing the action obtains verification from the other pilot that he is moving the correct control or switch prior to initiating the action.Any checklist action pertaining to a specific control, switch, or equipment that is duplicated in the cockpit is read to include its relative position and the action required (e.g., “Left Throttle – IDLE; Left Boost Pump – OFF”).

NOTTE: “Control” means responsible for flight control of the aircraft, whether manual or automatic.

Time Critical SituationsWhen the aircraft, passengers, and/or crew are in jeopardy, remember three things. FLY THE AIRCRAFT – Maintain aircraft control. RECOGNIZE CHALLENGE – Analyze the situation. RESPOND – Take appropriate action.




Aborted TakeoffsThe aborted takeoff procedure is a pre-planned maneuver; both crewmembers must be aware of and briefed on the types of malfunctions that mandate an abort. Assuming the crew trains to a firmly established SOP, either crew-member may call for an abort.The PF normally commands and executes the takeoff abort for directional control problems or catastrophic malfunctions. Additionally, any indication of the following malfunctions prior to V1 is cause for an abort: engine failure engine fire.

In addition to the above, the PF usually executes an abort prior to 65 KIAS for any abnormality observed.When the PM calls an abort, the PF announces “Abort.” or “Continue.” and executes the appropriate procedure.

Critical Malfunctions in FlightIn flight, the observing crewmember positively announces a malfunction. As time permits, the other crewmember makes every effort to confirm/identify the malfunction before initiating any emergency action.If the PM is the first to observe any indication of a critical failure, he announces it and simultaneously identifies the malfunction to the PF by pointing to the indicator/annunciator.After verifying the malfunction, the PF announces his decision and commands accomplishment of any checklist memory items. The PF monitors the PM during the accomplishment of those tasks assigned to him.

Non-Critical Malfunctions in FlightProcedures for recognizing and verifying a non-critical malfunction or impending malfunction are the same as those used for time critical situations: use positive oral and graphic communication to identify and direct the proper response. Time, however, is not as critical and allows a more deliberate response to the malfunction. Always use the appropriate checklist to accomplish the corrective action.

Radio Tuning and CommunicationThe PM accomplishes navigation and communication radio tuning, identification, and ground communication. For navigation radios, the PM tunes and identifies all navigation aids. Before tuning the PF’s radios, he announces the NAVAID to be set. In tuning the primary NAVAID, the PM coordinates with the PF to ensure proper selection sequencing with the autopilot mode. After tuning and identifying the PF’s NAVAID, the PM announces “(Facility) tuned and identified.”Monitor NDB audio output anytime the NDB is in use as the NAVAID. Use the marker beacon audio as backup to visual annunciation for marker passage confirmation.



In tuning the VHF radios for ATC communication, the PM places the newly assigned frequency in the head not in use (i.e., preselected) at the time of receipt. After contact on the new frequency, the PM retains the previously assigned frequency for a reasonable time period.

Altitude AssignmentThe PM sets the assigned altitude in the altitude alerter and points to the alerter while orally repeating the altitude. The PM continues to point to the altitude alerter until the PF confirms the altitude assignment and alerter setting.

Pre-Departure BriefingsThe PIC should conduct a pre-departure briefing prior to each flight to address potential problems, weather delays, safety considerations, and operational issues. Pre-departure briefings should include all crewmembers to enhance team-building and set the tone for the flight. The briefing may be formal or informal, but should include some standard items. The acronym AWARE works well to ensure no points are missed. This is also an opportunity to brief any takeoff or departure deviations from the SOP due to weather or runway conditions.

NOTTE: The acronym AWARE stands for the following. Aircraft status Weather Airport information Route of flight Extra

Advising of Aircraft Configuration ChangeIf the PF is about to make an aircraft control or configuration change, he alerts the PM to the forthcoming change (e.g., gear and flap selections). If time permits, he also announces any abrupt flight path changes so there is always mutual understanding of the intended flight path. Time permitting, a PA announcement to the passengers precedes maneuvers involving unusual pitch or bank angles.

Transitioning from Instrument to Visual ConditionsIf Visual Meteorological Conditions (VMC) are encountered during an instrument approach, the PM normally continues to make callouts for the instrument approach being conducted. However, the PF may request a changeover to visual traffic pattern callouts.




Phase of Flight SOP

Holding Short PF PM

CALLl: “Before Takeoff checklist.” ACTIONl: Complete Before Takeoff checklist.

CALLl: “Before Takeoff checklist complete.”

Takeoff Briefing

ACTIONl: Brief the following: Assigned Runway for Takeoff Initial Heading/ Course Type of Takeoff (Rolling or Standing) Initial Altitude Airspeed Limit (if applicable) Clearance Limit Emergency Return Plan SOP Deviations.

Consider the following: Impaired Runway Conditions Weather Obstacle Clearance Instrument Departures Procedures M.E.L. Abort

Cleared for Takeoff

ACTIONl: ”Complete Takeoff checklist.”

CALLl: “Takeoff checklist complete.”

ACTIONl: Confirm Assigned Runway for Takeoff and Check Heading Indicator Agreement

CALLl: “Assigned Runway Confirmed, Heading Checked”

CALLl: “Takeoff Checklist”

acTion: Complete Takeoff Checklist

CALLl: “Takeoff Checklist Complete



Takeoff RollPF PM

Setting Takeoff Power

CALLl: “Set takeoff power.” CALLl: “Ignition lights off.” CALLl: “Autofeather lights on.”

“Power Set.”

Initial Airspeed Indication

CALLl: “Airspeed alive.”

At 65 KIAS

CALLl: “65 Kts; power set.”

At V1/VR

ACTIONl: Move hand from throttles to yoke.

ACTIONl: Rotate to approximately 7° pitch attitude for takeoff.

CALLl: “Rotate.”




ClimbPF PM

At Positive Rate of Climb

Only after PM’s call,CALLl: “Gear up.”

CALLl: “Positive rate.”

CALLl: “Gear selected up.” When gear indicates up, “Gear indicates up.”

After Gear Retraction

ACTIONl: Immediately accomplish attitude correlation check. PF’s and PM’s ADI

displays agree. Pitch and bank angles

are acceptable.CALLl: “Attitudes check.” Or,

if a fault exists, give a concise statement of the discrepancy.

At VYSE and 400 Ft Above Airport Surface (Minimum)

CALLl: “Flaps up.” (if selected.)

CALLl: “VYSE”

CALLl: “Flaps selected up.” When indicator shows UP, “Flaps indicate UP.”



Climb (continued)PF PM

at 3,000 Ft above airport Surface and clear of Traffic

CALLl: “Climb power.”

CALLl: “Climb checklist.”

ACTIONl: Set climb power.CALLl: “Climb power set.”

ACTIONl: Complete Climb checklist.

At Transition Altitude

ACTIONl: Turn recognition lights off.CALLl: “29.92 set.”

CALLl: “29.92 set.”ACTIONl: “Climb checklist complete.”




CruisePF PM

At 1,000 Ft Below Assigned Altitude

CALLl: “____ (altitude) for ____ (altitude).” (e.g., “9,000 for 10,000.”)


After Level-Off and Acceleration

CALLl: “Cruise checklist.”ACTIONl: Complete Cruise checklist.

CALLl: “Cruise checklist complete.”

Altitude Deviation in Excess of 100 Ft

CALLl: “Correcting.”CALLl: “Altitude.”



DescentPF PM

CALLl: “Descent checklist.”ACTIONl: Complete Descent

checklist.CALLl: “Descent checklist

complete.”

At 1,000 Ft Above Assigned Altitude



At Transition Level

CALLl: “Altimeter set _____.” CALLl: “Altimeter set _____.”CALLl: “Transition Level checklist

complete.”

At 10,000 Ft

CALLl: “Check.” “Speed 250 Kts.”

CALLl: “10,000 ft.”

Maintain sterile cockpit below 10,000 ft above airport surface.




Descent (continued)PF PM

At Appropriate Workload Time

REVIEW REVIEW

Review the following: approach to be executed field elevation appropriate minimum sector altitude(s) inbound leg to FAF, procedure turn direction and altitude final approach course heading and intercept altitude timing required DA/MDA MAP (non-precision) VDP special procedures (DME step-down, arc, etc.) type of approach lights in use (and radio keying procedures,

if required) missed approach procedures runway information and conditions.

ACTIONl: Brief the following: configuration approach speed minimum safe altitude approach course FAF altitude DA/MDA altitude field elevation VDP missed approach

– heading – altitude – intentions

abnormal implications.

Accomplish as many checklist items as possible. The Approach checklist must be completed prior to the initial approach fix.



Precision ApproachPF PM

Prior to Initial Approach Fix

CALLl: “Approach checklist.”

CALLl: “Flaps APPROACH.”

ACTIONl: Complete Approach checklist.

CALLl: “Approach checklist complete.”

CALLl: “Flaps APPROACH.” When flaps indicate APPROACH, “Flaps indicate APPROACH.”

At Initial Convergence of Course Deviation Bar

CALLl: “Localizer/course alive.” CALLl: “Localizer/course alive.”

At Initial Downward Movement of Glideslope Raw Data Indicator

CALLl: “Glideslope alive.” CALLl: “Glideslope alive.”

When Annunciators Indicate Localizer Capture

CALLl: “Localizer captured.” CALLl: “Localizer captured.”




Precision Approach (continued)PF PM

At One Dot From Glideslope Intercept

CALLl: “Gear down. Before Landing checklist.”

CALLl: “One dot.”

CALLl: “Gear selected down.” When gear indicates down, “Gear indicates down.”

ACTIONl: Complete Before Landing checklist except for full flaps and autopilot/yaw damper.

When Annunciator Indicates Glideslope Capture

CALLl: “Glideslope captured.” CALLl: “Glideslope captured.”

If the VOR on the PM’s side is used for crosschecks on the intermediate segment, the PM’s localizer and glideslope status calls are accomplished at the time the PM changes to the ILS frequency. This should be no later than at completion of the FAF crosscheck, if required. The PM should tune and identify his NAV radios to the specific approach and monitor.




At FAF

CALLl: “Outer marker.” or “Final fix.”

ACTIONl: Start timing. Visually crosscheck

that both altimeters agree with crossing altitude.

Set final missed approach altitude in altitude alerter.

Check PF and PM instruments.

Call FAF inbound.CALLl: “Outer marker.” or

“Final fix.” “Altitude checks.”

At 1,000 Ft Above DA(H)

CALLl: “Check.” CALLl: “1,000 ft to minimums.”

At 500 Ft Above DA(H)

CALLl: “Check.”CALLl: “500 ft to minimums.”

NOTE: An approach window has the following parameters: within one dot deflection, both LOC and GS IVSI less than 1,000 fpm IAS within VAP ±10 Kts (no less than VREF) no flight instrument flags with the landing runway or visual references not in sight landing configuration, except for full flaps (circling or single engine approaches).

When within 500 ft above touchdown, the aircraft must be within the approach window. If the aircraft is not within this window, a missed approach must be executed.









At Point Where PM Sights Runway or Visual References

CALLl: “Going visual. Land.”CALLl: “Runway (or visual

reference) _____ o’clock.”

ACTIONl: As PF goes visual, PM transitions to instruments.

At DA(H)

ACTIONl: Announce intentions.CALLl: “Going visual. Land.” or

“Missed approach.”

CALLl: “Minimums. Runway (or visual reference) _____ o’clock.”

ACTIONl: As PF goes visual, PM transitions to instruments.



Precision Missed ApproachPF PM

At DA(H)

CALLl: “Missed approach.”ACTIONl: Apply power firmly

and positively. Activate go-around mode and initially rotate the nose to the flight director go-around attitude.

CALLl: “Minimums. Missed approach.”

ACTIONl: Assist PF in setting power for go-around.


CALLl: “Gear up.”CALLl: “Positive rate.”


ACTIONl: Announce heading and altitude for missed approach.




Precision Missed Approach (continued)PF PM


CALLl: “Flaps UP.”CALLl: “Flaps selected UP.”

When flaps indicate UP, “Flaps indicate UP.”

At 1,500 Ft (Minimum) Above Airport Surface and Workload Permitting

CALLl: “Climb checklist.”ACTIONl: Complete Climb checklist.

CALLl: “Climb checklist complete.”



Precision Approach DeviationsPF PM

± One Half Dot – Glideslope

CALLl: “Correcting.”

CALLl: “One half dot (high, low) and (increasing, holding, decreasing).”

± One Half Dot – Localizer


CALLl: “One half dot (right, left) and (increasing, holding, decreasing).”

VAP ± _____


CALLl: “Speed (plus or minus) _____ (Kts) and (increasing, holding, decreasing).”

At or Below VREF


CALLl: “VREF.” or “VREF minus _____ (Kts below VREF).”

Rate of Descent Exceeds 1,000 FPM


CALLl: “Sink _____ (amount) hundred and (increasing, holding, decreasing).”




Non Precision ApproachPF PM

Prior to Initial Approach Fix

CALLl: “Approach checklist.”


ACTIONl: Complete Approach checklist to flaps.

CALLl: “Flaps selected APPROACH.” Set APPROACH flaps. When flaps indicate APPROACH, “Flaps indicate APPROACH.”

At Initial Convergence of Course Deviation Bar

CALLl: “Localizer/course alive.” CALLl: “Localizer/course alive.”

When Annunciators Indicate Course Capture

CALLl: “Localizer/course captured.” CALLl: “Localizer/course captured.”

Prior to FAF

CALLl: “Gear down. Before Landing checklist.”

CALLl: “_____ (number) miles/minutes from FAF.”


ACTIONl: Complete Before Landing checklist except for full flaps.



Non-Precision Approach (continued)PF PM

At FAF

CALLl: “Outer marker.” or “Final fix.”

CALLl: “Outer marker.” or “Final fix.

ACTIONl: Start timing. Visually crosscheck

that both altimeters agree.

Set MDA (or nearest 100 ft above) in altitude alerter.

Check PF and PM instruments.

Call “FAF inbound.”

At 1,000 Ft Above MDA

CALLl: “Check.”CALLl: “1,000 ft to minimums.”

At 500 Ft Above MDA


NOTE: An approach window has the following parameters: within one dot CDI deflection or 5° bearing IVSI less than 1,000 fpm IAS within VAP ±10 Kts (no less than VREF) no flight instrument flags with the landing runway or visual references not in sight landing configuration, except for full flaps (circling or single engine approaches).

When within 500 ft above touchdown, the aircraft must be within the approach window. If the aircraft is not within this window, a missed approach must be executed.




Non-Precision Approach (continued)PF PM

At 200 Ft Above MDA


At 100 Ft Above MDA


At MDA

CALLl: “Autopilot engage.” (as required.)

CALLl: “Minimums. _____ (time) to go.” or “Minimums. ______ (distance) to go.”

ACTIONl: Engage autopilot.

At Point Where PM Sights Runway or Visual References

CALLl: “Going visual. Land.” CALLl: “Flaps DOWN.”

(or as briefed.)

CALLl: “Runway (or visual reference) _____ o’clock.”

CALLl: “Flaps selected DOWN.” When flaps indicate DOWN, “Flaps indicate DOWN.”



Non-Precision Missed ApproachPF PM

At MAP

CALLl: Missed approach.”ACTIONl: Apply power firmly

and positively. Activate go-around mode and initially rotate the nose to the flight director go-around attitude.

CALLl: “Missed approach point. Missed approach.”

ACTIONl: Assist PF in setting power for go-around.


CALLl: “Gear up.”

CALLl: “Positive rate.”


ACTIONl: Announce heading and altitude for missed approach.


CALLl: “Flaps UP.”

CALLl: “Flaps selected UP.” When flaps indicate UP, “Flaps indicate UP.”

At 1,500 Ft (Minimum) Above Airport Surface and Workload Permitting

CALLl: “Climb checklist.”

ACTIONl: Complete Climb checklist.CALLl: “Climb checklist complete.”




Non-Precision Approach DeviationsPF PM

± One Dot – Localizer/VOR


CALLl: “One dot (right, left) and (increasing, holding, decreasing).”

±5° At or Beyond Midpoint for NDB Approach


CALLl: “_____ (degrees off course) (right, left) and (increasing, holding, decreasing).”

VAP ±


CALLl: “Speed (plus or minus) _____ (Kts) and (increasing, holding, decreasing).”

At or Below VREF


CALLl: “VREF.” or “VREF minus _____ (Kts below VREF).”

Descent is ± 200 FPM of Briefed Rate


CALLl: “Sink _______ (amount) hundred and (increasing, holding, decreasing).”



Visual Traffic PatternsPF PM

Before Pattern Entry/Downwind (1,500 Ft Above Airport Surface)

CALLl: “APPROACH checklist.” ACTIONl: Complete APPROACH checklist to flaps.

Downwind


CALLl: “Gear down. ” Before Landing checklist.”

CALLl: “Flaps selected APPROACH.” When flaps indicate APPROACH, “Flaps indicate APPROACH.”


ACTIONl: Complete Before Landing checklist except for full flaps.




Visual Traffic Patterns (continued)PF PM

At 1,000 Ft Above Airport Surface

CALLl: “Check.”CALLl: “1,000 AGL.”

CALLl: “Check.”CALLl: “500 AGL.”

CALLl: “Check.”CALLl: “Flaps DOWN.”

CALLl: “200 AGL.”




LandingPF PM

At Point on Approach When PF Sights Runway or Visual Reference (Landing Assured)

CALLl: “Going visual. Land. Flaps – DOWN.”

ACTIONl: Push autopilot disconnect.

CALLl: “Autopilot/yaw down off.”


ACTIONl: Continue with: speed check vertical speed check callouts gear down verification flap verification autopilot/yaw damper

off.CALLl: “Final gear and flaps

recheck. Before Landing checklist complete.”

At 100 Ft Above Touchdown

CALLl: CALL “100 ft.”

At 50 Ft Above Touchdown

CALLl: “50 ft.”




Landing (continued)PF PM

At Touchdown

ACTIONl: Move power levers to Beta or reverse, as required.

ACTIONl: Move propeller levers full forward.

CALLl: “Props full.”

At Propeller Reverse Minimum Speed (40 KIAS)

ACTIONl: Move power levers out of reverse.

CALLl: “40 Kts.”




Maneuver Procedures 3D

King Air 200 3D-1December 2011


ContentsManeuversTwo Engine Operation

Taxiing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-7Before Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-8Takeoff – General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-8

Normal Standing Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-9Shout Field Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-9Rolling Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-9Crosswind Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-9Takeoff Rotation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-9Rejected Takeoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-10

Initial Climbout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-10Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-10Cruise . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-10

Power Setting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-10Cabin Temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-10

Turbulent Air Penetration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-11Operation in Icing Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-11Steep Turns . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-11Stall Recognition and Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-12

Approaches to Stalls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-12Clean Configuration – Flaps and Gear Up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-12Approach/Departure Configuration – 40% Flaps and Gear Up or Down . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-12Landing Configuration – Full Flaps and Gear Down . . . . . . . . . . . . . . . . . . . . . . . . .3D-13Wind Shear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-13

Unusual Attitudes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-13Recovery from Nose-High Attitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-13Recovery from Nose-Low Attitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-14

Instrument Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-14Holding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-14Flight Director . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-14Instrument Approach Considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-15Additional Instrument Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-16

King Air 2003D-2December 2011


Normal Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-16Pressurization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-16Anti-Icing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-16Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-16Emergency Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-17

Normal Visual Approach/ Balked Landing . . . . . . . . . . . . . . . . . . . . 3D-17Short Field Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-18Balked Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-18Bounced Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-18

Approaches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-18Checklist and Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-18Typical Precision ILS Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-19Typical Non-Precision Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-19Zero Flap Approach and Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-19Go-Around/ Missed Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-20Go-Around Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-20After a Missed Approach – Proceeding for Another Approach Accomplish the following . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-21After a Missed Approach – Departing Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-21Circling Approach/ Circling Pattern . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-21

Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-21Crosswind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-21Touch-and-Go Landings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-22Contaminated Runways . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-22

After Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-22Single Engine Operation

Engine Failure or Fire after V1/VR – Takeoff Continued . . . . . . . . . . . . . . . . . . . .3D-23Single Engine ILS Approach and Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-23Single Engine Go-Around/ Missed Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3D-24

Flight ProfilesKing Air Power Settings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3D-25

Figure: Normal Takeoff .............................................................. 3D-27Figure: Short Field Takeoff ........................................................ 3D-29Figure: Wind Shear on Takeoff ................................................. 3D-31Figure: Rejected Takeoff ........................................................... 3D-33Figure: Steep Turns .................................................................. 3D-35Figure: Approaches to Stalls ..................................................... 3D-37Figure: Visual Approach/Balked Landing .................................. 3D-39Figure: Short Field Approach/Bounced Landing ....................... 3D-41Figure: Wind Shear on Approach .............................................. 3D-43

Maneuver Procedures



Figure: Two Engine ILS Approach and Landing ...................3D-45Figure: Single Engine ILS Approach and Landing ................3D-47Figure: Two Engine Non-Precision Approach

and Landing ............................................................. 3D-49Figure: Single Engine Non-Precision Approach

and Landing ..............................................................3D-51Figure: Zero Flap Approach and Landing .............................3D-53Figure: Circling Approach/Circling Pattern ............................3D-55Figure: Engine Failure After Liftoff (Takeoff Continued) ........3D-57Figure: Go Around/Missed Approach ....................................3D-59




Maneuver Procedures



ManeuversThis chapter presents descriptions of various maneuvers and techniques applicable to normal and single engine operations.




Maneuver Procedures



Two Engine Operation

TaxiingCATTOO

Never taxi with a flat tire or flat shock strut. During taxi, pay particular attention to propeller tip clearance Employ extreme caution when operating on unimproved or irregular surfaces or when high winds exist .Use of reverse range in surface areas containing loose sand or small stones may cause propeller blade erosion .Do not taxi aircraft until attitude and heading flags are out of view.

Prior to taxiing the Super King Air 200/B200, complete the Before Taxi checklist. Obtain clearance from the appropriate control agency and ensure both pilots understand the taxi route prior to aircraft movement. Visually check the area around the aircraft for ground equipment, other obstructions, and personnel.Also, visually check the passenger cabin to note that baggage and equipment are stowed, emergency exit access is cleared and unlocked, galley equipment and supplies are secured, and passengers are seated with seat belts fastened. If necessary, make a verbal or PA announcement that the aircraft is taxiing.Exercise caution when taxiing on unimproved surfaces. Perform the runup on a hard surface that is free of sand and gravel, if possible, to prevent the pitting of either the propeller blades or the aircraft surfaces.The rudder pedals provide approximately 14° left or right nosewheel travel for takeoff and landing. Braking and differential power increase the steering authority to approximately 48° left or right for low-speed taxi and maneuvering in congested areas. If taxiing in a congested area and close to other aircraft, hangars, or other obstacles, use ground personnel to ensure adequate clearance.When ready to taxi, release the parking brake. Begin taxiing with a slow straight ahead roll using just enough power to start movement. Apply power and vary propeller pitch in the Beta range, with minimum blade erosion and to the point of N1 increase, to control speed. When clear of other aircraft, check both sets of brakes as soon as possible. Use care and good judgment to avoid propeller blast to other aircraft, personnel, equipment, and buildings.Make normal turns with the nosewheel steering; however, use full rudder and inside brake as necessary to tighten a turn. Do not start turns with brakes alone or pivot the aircraft sharply on one main gear. Observe proper response in the gyro instruments during turns.If a sharp turn after moving from the parking spot is necessary, maintain above idle power until gaining sufficient speed to complete the turn with idle power. The additional speed prevents the aircraft from stopping during the turn and then requiring excess power to move again.



Select a higher N1 speed on the condition lever to avoid excessive ITT during ground operations at high ambient temperatures, at high elevations, and during periods of high generator load. Even higher right engine N1 speeds may be necessary with an operating air conditioner compressor. If encountering an abnormally high ITT, particularly if accompanied by an N1 decrease, turn off the associated generator (and air conditioner compressor on the right engine) before attempting to accelerate the engine.Both pilots should maintain good lookout discipline while taxiing. Avoid tests, checks, and paperwork activity that compromise necessary visual clearing. Accomplish the Taxi and Before Takeoff checklists when visual clearing is not compromised.Keep taxi speed to the minimum practical for safety and passenger comfort. Whenever it is necessary to prevent aircraft movement with the engines running, maintain firm pressure on the brake pedals or set the parking brake. Plan ahead; be sure that the aircraft and its pilots and passengers are ready for flight before calling for takeoff clearance.

Before TakeoffPrior to takeoff, consider the following: use of flight director power application brake release runway alignment runway length proper use of controls proper rotation gear retraction power reduction to climb power adherence to airport area speed limits icing conditions.

The PF’s takeoff briefing, in accordance with SOP, should be clear, concise, and pertinent to the specific takeoff. Set airspeed indices (“bugs”) according to the SOP. Tune and identify navigation aids; set the specific courses. Set the altitude alerter to the proper altitude. When cleared for takeoff, complete the Takeoff checklist.

Takeoff – GeneralRefer to the profile on Page 3D-27.The primary instruments for setting normal takeoff power are the torque gauges (considering ITT when at high altitudes or high ambient temperatures). Obtain the required minimum takeoff power settings from the manufacturer’s AFM or from the CAE King Air 200 Operating Handbook. The manufacturer’s AFM and Operating Manual state that this power is set statically for normal takeoffs and that charted takeoff performance is based on required torque by 65 KIAS.Torque values will increase 50 to 150 lbs (22.6 to 68.0 kg) as the aircraft accelerates and more RAM air enters the engines.

Maneuver Procedures



Noumal StandFni TakroffHold the brakes firmly and advance the power levers. Allow the engines to spool up and stabilize at 2,000 prop RPM before advancing the power levers to the desired takeoff torque considering ITT limits. When power is set, check engine instruments and smoothly release the brakes.To optimize coordination, the PNF monitors the instruments and assists with the power levers to enable the PF to concentrate on directional control. At 65 KIAS, confirm torque is equal or greater than the minimum required.At V1, the PF’s right hand moves to the control wheel in preparation for takeoff rotation.

Shogt Frld TakroffRefer to the profile on Page 3D-29.The short field take off is a standing start takeoff, In order to minimize the takeoff distance, the aircraft should be positioned at the very beginning of the runway. The flaps should be set to 40% (approach setting), and the power should be set to maximum prior to brake release. Note; the performance charts are designed to account for ram air pressure. Power should be set such that at 65 Kts the power will be maximum, without exceeding redline. There will be approximately 50 to 150 lbs of torque increase as the aircraft accelerates. After brake release, the aircraft will accelerate rapidly. At VR, rotate to approximately Ten degrees pitch attitude, Press pitch sync button, and begin to climb at V2 speed. After obstacles are cleared, or if there are no obstacles, at 50 ft, lower the nose slightly and accelerate to the Vy speed and retract flaps.

RollFni TakroffA rolling takeoff is an option when actual runway length adequately exceeds takeoff field length and obstacle clearance is not a factor. Once the aircraft is aligned with the runway, apply the brakes and advance the power levers. Release the brakes and adjust power to the takeoff setting prior to 65 KIAS. The AFM takeoff field length data and takeoff settings assume a standing start.

CuosswFnd TakroffWhen necessary, a crosswind takeoff may be combined with any other takeoff. Directional and lateral control throughout a crosswind takeoff is critical. The pilot responsible for manipulating the control wheel uses the conventional aileron into the wind technique (i.e., applying full deflection at the beginning of the takeoff roll and slowly decreasing the deflection as airspeed increases to V1).

Takroff RotatFonAt VR, smoothly rotate to a takeoff pitch attitude of approximately +7°. Smooth rotation prevents a decrease in airspeed. Early or late rotation degrades takeoff performance.



Rrjrctrd TakroffFor abort prior to V1, immediately and simultaneously apply wheelbrakes, retard power levers to idle, lift power levers, and pull aft through the beta range to full reverse. Maintain directional control. Use caution with asymmetrical reverse; directional control could be a problem.Move power levers out of the reverse range by 40 KIAS to prevent propeller erosion. However, use maximum reverse thrust to a full stop if absolutely necessary. Use reverse thrust cautiously on wet or slippery runways. Also use caution during strong crosswind conditions because reverse thrust may aggravate any weather vaning tendency. Maintain directional control with nosewheel steering to remain on the runway centerline. Remember when on ice, zero thrust is approximately at the top of the red and white reverse area on the throttle Quadrant.

Initial ClimboutOnce the vertical speed indicator and altimeter indicate a positive rate of climb, move the landing gear lever to UP. Confirm gear retraction, and monitor annunciators and engine instruments. When clear of obstacles and the airspeed increases to a minimum of VYSE, retract the flaps (if selected). Set climb power.

ClimbAfter setting the climb power and when clear of the airport traffic area, both pilots complete the Climb checklist.Climb at 160 KIAS through 10,000 ft. Several power adjustments may be necessary during climb to maintain the specified setting from the climb charts. If encountering a temperature inversion during the climb, closely monitor the climb torque, ITT, and N1 to stay within limits. Observe the differential pressure/cabin altitude and cabin vertical speed gauges for proper programming and comfort rate.

Cruise

Powru SrttFniNormally, maintain climb power at level off until accelerating to the desired cruise speed, then adjust power to the appropriate setting. During the climb and acceleration to cruise speed, monitor the ITT.

CabFn TrmpruatgurMonitor the environmental control panel to ensure proper comfort level for the passengers and crew. A temperature control knob allows the crew to set specific temperature for cockpit and cabin.

Maneuver Procedures



Turbulent Air PenetrationAlthough the aircraft is not operationally restricted in rough air, avoid flight through severe turbulence if possible.Carefully plan turbulence avoidance strategy with an understanding of mountain wave dynamics, thunderstorm characteristics, and weight versus altitude stall margins. If encountering severe turbulence, the following steps are recommended.1. Maintain 170 KIAS and turn on the ignition system.2. Set power to maintain target airspeed. Change only for extreme

airspeed variation.3. Disengage the autopilot and keep control movements moderate and

smooth. Maintain wings level and desired pitch attitude. Use attitude indicator as the primary instrument. In extreme drafts, large attitude changes may occur. Do not make sudden large control movements.After establishing trim setting for penetration speed, do not change the pitch trim.

4. Large altitude changes are possible in severe turbulence. Allow the altitude to vary to maintain the desired attitude and airspeed. Do not chase altitude or airspeed.

5. Ensure yaw damper is engaged to reduce yaw/roll oscillations.6. If penetrating turbulence with the autopilot on, engage the Soft Ride

mode. Turn off the Altitude, Speed, or Vertical Speed Hold mode.7. Turn on the FASTEN BELT sign.

Operation in Icing ConditionsThe engine ice vanes, windshield, and propeller anti-ice systems prevent the accumulation of icing; activate them prior to encountering such conditions. Turning on the wing inspection lights illuminates the wing leading edges for ice detection during night operations.Check all anti-ice/deice systems prior to flights into known icing; they must be operational. Use engine ice vanes on the ground or in the air when the air temperature is +5°C or colder and flight free of visible moisture cannot be assured.

Steep TurnsRefer to the profile on Page 3D-35.Steep turns (e.g., 45° bank) confirm the aerodynamic principle that increasing bank requires increased pitch and power to maintain altitude and airspeed. At intermediate altitudes (e.g, approximately 10,000 ft MSL), practice steep turns at 180 KIAS.The initial engine power setting is approximately 1,100 ft-lbs torque and 1,700 RPM. When passing through 30° bank, increase power setting to approximately 1,300 ft-lbs.Trim out back pressure as needed. Lead the rollout heading approximately 15°, and reduce thrust and pitch to the original settings.



Stall Recognition and Recovery

Appuoachrs to Stalls

CATTOO The discussion on Stall Recognition and Recovery appears only in the context of recovery training . Do not execute deliberate stalls in high performance aircraft unless they are part of a supervised pilot training program. Safety of flight considerations dictate the employment of utmost caution during such exercises .

Refer to the profile on Page 3D-37.Continue the approach to stall only to the first evidence of a stall (i.e., stall warning horn). At the first warning indication, initiate an immediate recovery. Do not allow the aircraft to go into a full stall.Perform the approach to stall in clean, takeoff, and landing configurations. Practice altitude should be no lower than 5,000 ft above the terrain. Before practicing approaches to stalls, clear the cockpit area of loose articles and clear the practice area.

Clean Configuration – Flaps and Gear UpWhile maintaining altitude and heading (wings level), retard power levers to 200 to 300 ft-lbs and propellers to maximum RPM. As the aircraft slows, maintain altitude with back pressure. Use trim to reduce stick force. Stop trimming at 120 KIAS and use elevator control from 120 to stall warning. Do not trim all the way to stall. At the first evidence of stall (i.e., stall warning horn sounds), perform the following.1. Advance power levers to maximum allowable power.2. Reduce pitch attitude, maintain wings level.3. As airspeed increases, adjust pitch attitude to maintain altitude.4. Reduce power to maintain desired airspeed.

Approach/Departure Configuration – 40% Flaps and Gear Up or DownEstablish a level turn using 15 to 30° bank; retard power levers to 200 to 300 ft-lbs and propellers to maximum RPM. As the aircraft slows, maintain altitude with back pressure. Use trim to reduce stick forces; however, stop trimming at 120 KIAS. At the first evidence of a stall (i.e., stall warning horn sounds), perform the following.1. Advance power levers to maximum allowable power while rolling the

wings level.2. Reduce pitch attitude, maintain wings level.3. As the airspeed increases, adjust pitch attitude to maintain altitude.4. Reduce power to maintain desired airspeed.

Maneuver Procedures



Landing Configuration – Full Flaps and Gear DownWhile maintaining altitude and heading (wings level), retard power levers to 200 to 300 ft-lbs and propellers to maximum RPM. At the first evidence of a stall (i.e., stall warning horn sounds), perform the following.1. Advance power levers to maximum power. Reduce pitch attitude,

maintain wings level.2. Select flaps 40% and accelerate to VYSE. Establish a climb at VYSE to

regain altitude.3. After a positive rate of climb is achieved, retract landing gear.4. Select flaps up at VYSE.5. Reduce power as necessary to maintain desired airspeed and altitude.

WFnd ShrauRefer to the profiles on Pages 3D-31 and 3D-43.Wind shear at low altitude can be very dangerous. It can happen in the takeoff or landing phase of flight. A change in wind speed of 15 Kts, or 500 ft/mm is defined as severe wind shear, The wind shear is frequently associated with the gust front of a thunder storm, or possibly a microburst.In the Takeoff phase of flight, the airspeed and Vertical speed trends should be monitored closely. If wind shear is detected, Maximum power should be applied (it should already be set), and pitch attitude should be increased to the maximum value without stalling the aircraft. The wings should be maintained level, No turn should be attempted during the wind shear event. The aircraft configuration should be maintained, If the gear is still down, or if the flaps are set to approach, this configuration should be maintained until clear of the wind shear event.In the approach phase of flight, at the first indication of wind shear, initiate the Wind shear escape procedure. Maximum power should be applied, and pitch attitude should be increased to the maximum value without stalling the aircraft. The wings should be maintained level. No turn should be attempted during the wind shear event. The climb should be maintained until completely clear of the wind shear event. The aircraft configuration should be maintained. If the flaps are set to approach, this configuration should be maintained until clear of the wind shear event. If the Gear is down, If the aircraft is not close to the ground, the gear should be retracted. Follow the missed approach procedure or ATC vector. Remember, a pilot report should be made.

Unusual AttitudesA number of causes (e.g., failed attitude references, autopilot malfunction, wake turbulence encounter, pilot incapacitation) may result in unusual attitudes. Prior to executing the proper recovery, confirm the unusual attitude by crosschecking the altimeters and attitude, vertical speed, and airspeed indicators.

Rrcovruy fuom Nosr-HFih AttFtgdrAfter confirming a nose-high/low airspeed condition exists, apply full power, maintain bank, or roll into a bank (of approximately 45°) and allow nose to drop towards the horizon.



As the nose passes through the horizon and accelerating to a safe speed, smoothly roll to a wings level attitude and recover to level flight. If this procedure is followed, the aircraft should maintain approximately one G.

Rrcovruy fuom Nosr-Low AttFtgdrAfter confirming a nose-low attitude, reduce thrust to idle while simultaneously rolling to a wings-level attitude. Increase pitch attitude to recover to level or climbing flight. Use caution to avoid exceeding G-limits during recovery.

Instrument Procedures

HoldFniIf endurance is a factor, determine the recommended holding speed for the existing flight weight from the Holding chart in the CAE King Air 200 Operating Handbook or the manufacturer’s Performance Manual.Slow to holding speed within three minutes prior to reaching holding fix and after notifying ATC. Holding pattern recommended entries are parallel, teardrop, and direct.Outbound timing begins over or abeam the holding fix, whichever occurs later. If the abeam position is unknown, start timing when the turn to outbound is completed.Fly the initial outbound leg for one or one and one-half minute(s) as appropriate for altitude.Inbound leg time at 14,000 ft MSL or below is one minute. Above 14,000 ft MSL, the inbound leg time is one and one-half minutes.Adjust timing of subsequent outbound legs as necessary to achieve proper inbound leg time. For a crosswind correction, double the inbound drift correction on the outbound leg.

lFiht DFurctouThe flight director is effective for making an accurate approach in adverse weather conditions. If command bars are followed precisely, the fight director computes drift corrections based on track results. These computations command slow and deliberate corrections toward interception of track and glideslope.While following the flight director commands, remember to crosscheck the raw data presentations. The flight director is extremely reliable, but the command bar(s) display computed (i.e., trend) information only.Monitor the warning lights for indications of a malfunction. If the computer is not working properly, erroneous information may appear.

Maneuver Procedures



Instugmrnt Appuoach ConsFdruatFonsConsider several factors prior to commencing an approach in a high performance aircraft. The pilot must have a thorough knowledge of the destination and alternate weather conditions before descending out of the high altitude structure. Many weather and traffic advisory sources are available, including: Flight Service Stations that provide the latest destination and

alternate weather conditions Destination Tower and/or Approach Control ARTCC where controllers can obtain information (upon request)

pertaining to traffic delays and whether aircraft are successfully completing approaches

ATIS.If weather is at or near minimums for the approaches available, review time and fuel requirements to an alternate.To continue the approach to a landing after arrival at minimums, FAR 91.175 requires the following.(c) Operation below DA or MDA. Where a DA or MDA is applicable, no

pilot may operate an aircraft, except a military aircraft of the United States, at any airport below the authorized MDA or continue an approach below the authorized DA unless:

(1) The aircraft is continuously in a position from which a descent to a landing on the intended runway can be made at a normal rate of descent using normal maneuvers, and for operations conducted under part 121 or part 135 unless that descent rate will allow touchdown to occur within the touchdown zone of the runway of intended landing:

(2) The flight visibility is not less than the visibility prescribed in the standard instrument approach being used; and

(3) Except for a Category II or Category III approach where any necessary visual reference requirements are specified by the Administrator, at least one of the following visual references for the intended runway is distinctly visible and identifiable to the pilot:

(i) The approach light system, except that the pilot may not descend below 100 ft above the touchdown zone elevation using the approach lights as a reference unless the red terminating bars or the red side row bars are also distinctly visible and identifiable.

(ii) The threshold.(iii) The threshold markings.(iv) The threshold lights.(v) The runway end identifier lights.(vi) The visual approach slope indicator.(vii) The touchdown zone or touchdown zone markings.(viii) The touchdown zone lights.(ix) The runway or runway markings.(x) The runway lights.(d) Landing. No pilot operating an aircraft, except a military aircraft of the

United States, may land that aircraft when the flight visibility is less than the visibility prescribed in the standard instrument approach procedure being used.



AddFtFonal Instugmrnt SystrmsThe following additional equipment is available on most aircraft and should be set according to company SOP: radio altimeter terrain advisory voice encoding altimeter Vertical Navigation computer controller (VNAV) long-range navigation equipment.

Normal DescentPurssguFzatFonBefore initiating a descent, set the pressurization control for landing. Verify the pressure altitude of destination and set the cabin pressure controller. Refer to the chart below to determine correct pressure altitude setting.Continue to monitor the differential pressure, cabin altitude, and cabin vertical speed throughout descent. The most comfortable condition occurs when distributing cabin descent over the majority of the aircraft descent time.

AntF-IcFniAll anti-ice systems should be on when operating in visible moisture if the air temperature is 5°C or colder.

AppuoachDouble-check landing field information and estimated landing weight; check runway requirements, determine VREF, and set airspeed indices (“bugs”) in accordance with the SOP. When descending through the transition altitude, set the altimeters to field pressure and check for agreement.

PRESSURIZATION CONTROLLER SETTING FOR LANDINGClosest Altimeter Setting Add to Airport Elevation Closest Altimeter Setting Add to Airport Elevation

28 .00 . . . . . . . . . . . . . . . . . . . . 2400 29 .50 . . . . . . . . . . . . . . . . . . . . . 900 28 .10 . . . . . . . . . . . . . . . . . . . . 2300 29 .60 . . . . . . . . . . . . . . . . . . . . . 800 28 .20 . . . . . . . . . . . . . . . . . . . . 2200 29 .70 . . . . . . . . . . . . . . . . . . . . . 700 28 .30 . . . . . . . . . . . . . . . . . . . . 2100 29 .80 . . . . . . . . . . . . . . . . . . . . . 600 28 .40 . . . . . . . . . . . . . . . . . . . . 2000 29 .90 . . . . . . . . . . . . . . . . . . . . . 500 28 .50 . . . . . . . . . . . . . . . . . . . . 1900 30 .00 . . . . . . . . . . . . . . . . . . . . . 400 28 .60 . . . . . . . . . . . . . . . . . . . . 1800 30 .10 . . . . . . . . . . . . . . . . . . . . . 300 28 .70 . . . . . . . . . . . . . . . . . . . . 1700 30 .20 . . . . . . . . . . . . . . . . . . . . . 200 28 .80 . . . . . . . . . . . . . . . . . . . . 1600 30 .30 . . . . . . . . . . . . . . . . . . . . . 100 28 .90 . . . . . . . . . . . . . . . . . . . . 1500 30 .40 . . . . . . . . . . . . . . . . . . . . . . . 0 29 .00 . . . . . . . . . . . . . . . . . . . . 1400 30 .50 . . . . . . . . . . . . . . . . . . . . .-100 29 .10 . . . . . . . . . . . . . . . . . . . . 1300 30 .60 . . . . . . . . . . . . . . . . . . . . .-200 29 .20 . . . . . . . . . . . . . . . . . . . . 1200 30 .70 . . . . . . . . . . . . . . . . . . . . .-300 29 .30 . . . . . . . . . . . . . . . . . . . . 1100 30 .80 . . . . . . . . . . . . . . . . . . . . .-400 29 .40 . . . . . . . . . . . . . . . . . . . . 1000 30 .90 . . . . . . . . . . . . . . . . . . . . .-500

Maneuver Procedures



Set the radar altimeter to either the minimum descent altitude or the decision altitude, or as desired in VFR operation for terrain proximity warning.

Emruirncy DrscrntAn emergency descent moves the aircraft rapidly from a high altitude to a lower altitude and usually occurs in conjunction with a pressurization loss. Accomplish the following in quick succession.1. Put on oxygen masks.2. Establish communications.3. Disconnect autopilot.4. Retard power levers to idle.5. Move propellers to high RPM.6. Lower flaps to approach when below 200 KIAS.7. Lower gear when below 181 KIAS.8. Lower nose to 15° below the horizon.Adjust pitch as necessary not to exceed 181 KIAS.If flying in turbulent air or if structural integrity is questionable, make the descent at a lesser and more prudent speed. The PNF should set the transponder to 7700. When conditions permit, perform the following.1. Turn on the Fasten Seat belt/No Smoking sign.2. Check oxygen availability to passengers.3. Contact ATC for assistance and instructions.The PNF monitors the descent progress, establishes the minimum altitude for terrain avoidance, and completes the appropriate checklists on command.

NNTEe: Wind factor is half the steady wind plus all of the gust; it must not exceed 10 Kts.

Normal Visual Approach/ Balked LandingRefer to the profile on Page 3D-39.The traffic pattern altitude is normally at 1,500 ft AGL. At uncontrolled airports, comply with the prescribed traffic flow for that airport.In clean configuration, slow to a minimum of 140 KIAS. The target power setting is approximately 700 ft-lbs torque and 1,700 RPM.Before entering the downwind leg, complete the Approach checklist. Set flaps at approach, and initiate the Before Landing checklist.Abeam the end of the runway, select gear down and maintain airspeed (i.e., VREF + 25 KIAS + wind factor). Complete the Before Landing checklist.Fly the base turn at a minimum of VREF + 25 KIAS + wind factor. Establish a 500 to 600 fpm descent. Upon intercepting the glide path and when landing is assured, reduce airspeed to VREF + 20 KIAS + wind factor. Select full flaps and reduce to VREF + wind factor. Cross the threshold at VREF + wind factor. Disengage yaw damper prior to landing.



Shout Frld LandFniRefer to the profile on Page 3D-41.The Short Field Landing should be performed with the condition lever set to high, the propellers set full forward (high RPM), and full flaps. The accuracy of this landing is dependant on the accuracy of the stabilized approach. The power should be reduced at approximately 200 ft above the Touch down zone, witch should result in the recommended (ref) speed crossing the runway threshold, After touching down, maximum breaking effort (without skidding the tires) and maximum reverse thrust should be used to stop the aircraft.

NNTEe: To reduce the chance of FOD damage, minimize the use of reverse thrust below 60 Kts. Particularly on unimproved runways. Also, consider the use of Engine anti-ice to reduce FOD damage.

Balkrd LandFniRefer to the profile on Page 3D-41.The Balked landing is an aborted landing attempt that happens in the landing flare, This maneuver is typically initiated below 50 ft, and involves an obstruction. The first action is to apply full power. This will require the propellers to be set to maximum. The pitch attitude should be increase to an attitude to maintain Vx (100 Kts). This attitude will be approximately 13°. After Obstructions are cleared, lower the nose to a normal climb attitude (7°), and retract flaps all the way up and retract the landing gear.

Bogncrd LandFniRefer to the profile on Page 3D-41.The procedure for the bounced landing requires a quick decision on the part of the pilot. If the bounce is minor, and in the judgment of the pilot, If the Bounce is significant, the safe thing to do is an immediate rejected landing. The recommended profile for this maneuver is identical to the Balked landing (above). If the landing attempt is to be continued, the following factors must be considered.1. Back yoke pressure must be maintained. Do not allow the nose

gear to strike the ground. This will result in a series of bounces of increasing magnitude and will likely result in damage to the aircraft.

2. A slight increase in power to help slow the rate of descent in the subsequent landing may help.

3. Maintain directional control and all crosswind corrections.4. A significant increase in landing distance will result.

ApproachesChecklist and ConfigurationFor instrument approaches with a procedure turn, complete the Approach checklist and initiate the Before Landing checklist after turning outbound from the approach fix. Lower the flaps to the approach flap schedule, and maintain the airspeed at a minimum of VREF + 25 KIAS + wind factor. The target power setting is 700 ft-lbs and 1,700 RPM for the descent.

Maneuver Procedures



If the aircraft is receiving radar vectors for an approach, initiate the Before Landing checklist and aircraft configuration changes when abeam the FAF, or three to five miles before the FAF for a straight-in approach.At uncontrolled airports, make all required position/intention reports on the appropriate Common Traffic Advisory Frequency (CTAF).

TypFcal PurcFsFon ILS AppuoachRefer to the profile on Page 3D-45.Consider an ILS approach normal when all engines, the appropriate ILS facilities, and the airborne equipment are operating normally.1. When on the localizer inbound to the FAF, ensure flaps are set for

the approach flap schedule.2. Maintain airspeed at VREF + 20 KIAS + wind factor and initiate the

Before Landing checklist.3. When the glideslope indicates one dot prior to intercept, lower the

landing gear. Complete the Before Landing checklist to the autopilot/yaw damper.

4. At glideslope intercept, begin descent.5. Maintain airspeed at VREF + 20 KIAS + wind factor.6. At or before DA/DH, establish visual contact with the runway.7. Reduce power slightly to ensure crossing the runway threshold at

VREF + wind factor. Disengage the auto-pilot/yaw damper to complete the Landing checklist.

TypFcal Non-PurcFsFon AppuoachRefer to the profile on Page 3D-49.When on the inbound course to the FAF, perform the following.1. Set flaps to the approach flap schedule and initiate the Before

Landing checklist.2. Adjust airspeed to VREF + 25 KIAS + wind factor; the power setting

should be approximately 700 ft-lbs and 1,700 RPM.3. Extend landing gear and complete the Before Landing checklist

before the FAF.4. Upon crossing the FAF, start timing, notify ATC, and descend to the

MDA while maintaining VREF + 20 KIAS + wind factor. Vertical speed in descent should normally be 1,000 to 1,200 fpm.

5. After leveling off at MDA, increase power to maintain airspeed while proceeding to the VDP or the MAP.

6. With the runway environment in sight, set flaps to landing flap schedule and disengage the autopilot/yaw damper. Maintain speed while intercepting the proper visual glidepath for landing. Cross the landing threshold at VREF + wind factor.

Zruo lap Appuoach and LandFniRefer to the profile on Page 3D-53.Two major adjustments to a normal approach are the increased landing distance and the adjusted VREF for the current gross weight.During the initial phase of the approach, maintain a minimum speed of original VREF + 30 KIAS with target power at approximately 700 ft-lbs and 1,700 RPM.



Lower the landing gear early in the approach to help control airspeed.Once established on final, reduce speed to adjusted VREF + wind factor. The power setting may be as low as idle during the descent to maintain the approach speed. Ensure that the auto-pilot/yaw damper is off and that the Approach and Landing checklist are completed.During landing, the aircraft has a tendency to float in ground effect due to the increased speed and low drag configuration. Once touchdown occurs, use propeller reversing as normal to slow the aircraft.

NNTEe: Collins flight director – Heading mode is inoperative with go-around selected.

Go-Around/ Missed ApproachRefer to the profile on Page 3D-59.Accomplish the go-around/missed approach at the DA/DH or MDA with time expired (if applicable) and runway visual reference either not in sight or not in a position from which a normal visual landing approach can be accomplished.An approach with a Visual Descent Point (VDP) positions the aircraft for a normal glideslope to landing. When an aircraft proceeds beyond the VDP without visual reference to the runway, the probability of a missed approach is increased.

Go-Around ProcedureAccomplish the following.1. Apply go-around power.2. Push the go-around button; rotate to the flight director go-around

attitude (i.e, approximately 7° nose-up).3. With airspeed at a minimum of V2, retract gear at indication of a

positive rate of climb on both altimeter and VSI.4. When clear of obstacles and when at a minimum airspeed of VYSE,

retract the flaps and accelerate to 160 KIAS minimum. Adjust pitch attitude and power as necessary.

5. Reduce power to climb power. At the relatively light gross weight at which missed approaches are normally accomplished, the aircraft accelerates quickly. Adjust pitch and power accordingly.

6. Set the flight director as required. Use the heading bug and the Heading mode to fly a desired heading, and a navigation mode and the course selector to capture a desired radial/track. After the initial fixed (i.e., 7° nose-up) climb attitude is established, variable climb attitudes may be commanded with the touch control steering button on the control wheel. Use one of the vertical modes to capture and maintain desired climbs or altitudes.

7. Confirm the level-off altitude and heading/course needed for the missed approach. Comply with the published missed approach instructions unless other directions are received from ATC.

Maneuver Procedures



After a Missed Approach – Proceeding for Another Appuoach AccomplFsh thr followFni.1. After level-off, complete the Climb checklist and maintain 140 KIAS minimum.2. Review TOLD cards and bugs for next approach. Brief the approach

and complete the Approach checklist. Maintain a minimum of 140 KIAS until on a portion of the next approach and flaps are extended.

After a Missed Approach – Departing AreaAccomplish the following.1. Accelerate to normal climb speed.2. Complete the Climb checklist.3. Follow normal climbout procedures.

CFuclFni Appuoach/ CFuclFni PattrunRefer to the profile on Page 3D-55.A circling approach is an instrument approach requiring a heading change of 30° or more to align the aircraft with the landing runway. Once under visual conditions, the circling approach is a modified version of the VFR traffic pattern.Consider such factors as turbulence, strong winds, poor visibility, and low maneuvering altitude when planning a circling approach. Plan on using a minimum circling altitude and visibility appropriate to the airspeed or approach category. The King Air 200 normally falls into Category C for circling approaches.At uncontrolled airports, observe local traffic direction and restrictions.Fly the approach with the gear down and the flaps at the approach position until at a position from which a normal descent for landing can be made. At that time, begin descent and select flaps to the landing flap position.While maneuvering during the circling approach, fly a minimum of VREF + 25 KIAS + wind factor. When on final in the landing configuration, fly at VREF + 20 KIAS + wind factor until reducing power slightly to cross the runway threshold at VREF + wind factor.

LandingWith flaps at the landing flap position, cross the threshold at 50 ft AGL with a speed of VREF + wind factor.Reduce power slowly to idle and raise the nose slightly from the attitude maintained on final approach. Maintain attitude and allow the aircraft to fly onto the runway surface.Upon touchdown, move the propellers to high RPM, use reverse, and apply braking as necessary.

CuosswFndThe maximum demonstrated crosswind for the King Air 200 is 25 KIAS. On the final approach in a crosswind, either the crab approach or the wing-down method may be used.



Do not allow the aircraft to float with power off prior to touchdown. Fly to touchdown with little flare. During roll out, hold aileron control into the wind and maintain directional control with rudder and brakes. Use propeller reverse as required.

Touch-and-Go LandingsIf practicing touch-and-go landings, preplan and brief them. Do not use reverse thrust on landing. The PNF resets the flaps to the approach flap schedule, sets the stabilizer trim in the takeoff range, and confirms these settings to the PF before advancing the power levers to takeoff power.

ContamFnatrd RgnwaysLanding on a slick surface requires careful consideration of many factors: type of runway surface, approach hazards, aircraft weight/speed, wind conditions, temperature, ice, water, and snow. Do not rely on propeller reversing to ensure a reduced stopping distance.Exercise caution when using differential reversing on a slick runway. Be prepared to return to IDLE immediately if the aircraft starts to slide sideways. Strong consideration should be given for Beta power on slippery surfaces as zero thrust is in this range.Do not attempt single engine reverse on a slick runway.If there is a possibility of hydroplaning on surface water, slow below hydroplaning speed before using the wheel-brakes. Hydroplaning speed (Vh), based on NASA test data, is shown below.For takeoff:

Vh = 9 x √ tire pressureFor landing:

Vh = 7.7 x√ tire pressureThe difference is because the wheels are rolling for takeoff and not rolling prior to landing.For aircraft with standard main wheels, takeoff Vh is 88 KIAS while landing Vh is 75 KIAS. For aircraft with high flotation wheels, takeoff Vh is 71 KIAS while landing Vh is 61 KIAS. The nose wheel is common to both types of wheels. Nosewheel Vh is 69 KIAS for takeoff and 59 KIAS for landing.If braking action starts while the tires are hydroplaning, the hydroplaning can continue at a much slower speed.

After LandingAfter clearing the runway, complete the After Landing checklist. Operate the engines at idle for at least two minutes prior to shutdown, which may include taxi time. After parking the aircraft, complete the Shutdown checklist.

Maneuver Procedures



Single Engine Operation

EniFnr aFlgur ou Fur aftru V1/VR – Takeoff ContinuedRefer to the profile on Page 3D-57.With an engine fire or failure indication after V1/VR and takeoff continued, maintain directional control using the rudder. At V1/VR, rotate the aircraft to 7° nose-up and climb at V2. If the indication occurs after exceeding V2, maintain the existing airspeed. Retract the landing gear after establishing a positive rate of climb.When clear of obstacles (no lower than 400 ft AGL), accelerate to VYSE and retract the flaps if used for the takeoff. Continue the climb to the required altitude above the takeoff field elevation.In case of engine fire, consider fighting the fire after retracting the flaps and the aircraft is at VYSE or greater.When time and conditions permit, complete the Climb checklist and the Engine Failure/Fire checklist.

SFnilr EniFnr ILS Appuoach and LandFniRefer to the profile on Page 3D-47.Fly a single engine inoperative approach essentially the same as an approach with both engines operating. On final approach, however, do not extend the flaps beyond the approach flap schedule until landing is assured.Up to the final descent point, configure the aircraft as normal with the VREF + 30 KIAS + wind factor recommended speeds. A single engine power setting is approximately double the comparable two engine settings.Lower the flaps to the landing flap position when runway proximity does not require additional engine power to overcome flap drag to arrive at the normal touchdown point. Lower landing flaps and retard the operating engine power lever to idle as the aircraft crosses the runway threshold at VREF + wind factor.If using rudder trim during approach to counter asymmetric thrust, zero the rudder trim prior to or during the landing power reduction to prevent unwanted yaw. Power reduction and flare are similar to a normal landing. Power reduction should be slower than normal to counter roll due to yaw effect. Consequently, slightly less flare than normal is necessary to prevent floating.After touchdown, apply wheelbrakes, and keep the wings level. Use reverse on the operating engine if necessary.



Single Engine Go-Around/ Missed ApproachRefer to the profile on Page 3D-59.Apply go-around power on the operating engine and push the flight director go-around button to select the Go-Around mode. Rotate to approximately 7° nose-up as commanded by the flight director. While increasing power, apply rudder pressure as necessary to counter yaw.Maintain the go-around pitch attitude and minimum airspeed of VYSE. Retract the landing gear after establishing a positive rate of climb. Climb to 400 ft AGL (minimum), then retract flaps at VYSE (minimum). Set climb power, and continue the climb on the published missed approach.When time permits, the PNF sets the PF’s heading bug on the missed approach heading and selects the requested modes on the flight director. At the appropriate time, advise ATC of the missed approach and request further clearance (e.g., another approach, diversion to the alternate airport).

Maneuver Procedures



Flight ProfilesThe following flight profiles illustrate how selected maneuvers are performed. Each maneuver is broken down into sequential events that illustrate appropriate configurations. Normal Takeoff Rejected Takeoff Steep Turns Approaches to Stalls Visual Approach/Balked Landing Two Engine ILS Approach and Landing Single Engine ILS Approach and Landing Two Engine Non-Precision Approach and Landing Single Engine Non-Precision Approach and Landing Zero Flap Approach and Landing Circling Approach/Circling Pattern Engine Failure After V1/VR

Go Around/Missed Approach

King Air Power SettingsPower settings used in this chapter (Tables 3D-1 and 3D-2) work in the aircraft as well as in the simulator.Speed varies slightly with attitudes, weight, OAT and individual aircraft. The 100 lbs (45.3 kg) torque adjustment rule works in the speed ranges used in this chapter (100 lbs (45.3 kg)) Torque = 150 fpm or 10 Kts).

Two Engines Airspeed Torque RPM Gear FlapsApproach Area High Speed 150 700 1700 UP 0Approach Area Low Speed 130 700 1700 UP 40%ILS – OM Inbound (Gear at G/S 650 fpm)

130 700 1700 DOWN 40%

Non-Precision FAF Inbound (Gear at FAF 800/1000 fpm)

130 400 1700 DOWN 40%

Steep Turns 180 1100 1700 UP 0Circle to Land/ MDA 130 1000 1700 DOWN 40%

Table 3A-1: Two Engine Power Settings

Single Engine Airspeed Torque RPM Gear FlapsApproach Area High Speed 150 1300 2000 UP 0Approach Area Low Speed 130 1300 2000 UP 40%ILS – OM Inbound (Gear at G/S 650 fpm)

130 1300 2000 DOWN 40%

Non-Precision FAF Inbound (Gear at FAF 800/1000 fpm)

130 1000 2000 DOWN 40%

Circle to Land/ MDA 130 1800 2000 DOWN 40%

Table 3A-2: Single Engine Power Settings




King Air 200December 2011

3D-27For Training Purposes Only

Maneuver Procedures

Normal Takeoff

ROLLING TAKEOFF POWER – SET T/O POWER BY 65 KIAS

1 TAKEOFF FLIGHT DIRECTOR – SET BRAKES – HOLD TAKEOFF POWER – SET BRAKES – RELEASE

AT V RROTATE TO 7 °

4

AT 3,000 FT AGL MINIMUMCLIMB CHECKLIST – COMPLETE

2

6

5

3

AT POSITIVE RATE OF CLIMB PITCH – 7° CONFIGURATION – GEAR UP

PITCH – 7° AIRSPEED – ALLOW TO INCREASETO 160KIAS COFIGURATION – FLAPS UP (IF USED) POWER SET – PROPELLER 1900 RPM / ITT 725 (77O*) OR TORQUE 2,230 FT-LBS


3D-28 For Training Purposes Only




Maneuver Procedures

Short Field Takeoff

PITCH – 7 ° AIRSPEED – ALLOW TO INCREASE

TO VYSE

CONFIGURATION – FLAPS UP POWER SET – PROPELLER 1900 RPM / ITT 725 (770*) OR TORQUE 2,230 FT-LBS

1 FLAPS APPROACH FLIGHT DIRECTOR – SET BRAKES – HOLD TAKEOFF POWER – SET BRAKES – RELEASE

AT V RROTATE TO 7

MAINTAIN V2 TO 50 FT

AT 3,000 FT AGL MINIMUMCLIMB CHECKLIST – COMPLETEACCELERATE TO 160 KTS2

5

4

2°

AT POSITIVE RATE OF CLIMB CONFIGURATION – GEAR UP

3






Maneuver Procedures

Wind Shear on Takeoff

ROLLING TAKEOFF POWER – SET T/O POWER BY 65 KIAS DETECTED

1 TAKEOFF FLIGHT DIRECTOR – SET BRAKES – HOLD TAKEOFF POWER – SET BRAKES – RELEASE

AT V RROTATE TO 7 °

4

AT 3,000 FT AGL MINIMUMCLIMB CHECKLIST – COMPLETE

2

6

5

3

MAX PITCH ATTITUDE - AVOID STALLINGWINGS LEVEL UNTIL CLEAR OF WIND SHEARDO NOT CHANGE FLAP OR LANDING GEAR CONFIGURATIONVERIFY MAXIMUM POWER APPLIED

WIND SHEAR DETECTED

RESUME NORMAL CLIMB ATTITUDE RETRACT LANDING GEAR (IF NOT ALREADY UP) RETRACT FLAPS (IF APPROPRIATE) RESUME DEPARTURE PROCEDURE SUBMIT PILOT REPORT

AFTER CLEAR OF SHEAR;






Maneuver Procedures

Rejected Takeoff

STATIC TAKEOFF FLIGHT DIRECTOR – SET BRAKES – HOLD TAKEOFF POWER – SET BRAKES – RELEASE

1

ROLLING TAKEOFF TAKEOFF POWER – SET BY 65 KTS

PRIOR TO V1 CALL "ABORT" – ABORT AS BRIEFED POWER – IDLE BRAKES – APPLY OPERATING ENGINE – MAXIMUM REVERSE (IF REQUIRED) STOP – AIRCRAFT ATC – NOTIFY

3

2






Maneuver Procedures

Steep Turns

TOLERANCES: SPEED ± 10 KIAS ALTITUDE ± 100 FT BANK ± 5° HEADING ± 10 °

THIS MANEUVER MAY BE USED FOR A 180° OR 360° TURN,AND MAY BE FOLLOWED BY A REVERSAL IN THE OPPOSITE DIRECTION.

THE PNF MAY ASSIST AS DIRECTED BY THE PF.

ALTITUDE – MAINTAINSPEED – MAINTAINATTITUDE – MAINTAIN

2 BANK – SMOOTHLY ROLL TO 45 ° ALTITUDE – MAINTAIN TRIM – AS DESIRED PITCH – INCREASE TO APPROXIMATELY 3° (AFTER 30° OF BANK) POWER – PROPELLER 1700 RPM / TORQUE 1,300 FT-LBS

4

3

CLEAN CONFIGURATIONPOWER – PROPELLER 1700 RPM / TORQUE 1,100 FT-LBSAIRSPEED – 180 KIAS ATTITUDE – LEVEL

1LEAD ROLLOUT TO ASSIGNED HEADING BYAPPROXIMATELY 15 °

WINGS – SMOOTHLY ROLL LEVELTRIM – AS REQUIREDPITCH – AS REQUIRED AIRSPEED –PROPELLER 1700 RPM / TORQUE 1,200 FT-LBS






Maneuver Procedures

Approaches to Stalls

LANDING CONFIGURATION STALL – FLAPS FULL, GEAR EXTENDEDPOWER – TORQUE 200 FT-LBS

ALTITUDE – MAINTAIN EXPECT STALL WARNING – APPROXIMATELY 80 KIAS / 8° PITCH RECOVERY

- POWER – MAXIMUM - CONFIGURATION – FLAPS APPROACH PITCH – REDUCE ANGLE OF ATTACK - CONFIGURATION – GEAR UP AT POSITIVE RATE - AIRSPEED – ACCELERATE THROUGH VYSE - ALTITUDE – MINIMUM LOSS - FLAPS – UP - AIRSPEED – ACCELERATE TO 160 KIAS

BEFORE BEGINNING STALL PRACTICE V REF – COMPUTED CLEARING TURNS – COMPLETE (IN AIRCRAFT ONLY) PROPS – FULL FORWARD

TAKEOFF CONFIGURATION STALL – FLAPS APPROACH, GEAR EXTENDEDPOWER – TORQUE 200 FT-LBS

TRIM – V REF ALTITUDE – MAINTAIN BANK – 15 TO 20° EXPECT STALL WARNING – APPROXIMATELY 90 KIAS / 8 ° PITCH RECOVERY

- POWER – MAXIMUM

- ATTITUDE – WINGS LEVEL - CONFIGURATION – (GEAR UP AT POSTIVE RATE) - AIRSPEED – ACCELERATE THROUGH VYSE - ALTITUDE – MINIMUM LOSS - FLAPS – UP - AIRSPEED – ACCELERATE TO 160 KIAS

CONSTANT ALTITUDE

MINIMUM ALTITUDE 5,000 AGL

1

2

3

CLEAN CONFIGURATION STALL SET AIRSPEED BUG - VREFPOWER – TORQUE 200 FT-LBS TRIM – VREF ALTITUDE – MAINTAIN EXPECT STALL WARNING - APPROXIMATELY 100 KIAS / 10° PITCH RECOVERY

- POWER – MAXIMUM

- ATTITUDE – MAINTAIN - ALTITUDE – MINIMUM LOSS - AIRSPEED – ACCELERATE TO 160 KIAS

NOTE: In aircraft only.Minimum altitude – 5,000 ft.Conduct clearing turns.

PITCH – REDUCE ANGLE OF ATTACK

PITCH – REDUCE ANGLE OF ATTACK

- MAINTAIN HEADING

- MAINTAIN HEADING

-

-

-

- MAINTAIN HEADING

CONSTANT ALTITUDE






Maneuver Procedures

Visual Approach/Balked Landing

THRESHOLD – BALKED LANDING POWER – MAXIMUM PITCH – 10 ° AIRSPEED – 100 KIAS CONFIGURATION – FLAP FULL /

GEAR DOWN

AT 400 FT AGL MINIMUM AIRSPEED – 160 KIAS

CLEAR OBSTACLES AIRSPEED – VYSE CONFIGURATION – FLAPS UP GEAR – UP

THRESHOLD – LANDING AIRSPEED – VREF + WIND FACTOR POWER – IDLE

TOUCHDOWN BRAKES – AS REQUIRED REVERSE – AS REQUIRED

7A

7B

8

5

6A

6B

BASE LEGSINK RATE – ESTABLISH AT 500 TO 600 FPM POWER – PROPELLER 2000 RPM / TORQUE 500 FT-LBSAIRSPEED – 130 KIAS

4

PATTERN ENTRY (DOWNWIND 1,500 FT)POWER – PROPELLER 2000 RPM / TORQUE 700 FT-LBSAIRSPEED – 150 KIASCONFIGURATION – FLAPS APPROACH APPROACH CHECKLIST – COMPLETE AIRSPEED – 140 KIAS

ABEAM RUNWAY THRESHOLDLANDING GEAR – DOWN

23

BEFORE DESCENTDESCENT CHECKLIST – COMPLETEAIRSPEED BUGS – SET TO VAP / V REF

1

LANDING ASSUREDAIRSPEED – VAPCONFIGURATION – FLAPS FULLAIRSPEED – SLOWING TO V REF + WIND






Maneuver Procedures

Short Field Approach/Bounced Landing

THRESHOLD – BOUNCED LANDING POWER – MAXIMUM PITCH – 10 ° AIRSPEED – 100 KIAS CONFIGURATION – FLAP FULL /

GEAR DOWN

AT 400 FT AGL MINIMUM AIRSPEED – 160 KIAS

CLEAR OBSTACLES, AND STABILIZED AIRSPEED – VYSE CONFIGURATION – FLAPS UP GEAR – UP

THRESHOLD – LANDING AIRSPEED – VREF + WIND FACTOR POWER – IDLE

TOUCHDOWN BRAKES – AS REQUIRED MAXIMUM REVERSE(AS REQUIRED)

7A

5

6A

6B

BASE LEGSINK RATE – ESTABLISH AT 500 TO 600 FPM POWER – PROPELLER 2000 RPM / TORQUE 500 FT-LBSAIRSPEED – 130 KIAS

4

PATTERN ENTRY (DOWNWIND 1,500 FT)POWER – PROPELLER 2000 RPM / TORQUE 700 FT-LBSAIRSPEED – 150 KIASCONFIGURATION – FLAPS APPROACH APPROACH CHECKLIST – COMPLETE AIRSPEED – 140 KIAS CONDITION LEVERS HIGH IDLE

ABEAM RUNWAY THRESHOLDLANDING GEAR – DOWN

23

BEFORE DESCENTDESCENT CHECKLIST – COMPLETEAIRSPEED BUGS – SET TO VAP / V REF

1

LANDING ASSUREDAIRSPEED – VAPCONFIGURATION – FLAPS FULLAIRSPEED – SLOWING TO VREF+ WIND

BE PREPARED FOR POSSIBLE SECOND TOUCH DOWN –CONTINUE BOUNCED LANDING PROCEDURE

7B

8B

9B

6A & 7A – LANDING6B THRU 9B – REJECTED LANDING






Maneuver Procedures

Wind Shear on Approach

1

2

3

*B200

AT FIRST INDICATION OF WIND SHEAREXECUTE ESCAPE MANEUVER

FULL POWER PROPELLERS - MAX RPM PITCH FOR MAXIMUM CLIMB - AVOID STALLING WINGS LEVEL UNTIL CLEAR OF THE WIND SHEAR FLIGHT DIRECTOR - GO AROUND

AFTER CLEAR OF WIND SHEAR EVENT; RESUME MISSED APPROACH OR ATC VECTOR SUBMIT PILOT REPORT

DO NOT RETRACT FLAPS PAST APPROACHGEAR - AS REQUIRED






Maneuver Procedures

Two Engine ILS Approach and Landing

WITHIN 3 MINUTES OF IAF POWER – PROPELLER 1700 RPM / TORQUE 700 FT-LBS CONFIGURATION – GEAR, FLAPS UPAIRSPEED – 150 KIAS

IAF OUTBOUND TIMING – START APPROACH CHECKLIST – COMPLETE CONFIGURATION – FLAPS APPROACH AIRSPEED – 130 KIAS

THRESHOLD POWER – REDUCE TO IDLE AIRSPEED – VREF + WIND FACTOR

TOUCHDOWN POWER – REVERSE AS REQUIREDBRAKES – AS REQUIRED

GLIDESLOPE INTERCEPT GEAR – DOWN BEFORE LANDING CHECKLIST – COMPLETETIMING – STARTCHECK CROSSING ALT.

AIRSPEED – VAP

1

2

6

1A2A

4 7

3AWITHIN RANGE

POWER – PROPELLER 1700 RPM / TORQUE 700 FT-LBS

CONFIGURATION – GEAR, FLAP UP AIRSPEED – 150 KIAS

3 PROCEDURE TURN INBOUND POWER – PROPELLER 2000 RPM / TORQUE 600 FT-LBSAIRSPEED – 130 KIAS

LANDING ASSURED CONFIGURATION – FLAP FULL SPEED – SLOWING TO VREF + WIND POWER – PROPELLER 2000 RPM / TORQUE 600 FT-LBS LANDING CHECKLIST FINAL ITEMS – COMPLETE

5

TERMINAL AREA POWER – PROPELLER 1700 RPM /

TORQUE 700 FT-LBS AIRSPEED – 150 KIAS

WITHIN 5 TO 10 NM OF FAF APPROACH CHECKLIST – COMPLETE POWER – PROPELLER 1700 RPM /

TORQUE 700 FT-LBS CONFIGURATION – FLAP APPROACH AIRSPEED – 130 KIAS

RADAR VECTORSRADAR VECTORS RADAR VECTORS






Maneuver Procedures

Single Engine ILS Approach and Landing

WITHIN 3 MINUTES OF IAF POWER – PROPELLER 2000 RPM / TORQUE 1,300 FT-LBS AIRSPEED – 150 KIAS CONFIGURATION – GEAR, FLAPS UP

IAF OUTBOUND TIMING – START

THRESHOLD AIRSPEED – V REF + WIND FACTOR POWER – REDUCE TO IDLE

TOUCHDOWN POWER – REVERSE AS REQUIREDBRAKES – AS REQUIREDGLIDESLOPE INTERCEPT

GEAR – DOWN TIMING – START CHECK CROSSING ALTITUDE AIRSPEED – VREF + 30

1

2

6

1A2A

4

7

3AWITHIN RANGE

POWER – PROPELLER 2000 RPM / TORQUE 1,300 FT-LBS

AIRSPEED – 150 KIAS CONFIGURATION – GEAR, FLAP UP

3 PROCEDURE TURN INBOUND CONFIGURATION – FLAPS APPROACH AIRSPEED – 130 KIAS

LANDING ASSURED SPEED – SLOWING TO V REF + WIND CONFIGURATION – FLAP FULL (OR AS BRIEFED) POWER – PROPELLER 2000 RPM / TORQUE 1,300 FT-LBS LANDING CHECKLIST FINAL ITEMS – COMPLETE

5

TERMINAL AREA WITHIN 5 TO 10 NM OF FAF CONFIGURATION – FLAP APPROACH AIRSPEED – 130 KIAS

RADAR VECTORSRADAR VECTORS RADAR VECTORS






Maneuver Procedures

Two Engine Non-Precision Approach and Landing

1

2

7

1A

2A

48

3A

WITHIN 3 MINUTES OF IAF POWER – PROPELLERS 1700 RPM / 700 FT-LBS TORQUEAIRSPEED – 150 KIAS CONFIGURATION – GEAR, FLAPS UP

IAF OUTBOUND TIMING – STARTAPPROACH CHECKLIST – COMPLETE CONFIGURATION – FLAPS APPROACH AIRSPEED – 130 KIAS

THRESHOLD AIRSPEED – VREF + WIND FACTOR POWER – REDUCE TO IDLE

TOUCHDOWNPROPELLERS – FULL FORWARDBRAKES – AS REQUIRED REVERSE – AS REQUIRED

PRIOR TO FAF CONFIGURATION – GEAR DOWN LANDING CHECKLIST – COMPLETE

AT FAFGEAR – DOWN

POWER – PROPELLERS 1700 RPM / TORQUE 400 FT-LBSTIMING – STARTDESCENT – 1,500 FPM MAXIMUM

WITHIN RANGE POWER – PROPELLERS 1700 RPM /

TORQUE 700 FT-LBS AIRSPEED – 150 KIAS CONFIGURATION – GEAR, FLAPS UP

PROCEDURE TURN INBOUND POWER – PROPELLERS 1700 RPM / TORQUE 700 FT-LBSAIRSPEED – 130 KIAS

LANDING ASSURED POWER – PROPELLERS 1700 RPM / TORQUE 700 FT-LBS CONFIGURATION – FLAPS FULL AIRSPEED – SLOWING TO VREF BEFORE LANDING CHECKLIST FINAL ITEMS – COMPLETE

TERMINAL AREA POWER – PROPELLERS 1700 RPM /


WITHIN 5 TO 10 NM OF FAF POWER – PROPELLERS 1700 RPM /

TORQUE 700 FT-LBS APPROACH CHECKLIST – COMPLETE AIRSPEED – 130 KIAS CONFIGURATION – FLAPS APPROACH

6

3

AT MDA ALTITUDE – MAINTAIN POWER – PROPELLERS 1700 RPM / TORQUE 1000 FT-LBS AIRSPEED – 130 KIAS AUTOPILOT – CONSIDER

5

RADAR VECTORS

RADAR VECTORS RADAR VECTORS






Maneuver Procedures

Single Engine Non-Precision Approach and Landing

1A

2A 3A

WITHIN RANGE POWER – PROPELLER 2000 RPM /

TORQUE 1300 FT-LBS CONFIGURATION – GEAR, FLAPS UP AIRSPEED – 150 KIAS

TERMINAL AREA POWER – PROPELLER 2000 RPM /


WITHIN 5 TO 10 NM OF FAF POWER – PROPELLER 2000 RPM /

TORQUE 1300 FT-LBS CONFIGURATION – FLAPS APPROACH AIRSPEED – 130 KIAS

1

2

7

48

WITHIN 3 MINUTES OF IAF POWER – PROPELLERS 2000 RPM / TORQUE 1,300 FT-LBS CONFIGURATION – GEAR, FLAPS UPAIRSPEED – 150 KIAS

IAF OUTBOUND TIMING – START

THRESHOLD AIRSPEED – VREF + WIND FACTOR POWER – REDUCE TO IDLE

TOUCHDOWNBRAKES – AS REQUIRED REVERSE – AS REQUIRED

AT FAFTIMING – STARTDESCENT – 1,500 FPM MAXIMUM

POWER – PROPELLERS 2000 RPM / TORQUE 1000 FT-LBSCONFIGURATION – GEAR DOWN (OR AS BRIEFED)

– SEE CAUTIONAIRSPEED – 130 KIAS

LANDING ASSURED CONFIGURATION – FLAPS FULL (OR AS BRIEFED) LANDING CHECKLIST FINAL ITEMS – COMPLETE AIRSPEED – SLOWING TO V REF + WIND

6

3

AT MDA ALTITUDE – MAINTAIN POWER – PROPELLERS 2000 RPM / TORQUE 1800 FT-LBS AIRSPEED – 130 KIAS

5

PROCEDURE TURN INBOUND CONFIGURATION – FLAPS APPROACH AIRSPEED – 130 KIAS

RADAR VECTORS

RADAR VECTORS RADAR VECTORS

CAUTION: Under some condi-tions, level flight may not bepossible with gear extended.






Maneuver Procedures

Zero Flap Approach and Landing

THRESHOLD – LANDING POWER – IDLE AIRSPEED – VREF + WIND FACTOR

TOUCHDOWN REVERSE – AS REQUIRED BRAKES – AS REQUIRED

7

5

6

BASE LEGSINK RATE – ESTABLISH AT 500 TO 600 FPM POWER – PROPELLERS 2000 RPM / TORQUE 500 FT-LBSAIRSPEED – ZERO FLAPS V REF + 10 KTS

4

PATTERN ENTRY (DOWNWIND 1,500 FT)POWER – PROPELLERS 1700 RPM / TORQUE 700 FT-LBSCONFIGURATION – FLAPS UPFLAP UP LANDING CHECKLIST – BEGINAIRSPEED – 150 KIAS

ABEAM RUNWAY THRESHOLDLANDING GEAR – DOWNPOWER – PROPELLERS 2000 RPM / TORQUE 700 FT-LBSAIRSPEED – ZERO FLAP V REF + 10 KTS

23

BEFORE DESCENTDESCENT CHECKLIST – COMPLETEAIRSPEED BUGS – SET TO V AP / V REF

1

LANDING ASSUREDFLAP UP CHECKLIST – COMPLETE SINK RATE – 900FPMAIRSPEED – SLOWING ZERO FLAP V REF + WIND FACTOR






Maneuver Procedures

Circling Approach/Circling Pattern

RECOMMENDATIONS FLAPS – APPROACH GEAR – DOWN AIRSPEED – VREF + 20 + WIND FACTOR MINIMUM

(MAINTAIN CONSTANT SPEED FOR TIMING) POWER – PROPELLERS 2000 RPM / TORQUE 1000 FT-LBS F/D ALTITUDE HOLD – SELECT F/D HEADING – SELECT USE OF AUTOPILOT IS RECOMMENDED SLIGHT ADJUSTMENTS TO TIME OR HEADING

MAY BE USED TO ADJUST FOR WIND

ENTER BASIC PATTERN AS APPROPRIATEFOR AIRCRAFT POSITION

START TIMING ABEAM APPROACH END OF RUNWAY

START TURN TO FINAL, MAXIMUM 30° BANK

WITH RUNWAY IN SIGHT AND IN POSITION TO MAKEA NORMAL DESCENT TO LANDING

LANDING CHECKLIST FINAL ITEMS – COMPLETE POWER – PROPELLERS 2000 RPM / TORQUE 700 FT-LBS DESCENT FROM MDA – BEGIN AIRSPEED – VAP IF NOT IN A POSITION TO MAKE A NORMAL LANDING

- GO-AROUND – EXECUTE AT THRESHOLD

- AIRSPEED – V REF + WIND FACTOR

NOTESBASED ON 30° BANK TURNS USE CATEGORY C MINIMUMS A MINIMUM OF 300 FT OBSTACLE CLEARANCE

PROVIDED AT CATEGORY C CIRCLING MINIMUMS (MDA) WITHIN 1.7 NM FROM ANY RUNWAY.

CAUTION : FAR 91.175 requiresimmediate execution of themissed approach procedurewhen an identifiable part ofthe airport is not distinctlyvisible to the pilot during thecircling maneuver, unless theinability to see results from anormal bank of the aircraftduring the approach.

15 SEC

30 SEC

30 ° BANK

15 SECRUNWAY IN SIGHT ANDWITHIN CIRCLING APPROACH AREATURN 45° FROM RUNWAY CENTERLINETIMING – START

AFTER 30 SECONDS – TURN TO DOWNWIND

FLY 90 ° TO RUNWAY START TIMING CROSSING

RUNWAY CENTERLINE AFTER 15 SECONDS TURN

TO DOWNWIND

FLY OVER RUNWAY WHEN ESTABLISHED ON CENTERLINE

- 30° BANK TURN TO DOWNWIND

TURN OVER RUNWAY AT RUNWAY END, 30 °

BANKED TURN TO DOWNWIND

ABEAM POINT

BASIC CIRCLING PATTERN

45 °

KEY POINT

1

2

3

4

1

2 3

4

1

1 1

NOTE: If Missed Approachbecomes necessary after beginningthe circle, ensure aircraft remains incleared airspace by making first turntoward airport or Missed Approachcourse.






Maneuver Procedures

Engine Failure After Liftoff (Takeoff Continued)

ROLLING TAKEOFF POWER – SET T/O POWER BY 65 KIAS

1 TAKEOFF FLIGHT DIRECTOR – SET BRAKES – HOLD TAKEOFF POWER – SET BRAKES – RELEASE AT V 1/V R

V R CALL – PERFORMED RIGHT HAND – MOVE TO CONTROL WHEEL

ENGINE FAILURE RECOGNIZED POWER – MAXIMUM ROTATE TO 7° AIRSPEED – TAKEOFF SPEED OR ABOVE

AT POSITIVE RATE OF CLIMBCONFIGURATION – GEAR UPPROPELLER (INOP ENGINE) CONFIRM AUTO FEATHER OR FEATHERAIRSPEED – V2

4

AT 400 FT AGL (MINIMUM) OR CLEAR OF OBSTACLESAIRSPEED – ACCELERATE TO V YSE CONFIGURATION – FLAPS UP (IF SELECTED) POWER – SET MAXIMUM ALLOWABLE FAILED ENGINE – IDENTIFY

2

6

53 AIRSPEED – MAINTAIN V2 (MINIMUM)UNTIL 400 FT AGL (MINIMUM) OR CLEAROF OBSTACLES






Maneuver Procedures

Go Around/Missed Approach

MISSED APPROACH FLIGHT DIRECTOR – GO AROUND PITCH – 7° POWER – MAXIMUM ALLOWABLE AIRSPEED – V YSE

AT 400 FT AGL MINIMUM PITCH – 7 ° POWER – PROPELLERS 1900 RPM / ITT

725 ° / (770*) OR TORQUE 2,230 FT-LBS AIRSPEED – 160 KIAS

POSITIVE RATE OF CLIMB GEAR – UP PITCH – 7 ° AIRSPEED – V YSE CONFIGURATION – FLAPS UP

1

2

4

3

ADVISE ATC

*B200




Flight Planning 4



ContentsFlight PlanningGeneral InformationTrip Planning Data

Departure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7Enroute to Primary Destination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7Arrival at Primary Destination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7Enroute to Alternate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7Arrival at Alternate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7

DefinitionsTime, Distance and Fuel

Time, Distance, and Fuel – Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-13Time, Distance, and Fuel – Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-16Cruise Distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-18Cruise Power Setting and Fuel Requirement . . . . . . . . . . . . . . . . . . . .4-19Trip Summary to Destination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-21Fuel Load on Arrival at Destination . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-21Alternate Leg Planning – Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-21Alternate Leg Planning – Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-24Alternate Leg Planning – Cruise Distance, Time, Airspeed, and Fuel Requirement . . . . . . . . . . . . . . . . . . . . . . . . .4-26Trip Summary from Destination to Alternate . . . . . . . . . . . . . . . . . . . .4-28Maximum Holding Time at Primary Destination . . . . . . . . . . . . . . . . . .4-28

Weight and BalanceBasic Empty Weight and Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-31Beechcraft Super King Air B200/B200C . . . . . . . . . . . . . . . . . . . . . . . .4-31Payload Weight and Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-32Zero Fuel Weight and Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-38Fuel Weight and Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-38Ramp Weight and Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-40Takeoff Weight and Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-42Takeoff CG Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-43Landing Weight and Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-45Landing CG Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .4-48



PerformanceTOLD Card . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-51Airport Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-52Maximum Takeoff Weight for Enroute Climb . . . . . . . . . . . . . . . . . . 4-52Minimum Takeoff Power at 2,000 RPM . . . . . . . . . . . . . . . . . . . . . . . 4-54Takeoff Distance and VR . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-56Accelerate-Stop Distance and V1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-58Accelerate-Go Distance and V2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-60Single Engine Climb Gradient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-62Landing Distance Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-64

Supplemental InformationClimb Gradients . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-67ISA Conversion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-67Minimum Takeoff Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-67Primary Static System Blocked . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-67Range vs . Fuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-67Short Field Landing Distances . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-67Single Engine Service Ceiling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-68Takeoff Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4-68Temperature Conversions – Indicated OAT vs . True OAT . . . . . . . . 4-68

Flight Planning



Flight PlanningFlight planning is critical to safety of flight.This chapter provides instruction in, and examples of, flight planning procedures. Charts needed for procedures are provided opposite the applicable instructions.Italics represent example data.




Flight Planning



General InformationA preflight briefing may be obtained from a Flight Service Station by telephone, radio, or personal visit. The briefing should consist of weather, airport, enroute NAVAID information, and NOTAMS.Normally, plan the trip and compute weight and balance first. However, when conditions at the departure airport are near the maximum operating limits of the aircraft, determine takeoff performance data first. This prevents planning a trip and then discovering that takeoff is impossible with the planned passenger and fuel load.The performance tables require that the planned altitude and approximate aircraft weight be known. Aircraft weight decreases as fuel is consumed; fuel consumption can be estimated by scheduling 800 lbs (362.8 kg) for the first hour, 700 lbs (317.5 kg) for the second, 600 lbs (272.1 kg) for each subsequent hour.Modify the estimated time enroute for known delays (e.g., weather, diversions, ATC flow).If fuel conservation is more important than time to destination, consult the specific range vs. cruise wind tables in the King Air 200 Operating Manual for long range cruise information.For groundspeed considerations, this chapter on flight planning uses normal cruise power for the legs to the primary and alternate destinations. Assume a standard powerplant and landing gear configuration.




Flight Planning



Trip Planning Data

DepartureBasic Operating Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8,400 LBSPressure Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4,800 FTActive Runway . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .31Runway Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8,500 FTWeather . . . . . . . . . . . . . . . . . METAR KXXX 201955Z 30015 KT 10SM

SCT050 15/13 A 2992

Wind in Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .20 KT HEADWINDPassengers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5 AT 850 LBS TOTALBaggage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .135 LBSCabinet Supplies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .30LBS Ice Vanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .RETRACTEDFlaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0°

Enroute to Primary DestinationDistance to Destination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 538 NMCruise Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26,000 FTTemperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -25°CWind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .40 KT HEADWIND

Arrival at Primary DestinationWeather . . . . . . . . . . . . . . . . . . . . . . METAR KYYY 201955Z 01005 KT

1/2SM FG ROI/2000FT VV003 17/17 A2992

Pressure Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3,000 FTWind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .20 KT HEADWIND

Enroute to AlternateCruise Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16,000 FTWind . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . NONE

Arrival at AlternateDistance from Primary Destination . . . . . . . . . . . . . . . . . . . . . . 149 NMPressure Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 510 FTActive Runway . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .15



Runway Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6,200 FTWeather . . . . . . . . . . . . . . . . . .TAF KZZZ 201730Z 201818 13010 KTS

6SM SCT040

Flight Planning



DefinitionsTo understand flight planning, it is necessary to be thoroughly familiar with the terms involved. This section reviews the definitions for terms used throughout this chapter.Airspeed position error correction – Correction added to indicated airspeed to obtain calibrated airspeed.Balked Landing – A maximum effort rejected landing maneuver whereby a normal category turbine powered multi-engine airplane can maintain a steady gradient of climb of at least 2.5%with; not more than the power that is available on each engine eight

seconds after initiation of movement of the power controls from minimum flight-idle position.

the landing gear extended. the wing flaps in the landing position; and, a climb speed equal to VREF.

Cruise climb – A speed giving the best combination of altitude gain, fuel consumption, and distance covered.GS – Groundspeed. The speed of the aircraft relative to the ground.KCAS – Knots Calibrated Airspeed. Indicated airspeed corrected for position error due to location of the static ports. CAS equals Indicated Airspeed (IAS) plus airspeed position error correction.KIAS – Knots Indicated Airspeed. The speed of an aircraft, expressed in knots, as shown on it’s static airspeed indicator calibrated to reflect standard atmosphere adiabatic compressible flow at sea level uncorrected for airspeed system errors.M – Mach number. The ratio of true airspeed to local speed of sound.KTAS – Knots True Airspeed. Calibrated airspeed corrected for compressibility effects and air density different from sea level standard.

Use computer/calculator/chart.VA – Maneuvering Speed. Design maneuvering speed. Maximum speed at which abrupt flight control inputs should be attempted.VB – Turbulent Air Penetration Speed. Design speed for maximum vertical gust intensity in level flight as determined in the Flight Envelope in FAR 23.335 (d).VF – Design Flap Speed. The highest speed permissible at which wing flaps may be actuated.VFE – Maximum Flap Extended Speed. The highest speed permissible with wing flaps in a prescribed extended position.VLE – Maximum landing gear extended speed. The maximum speed at which an aircraft can be safely flown with the landing gear extended.VLO – The maximum speed at which the landing gear can be operated.VLOF – Liftoff Speed. The speed at which the aircraft lifts off the ground. VLOF is equal to VR plus three knots.



VMCA – Minimum Control Speed, Air. The minimum flight speed at which the aircraft is directionally controllable as determined in accordance with FAA regulations. The aircraft certification conditions include maximum gross weight, standard day at sea level, one engine becoming inoperative and windmilling, a 5° bank toward the operating engine, landing gear up, flaps in takeoff position, and most rearward CG. For some conditions of weight and altitude, stall can be encountered at speeds above VMCA as established by the certification procedure described above, in which event stall speed must be regarded as the limit of effective directional control.VMCG – Minimum Control Speed, Ground. The minimum speed at which directional control can be maintained if one engine becomes inoperative on takeoff roll with takeoff power on the operative engine.VMO /MMO – Maximum Operating Limit Speed. The speed limit that may not be deliberately exceeded in normal flight operations. VMO is expressed in knots and MMO in Mach number.VNE – Never Exceed Speed. The speed the aircraft is never to exceed.VR – Rotation speed. The speed at which the aircraft is rotated to the takeoff attitude. VR is equal to V1 but not less than 1.05 times VMCA. It must also be high enough to allow VX to be attained before the aircraft reaches a height of 50 ft.VREF – Approach Speed. Target approach speed for a given aircraft weight and flap configuration. VREF equals 1.3 times VS.VS – Stalling Speed. Minimum steady flight speed, the lowest speed at which the aircraft is controllable.VSO – Stalling Speed. Minimum steady flight speed at which the aircraft is controllable in the landing configuration.VSSE – Safe One Engine Inoperative Speed. A speed above both VMCA and stall speed, selected to provide a margin of lateral and directional control when one engine suddenly becomes inoperative. Intentional failing of one engine below this speed is prohibited.VX – Best Two Engine Angle Of Climb Speed. The airspeed that provides the greatest gain in altitude for the horizontal distance travelled.VXSE – Best Single Engine Angle Of Climb Speed. The airspeed that delivers the greatest gain of altitude in the shortest horizontal distance for single-engine operation.VY – Best Two Engine Rate Of Climb Speed. The airspeed that provides the greatest gain in altitude for the elapsed time.VYSE – Best Single Engine Rate Of Climb Speed. The airspeed that delivers the greatest gain in altitude in the shortest time for single-engine operation.V1 – Takeoff Decision Speed. In the event of an engine failure, the speed above which the takeoff is continued; below this speed, takeoffis aborted. V1 is equal to VR.

Flight Planning



V2 – Takeoff Safety Speed. FAR 23.51 (c)(4) defines V2 in terms of calibrated airspeed which allows the gradient of climb required in FAR 23.67 (c)(1-2). It must not be less than 1.10 VMC or 1.20 VS1. V2 is not a required speed for certification of Normal Category airplanes but is provided as additional information for an enhanced level of performance safety beyond the requirement of Part 23.67 (b).FAR 23 .67 (b) – Climb; One Engine Inoperative. Requires the steady gradient of climb, for normal category aircraft, at an altitude of 400 ft above the takeoff surface to be measurably positive with;

the critical engine inoperative and its propeller in the minimum drag position.

the remaining engine at takeoff power. the landing gear retracted. the wing flaps in the takeoff position. climb speed equal to that achieved at 50 ft in the demonstration

required by FAR 23.53.The steady gradient of climb at an altitude of 1,500 ft above the takeoff surface must be not less than .75 % with;

the critical engine inoperative and its propeller in the minimum drag position.

the remaining engine at not more than maximum continuous power.

the landing gear retracted. the wing flaps retracted; and, a climb speed not less than 1.20 VS1.




Flight Planning



Time, Distance and FuelThis section of the example begins with an estimation of trip time and fuel consumption. Using the estimated answers, compute actual time, distances, and fuel.

Time, Distance, and Fuel – ClimbBegin by determining the climb requirements from sea level to the field pressure altitude (4,800 ft). Next, determine the climb requirements from sea level to the final/selected cruise altitude (26,000 ft). Finally, subtract the former from the latter to determine the actual climb requirements. Use the Time, Fuel, and Distance to Climb chart (Figure 4-1).1. Enter the chart from the bottom left at the outside air temperature

at the takeoff airport (15°C). Move up to intersect the field pressure altitude (4,800 ft).

2. Move to the right to intersect the initial climb weight (12,500 lbs (5,669.9 kg)).3. Move down to the bottom of the chart and read the time, fuel, and

distance required in the climb.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 MINUTESDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .31 LBS

4. Enter the chart a second time from the bottom left at the outside air temperature at the cruise altitude (-27°C or ISA +10°C). Move up to intersect the cruise altitude (26,000 ft).


distance required in the climb.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18 MINUTESDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54 NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .223 LBS



Time, Fuel, and Distance to Climb

4,800 FT

-27°C 15°C

4 NM

31 LBS

2 MIN

54 NM

223 LBS

18 MIN

2 1

2

1 Sea Level to Field Pressure Altitude

Sea Level to Cruise Altitude

Figure 4-1: Time, Fuel, and Distance To Climb Chart

Flight Planning



7. Subtract the sea level to field pressure climb figures from the figures for sea level to cruise altitude; these are the actual figures for the climb.

Time to 26,000 ft 18 minutesTime to 4,800 ft -2ACTUAL TIME REQUIRED 16 minutes

Distance to 26,000 ft 54 NMDistance to 4,800 ft -4ACTUAL DISTANCE REQUIRED 50 NM

Fuel to 26,000 ft 223 lbsFuel to 4,800 ft -31ACTUAL FUEL REQUIRED 192 lbs

8. Determine the average groundspeed by dividing the distance travelled by the time required.Because the time required is in minutes, multiply the result by 60 to determine the speed in knots.

50 NM16 (minutes)

9. Correct for the existent wind by adding or subtracting the headwind component to determine groundspeedFor this example, there is a headwind of 20 Kts.

187.5 -20 = 167.5 Kts ground speed10. Multiply the groundspeed by the time to climb to determine the actual

distance travelled during the climb.Use calculator/navigation computer.

16 minutes at 167.5 Kts = 45 NM

x 60 (minutes) = 187.5 Kts

DistanceTime

= x 60 = Speed

Speed ± Headwind Component = Actual Groundspeed



Time, Distance, and Fuel – DescentDetermine the descent requirements from the cruise altitude (26,000 ft) to sea level, then the field pressure altitude (3,000 ft) to sea level. Finally, subtract the former from the latter to determine the actual descent requirements. Use the Time, Fuel, and Distance to Descend chart (Figure 4-2).

For this example, anticipate holding at 5,000 ft at the destination because the weather is marginal. Assume, therefore, that you are descending only to that altitude (i.e., 5,000 ft) instead of the destination pressure altitude (i.e., 3,000 ft).

1. Enter the chart from the left at the cruise altitude (26,000 ft). Move to the right to the diagonal reference line.

2. Move down to the bottom of the chart and read the time, fuel, and distance requirements.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .17.4 MINUTESDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82 NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .165 LBS

3. Enter the chart from the left at the destination pressure altitude (holding altitude, 5,000 ft). Move to the right to the diagonal reference line.

4. Move down to the bottom of the chart and read the time, fuel, and distance requirements.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .3.3 MINUTESDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.5 NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .40 LBS

Flight Planning



Time, Fuel, and Distance to Descend

1 Cruise Altitudeto Sea Level

2 Holding Altitudeto Sea Level

2 1

17.4 MIN

165 LBS

82 NM

3.3 MIN

40 LBS

14.5 NM

26,000 FT

Figure 4-2: Time, Fuel, and Distance To Descend Chart



5. Subtract the destination pressure altitude (i.e., holding altitude) to sea level descent figures from the figures for cruise altitude to sea level descent figures; these are the actual figures for the descent.

Time from 26,000 ft 17.4 minutesTime from 5,000 ft -3.3ACTUAL TIME REQUIRED 14.1 minutesRound the time down to an even 14 min.

Distance from 26,000 ft 82.0 NMDistance to 5,000 ft -14.5ACTUAL DISTANCE REQUIRED 67.5 NM

Fuel from 26,000 ft 165 lbsFuel to 4,800 ft -40ACTUAL FUEL REQUIRED 125 lbs

6. Determine the average groundspeed by dividing the distance travelled by the time required.

67.5 (NM)14 (minutes)

For future calculations, round this figure to 289 Kts.7. Correct the calculated speed for the existent wind by adding or

subtracting the wind component.For this example, there is a headwind of 20 Kts.

289 - 20 = 269 Kts8. Multiply the GS by the time to determine the actual distance travelled

during the descent.14 minutes at 269 Kts = 63 NM

Cruise DistanceThe total cruise distance is the total distance from takeoff to the primary destination (538 NM) minus the distances for climb and descent (45 and 63 NM, respectively).

538 - (45 + 63) = 430 NMRemember, however, the aircraft has not landed at 538 NM; it remains at a holding altitude of 5,000 ft.

= x 60 (minutes) = 289.2 Kts

DistanceTime

= x 60 = Speed

Speed±Wind

Actual Groundspeed

C = T - (L + D)where:C = Cruise DistanceT = Total DistanceL = Climb DistanceD = Descent Distance

Flight Planning



Cruise Power Setting and Fuel RequirementDetermine the cruise power setting and fuel requirement from the appropriate Normal Cruise power chart (Figure 4-3).

For this example, the appropriate Normal Cruise Power chart is that for 1,700 RPM and ISA + 10°C.

1. Compute groundspeed from the airspeed shown on the chart (274 KIAS) corrected for the existent wind (a 40 kts headwind) by subtracting the headwind component.

274 - 40 = 234 KTAS2. To determine the time at cruise, divide the distance by the speed.

430 (NM)234 KTAS

= 1.84 hr (1 hr; 50 minutes)

Tabulated Airspeed±Wind

Groundspeed

DistanceSpeed

= Time



Normal Cruise Power – 1,700 RPMISA + 10°C

NOTE: IOAT, torque, and fuelow based on 11,000 pounds.

Figure 4-3: Normal Cruise Power Chart

Flight Planning



Trip Summary to DestinationTo this point, only the time, distance, and fuel figures to the primary destination are known; find the sum of the figures for each phase of flight to determine the total figures.

Time for Taxi 0 minutesClimb to 26,000 ft 16 Cruise (60 + 50) 110 Descent to 5,000 ft +14 TOTAL TIME REQUIRED 140 minutesTotal time is 140 minutes, or to 2 hours and 20 minutes.

Taxi 0 NMClimb to 26,000 ft 45 Cruise 430 Descent to 5,000 ft +63 TOTAL DISTANCE REQUIRED 538 NM

Taxi/Takeoff (Standard) 90 lbsClimb to 26,000 ft 192 Cruise 1,034 Descent to 5,000 ft +125 TOTAL FUEL REQUIRED 1,441 lbs

Remember, the example does not actually reflect a landing at the destination; rather, plan to hold at 5,000 ft.

Fuel Load on Arrival at DestinationNext, determine the amount of fuel remaining upon arrival at the destination by subtracting the fuel consumed from the total fuel load.

Assume the aircraft loads fuel to the maximum ramp weight, or 12,590 lbs (5,710.7 kg). The basic operating weight plus the crew, passengers, baggage, and miscellaneous supplies is only 9,515 lbs (4,315.9 kg); the remaining weight, 3,075 lbs (1,394.7 kg), is made up in fuel.Subtract 90 lbs (40.8 kg) as the standard fuel requirement for start, taxi, and takeoff operations. This results in a takeoff weight of 12,500 lbs (5,669.9 kg), the maximum allowable.Subtracting the fuel consumed on the leg to the destination (1,441 lbs (653.6 kg)) from the 2,985 lbs (1,353.9 kg) takeoff fuel load yields the fuel remaining upon arrival at the destination, or 1,544 lbs (700.3 kg).

Alternate Leg Planning – ClimbA diversion to the alternate requires additional flight planning. The reserve fuel load on landing at the alternate is 500 lbs (226.7 kg). Calculate the climb from the holding altitude to the alternate cruise leg altitude; use the same technique as for the climb from takeoff to the cruise altitude.



Assume the weather conditions enroute to the planned alternate do not required ice vanes. The winds aloft are light and variable with temperatures remaining at ISA + 10°C at all levels.For this example, climb from the holding pattern altitude, 5,000 ft, to 16,000 ft for cruise to the alternate. Recall that the distance to the alternate is 149 NM.

Begin by determining the climb requirements from sea level to the holding altitude (5,000 ft). Next, determine the climb requirements from sea level to the new cruise altitude (16,000 ft). Finally, subtract the former from the latter to determine the actual climb requirements. Use the Time, Fuel, and Distance to Climb chart (Figure 4-4).1. Enter the chart from the bottom at the ISA diagonal. Move up to the

pressure altitude (5,000 ft); correct for the temperature (ISA +10°C).2. Move to the right to intersect the initial climb weight.

When the fuel requirement to the destination, 1,441 lbs (653.6 kg), is subtracted from the takeoff weight, the result is 11,059 lbs (5,016.2 kg). Round this figure up to an even 11,100 lbs (5,034.8 kg).

3. Move down to the bottom of the chart and read the time, fuel, and distance required in the climb.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 MINUTESDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.2 NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .25 LBS

4. Enter the chart a second time from the bottom at the ISA diagonal. Move up to the pressure altitude (16,000 ft). Correct for the temperature (ISA + 10°C).


distance required in the climb.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 MINUTESDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22 NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112 LBS

7. Subtract the sea level to holding altitude climb figures from the figures for sea level to cruise altitude; these are the actual figures for the alternate climb.

Time to 16,000 ft 8 minutesTime to 5,000 ft -2ACTUAL TIME REQUIRED 6 minutes

Distance to 16,000 ft 22.0 NMDistance to 5,000 ft -4.2ACTUAL DISTANCE REQUIRED 17.8 NM

Fuel to 16,000 ft 112 lbsFuel to 5,000 ft -25ACTUAL FUEL REQUIRED 87 lbs

Because no wind correction is necessary, these are the figures for the climb to the cruise altitude.

Flight Planning



Time, Fuel, and Distance to Climb

1 Correction For ISA + 10°C

2 MIN 8 MIN

25 LBS 112 LBS

4.2 NM 22 NM

1 1

5,000 FT

Figure 4-4: Time, Fuel, and Distance To Climb Chart



Alternate Leg Planning – DescentDetermine the descent requirements from the cruise altitude (16,000 ft) to sea level, then the field pressure altitude (510 ft) to sea level. Finally, subtract the former from the latter to determine the actual descent requirements. Use the Time, Fuel, and Distance to Descend chart (Figure 4-5).1. Enter the chart from the left at the cruise altitude (16,000 ft). Move to

the right to the diagonal reference line.2. Move down to the bottom of the chart and read the time, fuel, and

distance requirements.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .10.5 MINUTESDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48 NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115.5 LBS

3. Enter the chart from the left at the destination pressure altitude (510 ft). Move to the right to the diagonal reference line.

4. Move down to the bottom of the chart and read the time, fuel, and distance requirements.Time . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .0.5 MINUTEDistance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 NMFuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5 LBS

5. Subtract the figures for the descent from the alternate destination pressure altitude to sea level from the figures for the descent from the cruise altitude to sea level; these are the actual figures for the descent.

Time from 16,000 ft 10.5 minutesTime from 510 -0.5ACTUAL TIME REQUIRED 10.0 minutes

Distance from 16,000 ft 48 NMDistance from 510 ft -2ACTUAL DISTANCE REQUIRED 46 NM

Fuel from 16,000 ft 115.5 lbsFuel to 510 ft -5.0ACTUAL FUEL REQUIRED 110.5 lbs

Because no wind correction is necessary, these are the figures for the descent from the cruise altitude.

Flight Planning



Time, Fuel, and Distance to Descend

16,000 FT

51 FT

0.5 MIN 10.5 MIN

48 NM2 NM

115.55 LBS

Figure 4-5: Time, Fuel, and Distance To Descend Chart



Alternate Leg Planning – Cruise Distance, Time, Airspeed, and Fuel RequirementNext, determine the cruise distance between the primary destination and the alternate.1. The cruise distance is the distance from the primary destination to

the alternate (149 NM) minus the distances for climb and descent (21.8 and 46 NM, respectively).

149 -17.8 +46 = 85.2 NM2. Determine the airspeed from the Normal Cruise Power table

(Figure 4-6). Enter the table from the left at the cruise altitude (16,000 ft). Move to the right to the cruise weight (11,000 lbs (4,989.5 kg)). Read the applicable TAS (281).

3. To determine the time at cruise, divide the distance by the speed. The result is terms of hours; multiply it by 60 to convert to minutes.

85.2 (NM)281 (KTAS)

0.30 (hours) x 60 (minutes) = 18.0 minutes4. Multiply the fuel flow shown on the Normal Cruise Power table by the

hour figure determined in step 3.758 (lbs/hr) x 0.30 (hr) = 227.4 (lbs (1,031.4 kg))

Round this figure to an even 230 lbs (104.3 kg).

= 0.30 hours

C = T - (L + D)where:C = Cruise DistanceT = Total DistanceL = Climb DistanceD = Descent Distance

DistanceGroundspeed

= Time

Flight Planning



Normal Cruise Power – 1,700 RPMISA + 10°C

NOTE: IOAT, torque, and fuelow based on 11,000 pounds.

Figure 4-6: Normal Cruise Power Table



Trip Summary from Destination to AlternateFind the sum of the figures for each phase of the flight to the alternate to determine the total figures involved. Be sure to add 500 lbs (226.7 kg) to the fuel required as the planned reserve.

Climb from 5,000 ft to 16,000 ft 6 minutesCruise at 16,000 ft 17Descent to 510 ft + 10TOTAL TIME REQUIRED 35 minutes

Climb from 5,000 ft to 16,000 ft 17.8 NMCruise at 16,000 ft 85.2Descent to 510 ft + 46.0TOTAL DISTANCE REQUIRED 149.0 NM

Climb from 5,000 ft to 16,000 ft 87.0 lbsCruise at 16,000 ft 230.0Descent to 510 ft 110.5Planned Reserve +500.0TOTAL FUEL REQUIRED 927.5 lbsRound off the total fuel required to an even 928 lbs.

Maximum Holding Time at Primary DestinationPerform the following to determine how long the aircraft can hold at the primary destination before diverting to the alternate.1. Subtract the fuel required for the trip to the alternate (including the 500

lb reserve) from the fuel remaining on arrival at the primary destination.

Fuel Supply on Arrival at Destination 1,544 lbsFuel Required to Alternate -928TOTAL FUEL AVAILABLE FOR HOLDING 616 lbs

2. Enter the Holding Time chart (Figure 4-7) from the bottom with the available fuel (616 lbs (279.4 kg)).

3. Move up to intersect the diagonal for the holding altitude (5,000 ft).4. Move left to the edge of the chart and read the maximum holding

time available (1.20 hours, or 1 hour, 12 minutes).

Flight Planning



Holding TimePower Setting 800 ft-lbs at 1,700 rpm Applicable for All Temperatures

616 LBS

1.20

Figure 4-7: Holding Time Chart




Flight Planning



Weight and BalancePrecise weight and balance computations are essential elements of flight planning. Accuracy in these computations helps to ensure a safe flight. This section reviews the procedures for computing weight and balance data using the manufacturer’s Airplane Flight Manual.

Basic Empty Weight and MomentBasic empty weight is the weight of the aircraft including all fixed operating equipment, unusable fuel, and engine oil. This weight and its moment are noted on the aircraft weighing form. Alternations affect the aircraft weight and moment, refer to the weight and balance record for the corrected information.

Assume that the basic empty weight is 8,160 lbs (3,701.3 kg) with a moment of 15,220. Record this information on the Weight and Balance Loading form (hereafter referred to as the loading schedule) (Figure 4-8).

Beechcraft Super King Air B200/B200CWeight and Balance Loading FormSERIAL REGISTRATION NO . DATE

PASSENGERS OR CARGO REF ITEM WEIGHT MOM/100ITEM WEIGHT MOM/100 1. Basic Empty Cond. 8,160 15,220

Location (Row, F .S ., Etc .) 2. Pilot

3. Passengers or Cargo4. Baggage5. Cabinet Contents6. Sub Total

Zero Fuel Condition Do Not Exceed 11,000 lbs

7. Fuel Loading8. Sub Total

Ramp Condition9. Less Fuel For Start,

Taxi, and Take Off10. Sub Total

Take Off Condition11. Less Fuel to Destination

TOTAL PASSENGERS OR CARGO 12. Landing Condition

* Fuel for start, taxi and take-off is normally 90 lbs (40.8 kg) at an average moment/100 or 177.Figure 4-8: Weight and Balance Loading Form



Payload Weight and MomentPayload is defined as the passengers, cabinet contents, and baggage. The payload moment is the sum of the individual moments of each of these components. Individual moments are calculated by multiplying a flight station’s weight by its arm, or the distance from the reference datum line. Divide moment by 100 to facilitate further computations.1. Locate the appropriate table (Occupants, Baggage, and Cabinet

Contents).In this example, use the seating arrangement (Figure 4-9) and loading table (Figure 4-10) provided.

2. Find the appropriate weight in the left column, then move to the right to the appropriate flight station column to obtain the moment. Record the derived information in the appropriate column of the loading schedule (Figure 4-11).Assume the following passenger assignment.a. Passenger A sits at flight station 176 and weighs 170 lbs (77.11 kg).

His moment is 299.b. Passengers B and C sit at flight station 215; each weighs

170 lbs (77.1 kg). Their moments, individually, are 366.c. Passengers D and E sit at flight station 259; each weighs

170 lbs (77.1 kg). Their moments, individually, are 440.3. Add the passenger weights and moments, and record the results on

the loading schedule both at the bottom of the left column and at row 3 in the right column.

Weight (Lbs) MomentPassenger A 170 299Passenger B 170 366Passenger C 170 366Passenger D 170 440Passenger E +170 +440TOTALS 850 +1,911

NTTE: Use actual or estimated passenger weights versus “FAA Standard” 170 lbs (77.11 kg).

Moment = Weight x Arm

Flight Planning



Cabin Arrangement

9

Figure 4-9: Seating Arrangement

Useful Load Weights and Moments

Figure 4-10: Useful Load and Moments Table



Beechcraft Super King Air B200/B200CWeight and Balance Loading Form

4. Record the total weight and moment for the crew on the loading schedule (Figure 4-13).Both pilots sit at flight station 129. Because each weighs 170 lbs (77.1 kg) with a moment of 219 (Figure 4-12), their combined weight is 340 lbs (154.2 kg) at a moment of 438.

5. Record the weight and moment of the baggage (Figure 4-14).The baggage, which is stowed in the aft cabin compartment (flight station 325), weighs 135 lbs (61.2 kg). Separate the baggage weight into two quantities: 100 lbs (45.3 kg) and 35 lbs (15.8 kg). Find the moment for the 100 lbs (45.3 kg) figure, or 325. Next, interpolate the moment for 35 lbs (15.8 kg) of baggage, or 114. Add the moments together for the total baggage moment, or 439.

6. Record the weight and moment of the cabinet contents (Figure 4-13).The cabinet contains 30 lbs (13.6 kg) of goods, which have a combined moment of 59.

SERIAL REGISTRATION NO . DATE


Location (Row, F .S ., Etc .) 2. Pilot 850 1,911

A, F.S. 176 170 299 3. Passengers or CargoB, F.S. 215 170 366 4. BaggageC, F.S. 215 170 366 5. Cabinet ContentsD, F.S. 259 170 440 6. Sub Total


E, F.S. 259 170 440 7. Fuel Loading8. Sub Total




TOTAL PASSENGERS OR CARGO 850 1,911 12. Landing Condition

* Fuel for start, taxi and take-off is normally 90 lbs (40.8 kg) at an average moment/100 or 177.Figure 4-11: Weight and Balance Loading Schedule

Maximum Weight in Baggage Compartment

BB-1052; BB-1091 and Subse-quent; BL-58 and Subsequent; Prior Aircraft with Beech Kit 101-5068-1E:With Fold-Up Seats . . 510 LBSWithoutFold-Up Seats . . . . . 550 LBSAircraft Prior to BB-1052;BB-1091; BL-58E:With Fold-Up Seats . . 370 LBSWithoutFold-Up Seats . . . . . 410 LBS

Flight Planning



Useful Load Weights and MomentsOccupants

Figure 4-12: Useful Load Weights and Moments






Location (Row, F .S ., Etc .) 2. Pilot 340 438

A, F.S. 176 170 299 3. Passengers or Cargo 850 1,911B, F.S. 215 170 366 4. Baggage 135 439C, F.S. 215 170 366 5. Cabinet Contents 30 59D, F.S. 259 170 440 6. Sub Total







* Fuel for start, taxi and take-off is normally 90 lbs (40.8 kg) at an average moment/100 or 177.

Figure 4-13: Weight and Balance Loading Form

Flight Planning




11435 LBS

Figure 4-14: Weight and Moment Of The Baggage



Zero Fuel Weight and MomentThe zero fuel weight is the sum of the basic empty weight plus the weights for the crew, payload (passengers or cargo), baggage, and cabinet contents. The zero fuel weight moment is the sum of the moments for the same components. Record the zero fuel weight and moment on the loading schedule (Figure 4-15).

Weight (Lbs) Moment

Basic Empty Weight 8,160 15,220Pilot (2) 340 438Payload 850 1,911Baggage 135 439Cabinet Contents 30 59ZERO FUEL WEIGHT AND MOMENT 9,515 18,067


Figure 4-15: Zero Fuel Weight and Moment On The Loading Schedule

Fuel Weight and MomentUse the Usable Fuel table (Figure 4-16) to determine the momentfor the fuel requirement for the flight.1. Enter the table and determine the moment for the fuel load (3,075 lbs).











* Fuel for start, taxi and take-off is normally 90 lbs (40.8 kg) at an average moment/100 or 177

Flight Planning



The fuel load is 3,075 lbs (1,394.7 kg) (2,985 lbs (1,353.9 kg)) trip fuel plus 90 lbs (40.8 kg) start/ taxi/takeoff fuel). After interpolation, the moment for this weight is 5,760.

2. Record this information on the loading schedule (Figure 4-17).

Useful Load Weights and MomentsUsable Fuel

1 2

1 3,075 LBS

2 5,760 MOM

Figure 4-16: Usable Fuel Table




Figure 4-17: Weight and Balance Loading Form

Ramp Weight and MomentThe ramp weight is the gross weight of the aircraft fully loaded for flight. It is the sum of the zero fuel weight and the weight of the fuel load. The ramp weight moment is the sum of the zero fuel weight moment and the fuel load moment. Record this information on the loading schedule (Figure 4-18).

Because the zero fuel weight is 9,515 lbs (4,315.9 kg) and the fuel load weight is 3,075 lbs (1,394.7 kg), the ramp weight is 12,590 lbs (710.7 kg). Likewise, because the zero fuel weight moment is 18,067 and the fuel load moment is 5,760, the ramp weight moment is 23,827.

Weight LimitsZero Fuel Weight . . . 11,000 LBS Ramp Weight . . . . . . 12,590 LBS






9,515 18,067

E, F.S. 259 170 440 7. Fuel Loading 3,075 5,7608. Sub Total






Flight Planning



Beechcraft Super King Air B200/B200CWeight and Balance Loading FormSERIAL REGISTRATION NO . DATE





9,515 18,067


Ramp Condition12,590 23,827

9. Less Fuel For Start, Taxi, and Take Off

10. Sub TotalTake Off Condition

11. Less Fuel to DestinationTOTAL PASSENGERS OR CARGO 850 1,911 12. Landing Condition

Figure 4-18: Weight and Moment Loading Schedule* Fuel for start, taxi and take-off is normally 90 lbs (40.8 kg) at an average moment/100 or 177.



Takeoff Weight and MomentThe takeoff weight is the ramp weight minus the fuel used during start, taxi, and takeoff. Beech calls for a standard burn of 90 lbs 40.8 kg), which has a moment of 177.1. Subtract the start/taxi/takeoff fuel from the ramp weight; the result is

the takeoff weight.Subtracting 90 from the ramp weight of 12,590 lbs (5,710.7 kg) leaves 12,500 lbs (5,669.9 kg), or the maximum takeoff weight.

2. Subtract the start/taxi takeoff fuel moment from the ramp weight moment; the result is the takeoff weight moment.Subtracting 177 from 23,827 leaves 23,650.

3. Record this information on the loading schedule (Figure 4-19).


Figure 4-19: Takeoff Weight and Moment Loading Schedule






9,515 18,067




90 177


12,500 23,650

11. Less Fuel to DestinationTOTAL PASSENGERS OR CARGO 850 1,911 12. Landing Condition


Takeoff Weight LimitNon-FAR 135Operations . . . . . 12,500 LBS Far 135Operations . . . . .PER MAXIMUMENROUTE WEIGHT CHART, AFM SECTION 5

Flight Planning



Takeoff CG LimitsUse the Moment Limits vs Weight chart (Figure 4-20) to determine whether the calculated takeoff moment is within limits.1. Enter the chart from the left with the calculated takeoff weight

(12,500 lbs (5,669.9 kg)).2. Move to the right to intersect the takeoff moment (23,650).3. If the intersection of these values lies within the CG envelope

(unshaded area), the aircraft is within limits.The intersection of the values is within takeoff CG limits.

Alternatively, calculate the CG arm by dividing the total moment by the total weight. Because moment is divided by 100 to simplify computations, multiply the result by 100 to derive the actual arm.

23,65012,500

1.89 x 100 = 189Read up from this number to the intersection of the takeoff weight line to determine whether or not the aircraft is within CG limits for takeoff.

Projecting up from 189 to the takeoff weight line, 12,500, confirms that the aircraft is within limits for takeoff.

The forward CG limit for any given weight between 11,279 and 12,500 lbs (5,116.0 and 5,669.9 kg) can be derived from the following equation:

Wt305.25

Where:Wt = Takeoff WeightCG = Center of Gravity

= 1.89

+ 141.05Fwd CG =

Center of Gravity Limits

Aft limit is 196.4 inches aft of datum at all weights.

Forward limit is 185.0 inches aft of datum at 12,500 lbs, with straight line variation to 181.0 inches aft of datum at 11,279 lbs. Forward limit is 181.0 inches aft of datum at 11,279 lbs or less.

Reference datum is 83.5 inches forward of the center of the front jack point.



Moment Limits vs Weight

23,650

189

Figure 4-20: Limits Vs Weight Chart

Flight Planning



Landing Weight and MomentThe landing weight is the takeoff weight minus the weight of the fuel used to reach the destination. The landing weight moment is takeoff weight minus the difference between takeoff fuel moment and landing fuel moment. Record the appropriate numbers on the loading schedule (Figure 4-21).

For this example, assume that upon arrival at the primary destination, conditions require an immediate diversion to the alternate. The example that follows assumes arrival at the alternate.

1. Subtract the fuel consumed from the fuel load at takeoff. This is the landing fuel load. Interpolate its moment.First determine the total fuel consumed; record this figure on the loading schedule.

Fuel to Primary Destination 1 1,441 lbsFuel to the Alternate + 928TOTAL FUEL CONSUMED 2,369 lbs

1 Fuel burned from takeoff rotation to the primary destination; does not include the standard 90 lbs (40.8 kg) of fuel used for start, taxi, and takeoff.

Next, subtract the fuel burn from the takeoff fuel load.

Takeoff Fuel Load 1 2,985 lbsTotal Fuel Consumed -2,369LANDING FUEL LOAD 616 lbs

1 Fuel on board at takeoff rotation; does not include the standard 90 lbs (40.8 kg) of fuel used for start, taxi, and takeoff.

Using the Useful Load Weights and Moments, Usable Fuel table from the AFM (Figure 4-22) to interpolate the moment for 616 lbs (279.4 kg)results in a moment of 1,294.

2. Compute the difference between the takeoff and landing fuel moments.

Takeoff Fuel Moment 5,583 lbsLanding Fuel Moment -1,095MOMENT CHANGE 4,488 lbs

Record this figure on the loading schedule as the moment of the fuel consumed.

3. Determine the landing weight and moment by subtracting the consumed fuel weight from the takeoff weight and by subtract-ing the difference between the takeoff and landing fuel moments from the takeoff weight moment.

Takeoff Weight 12,500 lbsFuel Consumed -2,369LANDING WEIGHT 10,131 lbs

Wt- Wf= Wlwhere:Wt = Takeoff WeightWf = Weight of Fuel ConsumedWl = Landing Weight

Mt- (Mft- Mfl) = Mlwhere:Mt = Takeoff Weight MomentMft = Takeoff Fuel MomentMfl = Landing Fuel MomentMl = Landing Weight Moment

Landing Weight LimitWeight . . . . . . . . . . . .12,500 LBS



Takeoff Weight MOMENT 23,650 Difference Between Takeoff and Landing Fuel Moments

-4,488

LANDING WEIGHT MOMENT 19,162

Record these figures on the loading schedule.


Figure 4-21: Landing Weight Weight and Moment Loading Schedule






9,515 18,067




90 177


12,500 23,650

11. Less Fuel to Destination 2,369 4,488TOTAL PASSENGERS OR CARGO 850 1,911 12. Landing Condition 10,131 19,162


Flight Planning




2

1 616LBS

2 1095MOM

1

Figure 4-22: Usable Fuel Table



Landing CG LimitsUse the Moment Limits vs Weight chart (Figure 4-23) to determine whether the calculated landing moment is within limits.1. Enter the chart from the left with the calculated landing weight

(10,131 lbs(4,595.3 kg)).2. Move to the right to intersect the landing moment (19,162).3. If the intersection of these values lies within the CG envelope

(unshaded area), the aircraft is within limits.The intersection of the values is within landing CG limits.

Use the alternate method to calculate the CG arm. That is, divide the landing moment by the landing weight. Because moment is divided by 100 to simplify computations, multiply the result by 100 to derive the actual arm.

19,16210,131

1.89 x 100 = 189Read up from this number to the intersection of the landing weight line to determine whether or not the aircraft is within CG limits for landing.

Projecting up from 189 to the landing weight line, 10,131 lbs (4,595.3 kg), confirms that the aircraft is within limits for landing.

See forward CG limit computation on page 4-50.

= 1.8914

Center of Gravity Limits

Aft limit is 196.4 inches aft of datum at all weights.

Forward limit is 185.0 inches aft of datum at 12,500 lbs, with straight line variation to 181.0 inches aft of datum at 11,279 lbs. Forward limit is 181.0 inches aft of datum at 11,279 lbs or less.

Reference datum is 83.5 inches forward of the center of the front jack point.

Flight Planning



Moment Limits vs Weight

19,162

189

10,131 LBS

Figure 4-23: Limits vs Weight Chart




Flight Planning



Performance

TOLD CardUse a TakeOff and Landing Data (TOLD) card to record takeoff and landing data. It serves as a convenient reference aid in the cockpit.The Takeoff side of the card provides spaces for the following information: ATIS V1/VR – Takeoff Decision Speed/Rotation Speed V2 – Takeoff Safety Speed VYSE – Best Single Engine Rate of Climb Speed WEIGHT – Takeoff Weight FLAPS – Takeoff Flap Setting T/O MIN. TORQ. – Minimum Takeoff Torque Setting RWY RQD – Computed Takeoff Field Length CLEARANCE.

The Approach side of the card provides spaces for the following information: ATIS VREF – Landing Reference Speed VAP – Target Approach Speed VZF – Zero Flap Approach Speed WEIGHT – Landing Weight BALKED LANDING – 100 (always 100 Kts at all weights) VYSE – Best Single Engine Climb Speed RWY RQD – Computed Landing Field Length NOTES.

WEIGHT

BALKED LANDING

V REF

V AP

RWY

RQD

KING AIR

FT.

V ZF

100

V YSE

NOTES:

WEIGHT

FLAPS

V 1

V 2

V YSE

CLEARANCE

T/O MIN. TORQ.

RWY

RQD

KING AIR

V R/

ATIS

GO

AROUND

ATIS

APPROACH

TAKEOFF



Airport InformationObtain airport information from the standard sources.

In this case, use the trip planning data provided and assume a forecast runway wind of 15 Kts headwind.

Maximum Takeoff Weight for Enroute ClimbUse the Maximum Take-Off Weight Permitted by Enroute Climb Requirement chart (Figure 4-24) to determine the maximum takeoff weight permitted by climb requirements.

Notice that the King Air 200 is not limited by enroute climb.

Flight Planning



Maximum Takeoff WeightPermitted by Enroute Climb Requirement

NOTE: For operation with icevanes extended. No o loading isrequired.

Figure 4-24: Enroute Climb Requirement Chart



Minimum Takeoff Power at 2,000 RPMDetermine the minimum power at which the AFM performancecharts are accurate.

A power (torque) lower than that derived by this process inval idates information derived from the charts in AFM Section V.

1. Enter the appropriate Minimum Take-Off Power at 2,000 RPM chart (ice vanes retracted) (Figure 4-25) from the bottom at the ambient temperature (15°C). Move up to intersect the pressure altitude (4,800 ft).

2. Move left to the edge of the chart to read the minimum power setting.For this example, the minimum power setting is 2,140 ft-lbs. If the engines do not produce 2,140 ft-lbs of torque, the charts in AFM Section V indicate better performance than can be expected.

3. Record the power setting on the TOLD card.

WEIGHT

FLAPS

V 1

V 2

V YSE

CLEARANCE

T/O MIN. TORQ.

RWY

RQD

KING AIR

V R/

ATIS

TAKEOFF

Flight Planning



Minimum Takeoff Power at 2,000 RPMWith Ice Vanes Retracted – 65 Kts

NOTE: Torque increasesapproximately 20 ft-lbs from zero to65 kts. The power (torque) indicatedis the minimum value for which take-o performance in this section can beobtained. Excess power which can bedeveloped without exceeding enginelimitations may be utilized.

2,140FT LBS

15°C

4,800 FT

Figure 4-25: Minimum Take-Off Power At 2,000 RPM Chart



Takeoff Distance and VRUse the appropriate Take-Off Distance chart (flaps 0°) (Figure 4-26)to determine the takeoff distance required.1. Enter the chart from the bottom left at the ambient temperature

(15°C). Move up to the pressure altitude (4,800 ft).2. Move to the right to the weight reference line. Parallel the guide lines

down and right to the actual takeoff weight, then proceed to the wind reference line.Because the weight reference line is also the intersection for the takeoff weight, 12,500 lbs (5,669.9 kg), move straight to the right to the wind reference line.

3. Parallel the wind guidelines to the existent wind component (down and right for 15 kt headwind), then move straight to the right to the obstacle height reference line.

4. Parallel the guidelines upward to the edge of the chart. The value indicated is the distance required to clear a 50 ft obstacle.In this case, 3,900 ft are required; however, there is no obstacle indicated for this flight. Move straight right from the reference line to read 2300 ft.

5. Consult the table at the top of the chart and determine VR for the takeoff weight. Record this value on the TOLD card.For all takeoff weights, VR is 95 Kts.

WEIGHT

FLAPS

V 1

V 2

V YSE

CLEARANCE

T/O MIN. TORQ.

RWY

RQD

KING AIR

V R/

ATIS

TAKEOFF

Flight Planning



Takeoff Distance – Flaps 0%

15°C

2,30

0 FT

3,90

0 FT

Figure 4-26: Take-Off Distance Chart



Accelerate-Stop Distance and V1Determine the distance required to accelerate the aircraft to just shortof V1 and then bring it to a full stop; use the appropriate Accelerate-Stop Distance chart (flaps 0°) (Figure 4-27).1. Enter the chart from the bottom left at the ambient temperature

(15°C). Move up to the pressure altitude (4,800 ft).2. Move to the right to the takeoff weight reference line, then follow the

guidelines down and right to the actual takeoff weight. From that point, move to the right to the wind reference line.Because the weight reference line is also the intersection for the takeoff weight, 12,500 lbs (5,669.9 kg), move straight to the right to the wind reference line.

3. Parallel the guidelines (down and right) to the appropriate wind component (15 Kts), then move to the right to the edge of the chart and read the accelerate-stop distance required.For this example, a distance of 3,800 ft is required. Because the runway available is 8,500 ft in length, this is not limiting.

4. Consult the table above the chart and determine V1 for the takeoff weight. Record this value on the TOLD card.For all takeoff weights, V1 is 95 Kts.

WEIGHT

FLAPS

V 1

V 2

V YSE

CLEARANCE

T/O MIN. TORQ.

RWY

RQD

KING AIR

V R/

ATIS

TAKEOFF

Flight Planning



Accelerate-Stop – Flaps 0%

15°C

3,80

0 FT

Figure 4-27: Accelerate-Stop Distance Chart



Accelerate-Go Distance and V2Determine the distance required to clear a 35 ft obstacle afteraccelerating to V1 and then suddenly losing power on one engine; use the appropriate Accelerate-Go chart (flaps 0°) (Figure 4-28).1. Enter the chart from the lower left at the ambient temperature (15°C).

Move up to the pressure altitude (4,800 ft).2. Move to the right to the takeoff weight reference line, then follow the

guidelines down and right to the actual takeoff weight. From that point, move to the right to the wind reference line.Because the weight reference line is also the intersection for the takeoff weight, 12,500 lbs (5,669.9 kg), move straight to the right to the wind reference line.

3. Parallel the guidelines (down and right for headwind) to the appropriate wind component (15 Kts), then move to the right to the edge of the chart and read the accelerate-go distance.For this example, a distance of 8,200 ft is required.

4. Consult the table above the chart and determine V2 for the takeoff weight. Record this value on the TOLD card.For 12,500 lb takeoff weight, V2 is 121 Kts.

WEIGHT

FLAPS

V 1

V 2

V YSE

CLEARANCE

T/O MIN. TORQ.

RWY

RQD

TAKEOFF KING AIR

V R/

ATIS

Flight Planning



Accelerate-Go – Flaps 0%

15°C

8,20

0 FT

Figure 4-28: Accelerate-Go Chart



Single Engine Climb GradientIf an engine fails after V1, the aircraft’s climb gradient becomes especially critical if an obstacle is in the first or second segment. Use the Climb – One Engine Inoperative chart (Figure 4-29) to determine the single engine climb gradient.1. Enter the chart from the bottom left at the ambient temperature

(15°C). Move up to the pressure altitude (4,800 ft).2. Move to the right to the takeoff weight reference line, then parallel

the guidelines up and right to the takeoff weight.Because the weight reference line is also the intersection for the takeoff weight, 12,500 lbs (5,669.9 kg); proceed to step 3.

3. From that point, move straight to the right to the edge of the chart. Read the rate of climb and the climb gradient.For this example, the rate of climb is 650 ft/min with a 4.6% gradient.

4. Consult the table at the top of the chart and read the climb speed appropriate for the takeoff weight.For a takeoff weight of 12,500 lbs (5,669.9 kg), climb speed is 121 Kts.

Flight Planning



Climb – One Engine Inoperative

650

FT /

MIN

15°C

4.6

%

Figure 4-29: Engine Inoperative Chart



Landing Distance RequiredLanding distance varies with landing weight, flap setting, and use of propeller reverse. Use the appropriate Normal Landing Distance chart (without propeller reverse, flaps 100%) (Figure 4-30) to determine the distance required for the existent conditions.

Remember that the planned landing weight at the alternate is 10,131 lbs (4,595.3 kg).

1. Enter the chart from the bottom left at the ambient temperature (15°C). Move up to the pressure altitude (510 ft).

2. Move right to the landing weight reference line. Follow the guidelines down and right to the landing weight (10,131 lbs (4,595.3 kg)). From that point, move straight to the right to the wind component reference line.

3. Follow the guidelines (down and right) to the wind component, 10 Kts, then move straight to the right to the obstacle height reference line.

4. Follow the guidelines up and right to the edge of the chart. Read the distance required from 50 ft above the runway to a full stop (2,050 ft).

5. Return to the intersection at the obstacle height reference line. Move straight to the right to the edge of the chart to determine the distance required for ground roll without obstacle (1,200 ft).Because the runway at the alternate is 6,200 ft, landing distance is not limiting.

ATIS

WEIGHT

BALKED

VREF

VAP

RWY

RQD

KING AIR

FT.

VZF

100

VYSE

GO

AROUND

NOTES:

APPROACH

Flight Planning



Normal Landing Distance without Propeller ReversingFlaps 100%

15°C

10,1

31 L

BS

10 K

TS

2,05

0 FT

1,20

0 FT

Figure 4-30: Normal Landing Distance Chart




Flight Planning



Supplemental InformationThis section promotes an understanding of flight planning issues not encountered during the usual flight planning process.Planning, consideration of all factors, and acknowledgment of the potential for change promote higher levels of safety and efficiency.

Climb GradientsThe aircraft is certified under FAR Part 23; however, Beechcraft includes several FAR Part 25 performance charts to assist in critical situations. Part 23 requires that takeoff weight with engine failure at V1/VR allow a positive rate-of-climb.While this weight assures clearing 35 ft at the end of the runway, an obstacle several hundred feet high at a distance of a mile or more from the airport requires more attention. Use the Net Gradient of Climb charts in AFM Section 5 to determine the performance (i.e., weight) to clear such an obstacle.

ISA ConversionBeing familiar with aircraft performance under standard (i.e., ISA) conditions at operating altitudes facilitates identification of performance issues under warmer than standard (e.g., ISA + 10°C) conditions, which tend to degrade performance.

Minimum Takeoff PowerThe performance of engines nearing hot section or overhaul times may deteriorate gradually, so that the performance loss may not be noticed readily. Charts indicating minimum takeoff power (with and without ice vanes extended) provide a guideline to the minimum torque acceptable to meet established data for performance. Performance chart data assumes average engines.

Primary Static System BlockedEffects in airspeed at normal cruise or approach speeds with blocked static ports are minimal; however, altimeter readings at approach speeds can be 75 ft higher than the actual aircraft altitude when using the alternate static system. Confirm these instrument errors prior to performing approaches.

Range vs . FuelAn engine loss does not adversely affect the range of the aircraft; cruise altitude and time are affected, but not range.



Short Field Landing DistancesData for short fields (with and without flaps, with and without propeller reverse) are available on several charts. Use the landing distance charts found at the end of AFM Section V.

Single Engine Service CeilingExcept for extreme atmospheric conditions, aircraft performance exceeds routing requirements within North America.

Takeoff PerformanceWhen deciding between approach flaps or no flaps for the takeoff, compare the performance figures, especially on high altitude/hot temperature takeoffs. Be sure to check the accelerate-go requirements and net gradient of climb.

Temperature Conversions – Indicated OAT vs . True OATIndicated Outside Air Temperature (OAT) does not always reflect true OAT. For example, the true temperature at 25,000 ft and 200 Kts is 8°C colder than the indicated temperature. The Pilot’s Operating Handbook provides tables to convert indicated OAT to true OAT.

Systems



SystemsSeveral chapters contain multiple systems to facilitate a more Systems coherent presentation of information. The systems covered are listed below in alphabetical order opposite the chapter in which they are located. ATA codes note in parentheses.

SYSTEM (ATA Code) CHAPTER

Air Conditioning (21) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PNEUMATICAircraft Structure (51) . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWBrakes (32) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LANDING GEARCommunications (23) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AVIONICSDimensions and Areas (6) . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWDoors (52) . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWElectrical (24). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ELECTRICALEmergency Equipment (25) . . . . . . . . . . . . . . . . . . . MISCELLANEOUSEngine (71) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTEngine Controls (76) . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTEngine Fuel and Control (73) . . . . . . . . . . . . . . . . . . . . POWERPLANTEngine Indicating (77) . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTEquipment/Furnishings (25) . . . . . . . . . . . . . . AIRCRAFT OVERVIEW Fire Protection (26) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .FIRE Flight Controls (27) . . . . . . . . . . . . . . . . . . . . . . . .FLIGHT CONTROLSFuel (28) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FUELFuselage (53) . . . . . . . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWHydraulics (29) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . LANDING GEARIce and Rain Protection (30) . . . . . . . . . . . . . . . . . . . . . .ICE AND RAINIgnition (74) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTLanding Gear (32) . . . . . . . . . . . . . . . . . . . . . . . . . . . . LANDING GEARLighting (33). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ELECTRICALNavigation (34) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AVIONICSOil (79) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTOxygen (35) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MISCELLANEOUSPitot/Static (34) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AVIONICSPneumatic (36) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .PNEUMATIC\Pressurization (21) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PNEUMATICPropeller (61) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . POWERPLANTStabilizers (55). . . . . . . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWStall Warning (27) . . . . . . . . . . . . . . . . . . . . . . . . .FLIGHT CONTROLSWarning Lights (33) . . . . . . . . . . . . . . . . . . . . . . . . . MISCELLANEOUS



Windows (56). . . . . . . . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEWWings (57) . . . . . . . . . . . . . . . . . . . . . . . . . . . . AIRCRAFT OVERVIEW

Aircraft Overview 5A



ContentsAircraft OverviewAircraft Description

Engines .....................................................................................................5A-5Fuselage ...................................................................................................5A-5

Nose Section .......................................................................................5A-6Center Section .....................................................................................5A-6Empennage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5A-13

Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5A-13Fuel Tank System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5A-14

Aircraft DimensionsDanger Areas

Radar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5A-17Engine Exhaust Plume/ Propeller Wake . . . . . . . . . . . . . . . . . . . . . .5A-17

Modifications and Service InstructionsModifications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5A-19Selected Service Instructions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5A-19Airworthiness Directives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5A-19




Aircraft Overview



Aircraft OverviewThis section presents an overview of the Super King Air 200 series aircraft (200/B200, 200C/B200C, 200T/B200T, 200CT/ B200CT). It includes major features, airframe structures, dimensions, and danger areas, as well as a list of the modifications and service bulletins discussed in this manual.This manual references the manufacturer’s series numbers or individual aircraft serial numbers, as applicable, and where system differences warrant, it publishes separate data and schematics. Please refer to Modifications in this chapter for aircraft serial numbers that correspond to each of the 200 model series.Serial numbers, assigned consecutively at construction, remain with the aircraft unless factory modifications place the aircraft in another category (e.g., BB-870 [200 series] became BL-36 [200C]).



Aircraft Features

WING TIP TANK

200T/B200T 1

1

SIDE WINDOW

RADOME

NACELLE

AILERON

AILERON TRIM TAB HORIZONTAL STABILIZER

AIRSTAIR DOOR

FLAPS

EMERGENCY EXIT

ELEVATOR

ELEVATOR TRIM TABS

DORSAL FIN

RUDDER

RUDDER TRIM TAB

STORM WINDOW

CARGO DOOR (200C/B200C)

Aircraft Overview



Aircraft DescriptionThe Super King Air 200/B200 series is a transport category aircraft, certified in accordance with FAR Part 23 airworthiness standards for single pilot, IFR, VFR, day, night, and icing conditions operation. The low-wing cabin class turboprop has an optimum range of approximately 1,800 nautical miles with full fuel with a maximum cruise speed of 290 Kts TAS. The 200 prior to BB-54 except 38, 39, and 44 is certified to an operating altitude of 31,000 ft and all other 200s to 35,000 ft.The aircraft has a retractable tricycle landing gear, which on models 200 and B200, S/N BB-1 through 1192 are electrically operated and on BB 1193 and subsequent are hydraulically operated and electrically powered.The 200T/B200T and 200CT/ B200CT series have optional wing tip fuel tanks.

EnginesTwo wing-mounted, Pratt & Whitney PT6A-41 (200) or PT6A-42 (B200) turbopropeller engines power the aircraft. Each engine produces 850 shaft horsepower and a maximum 2,230 ft-lbs of torque at sea level. Engine design includes anti-ice and fire protection systems.Depending on model, serial number or modifications, the aircraft uses three or four-bladed constant speed, full-feathering, reversible Hartzell or McCauley propellers. Supplemental Type Certificates (STCs) exist to replace three-bladed propellers with four-bladed units that provide increased performance and reduced noise.

FuselageThe all-metal, semimonocoque fuselage comprises the nose section, pressurized center section, and the aft fuselage.

Figure 5A-1: Fuselage



Nose SectionThe unpressurized nose section comprises the radome, which houses the radar antenna, many avionics components of the AC system and an equipment compartment between the radome and the forward bulkhead. A door on each side of the nose provides access to avionics, electrical, and other equipment.

Center SectionThe pressurized center section consists of the cockpit, cabin, and aft baggage compartment. On all aircraft S/N BB1439 and BB-1444 and subsequent, fuselage Dynamic Vibration Absorbers (DVAs) were installed to suppress noise and vibration.

CockpitThe crew compartment has dual flight controls and flight instrumentation; however, the aircraft is certified for single-pilot operation. A center console supports the power quadrant, pressurization controls, and automatic flight controls. Additional space is provided for optional equipment. Cockpit lighting controls are in an overhead panel. The left side panel supports fuel controls, indicators and related circuit breakers; additional circuit breakers are on the right side panel.Each crewmember’s seat is adjustable vertically, forward, and aft; adjustment controls are under the forward edge of the seat. In addition, the armrests pivot at the aft end and raise up to facilitate cockpit entry and egress.The seatbelt/shoulder harness is on an inertia reel that keeps the harness snug but allows normal movement during flight operations. A locking device in the reel secures the harness if there is sudden forward movement or impact.Sliding door panels separate the cockpit from the cabin area.

Figure 5A-2: Cockpit

Aircraft Overview



Windshield/Cockpit WindowsThe two-piece layered glass windshield is mounted to form an integral part of the pressure vessel. The inner and outer layers are glass; electric heating elements are bonded to an inner polyvinyl layer. Warm air outlets defog the interior windshield. An electrically operated windshield wiper has two speeds. (See the Ice and Rain chapter.)Two storm windows and two side windows, one of each on each side of the cockpit, are of single-layer acrylic plastic. Both the left and right storm windows open inward; the crew side windows do not open.

Figure 5A-3: Two-Piece Layered Glass Windshield

Cabin SeatingThe cabin can accommodate up to 13 passengers when utilizing the foyer and rear baggage area. Configurations vary in each aircraft, depending on the owner’s choice. A typical seating arrangement accommodates seven passengers in the forward cabin.

Figure 5A-4: Cabin Seating Arrangement



Combinations of non-adjustable couches and adjustable passenger chairs are available for installation on continuous tracks mounted in the cabin floor. Adjustable headrests are on all aft-facing chairs and on all forward-facing chairs with optional shoulder harnesses. All passenger seats have adjustable seat belts.Oxygen masks stowed in the headliner deploy automatically in the event of cabin depressurization. In addition, the crew has cockpit controls to manually deploy passenger masks as required.

Cabin WindowsOn the 200, B200, 200T, and B200T, there are 10 circular, layered acrylic plastic windows in the forward passenger cabin area, five on each side of the aircraft. On the 200C, B200C, 200CT, and B200CT, the fifth window on the left fuselage is incorporated into the cargo door. On all models, the forward right window is in the emergency exit, and there are two oval-shaped windows in the aft cabin area, one on each side of the aircraft. In addition, an optional window may be in the toilet compartment, which is opposite the entry door.

Figure 5A-5: Cabin Windows

Two acrylic polarizing panes in each cabin window provide light adjustment; the inner pane rotates to regulate light transmission.

NNTTE: See the Miscellaneous chapter for operation of the 200C/B200C cargo door and built-in airstair door.

Aircraft Overview



ToiletA partitioned toilet compartment is in the foyer opposite the entry door. Hinged seat-cushions and a seat belt mounted on top of the electric flushing toilet form an extra passenger seat when the toilet is not in use.

Aft Cabin Area/ Baggage CompartmentThe entire cabin area aft of the foyer is normally used as a baggage compartment; items stowed there are accessible during flight. The aft cabin area may be separated from the foyer by a nylon webb, and, if installed, an optional baggage compartment curtain that pulls across the opening and secures with snap fasteners. One or two optional folding seats mount on the aft cabin baggage area sidewall.

Figure 5A-6: Aft Cabin Area/Baggage Compartment



Cabin Airstair Entrance Door (200/B200)

WARIRN If any condition specified in this door-latching procedure is not met, do not take off.

WARIRN The door opening button can be overpowered resulting in the door opening while pressurized.

WAUIIR Only one person should be on the airstair door stairway at any one time.

The bottom-hinged cabin entrance door, just aft of the left wing trailing edge, opens outward and down. A hydraulic damper slows the door opening for safety, and integral steps fold out when the door is down. A vinyl-encased cable supports the open door, serves as a handrail and a door pull for closing from inside.

Figure 5A-7: Bottom-Hinged Cabin Entrance Door

When weight is off the landing gear, engine bleed air supplies pressure to inflate the door seal, which provides a positive pressure-vessel seal around the door. The outside door handle key-locks for security of the aircraft on the ground.To open the door, press the release button adjacent to either the inside or outside handle; rotate the inside handle counterclockwise or the outside handle clockwise. The release button contains a differential pressure-sensitive mechanism to prevent the door from opening when the aircraft is pressurized. Extreme caution should be exersized when opening the door when any possibility of pressurization exits, as this pressure sensing mechanism can be over powered resulting in the door opening while pressurized. This could result in injury or death. Extreme caution should be exercized when any possibility of pressurization exists.

Aircraft Overview



To close the door from outside the aircraft, lift up the free end of the door and push it up against the door frame as far as possible. Grasp the handle with one hand and rotate it clockwise as far as it will go. After the door moves into the fully closed position, rotate the handle counterclockwise to point aft, as far as it will go; the release button should pop out. Check the security of the door by attempting to rotate the handle clockwise without pressing the release button; the handle should not move.To close the door from inside the aircraft, grasp the handrail cable and pull the door up against the door frame. Grasp the handle with one hand and rotate it counterclockwise as far as it will go, continuing to pull the door inward. After the door moves into the fully closed position, turn handle clockwise to point downward as far as it will go. Check the security of the door by attempting to rotate the handle counterclockwise without pressing the release button; the handle should not move.Next, lift the folded stairstep that is just below the handle. An observation window with a diagram shows proper configuration when in locked position. The red portion should appear around the locking mechanism. A push type light switch is available to assist in low light conditions. A round observation window is at each corner of the door. Observe the bayonet locking pins are properly aligned (green to green). Press the red pushbutton adjacent to the observation port to illuminate the locking mechanisms inside the door and visually check through the window that the arm is properly positioned around the shaft, as shown on a placard on the door.Finally, switch the battery switch on and verify that the CABIN DOOR annunciator is extinguished. The CABIN DOOR annunciator illuminates if the battery switch is on and the airstair door is not fully closed or if the cabin door handle is rotated to the open position. Prior to the first flight of the day, perform the Cabin/Cargo Door Circuitry Check in the Expanded Normal Procedures chapter.

Figure 5A-8: Exit-Lock Handle



Emergency ExitThe emergency exit door is on the right fuselage at the forward end of the passenger compartment. To open from the inside, pull down on the EXIT-PULL handle. To open from outside, release the door with a flush-mounted, pull-out handle. The non-hinged, plug-type door removes completely from the frame into the cabin.On the 200/B200, the emergency exit door can be locked from inside the cabin to ensure security when the aircraft is parked. To lock, place the lock lever in the down position. Ensure this lock lever is in the up, or unlocked, position prior to flight to allow removal of the door from the outside in the event of an emergency. Removal of the door from the inside is possible at all times by using the EXIT PULL handle, which is not locked by the lock lever. An exit lock placard on the lock lever is visible when the lever is in the locked position.On the early 200s, the emergency exit door key-locks from the inside to prevent opening from the outside. However, the inside handle unlatches the door, whether it is locked or not, by overriding the locking mechanism. Unlock this key lock prior to flight to allow removal of the door from the outside in the event of an emergency. The keyhole is horizontal when the door is locked; the key cannot be removed from this position.A wiper-type disconnect for the air duct that supplies air to the eyeball outlet in the emergency exit door is on the upper aft edge of the door. As the door is removed, the duct disconnects because it is an integral part of the door. In addition, an electrical disconnect for the wiring that goes to the reading light and the fluorescent light in the emergency exit door unplugs when the door is removed. When reinstalling the door, reconnect the electrical disconnect before moving the door into the closed position.

Figure 5A-9: Emergency Exit Door Figure 5A-10: Exit pull handel

Aircraft Overview



TmpennageThe T-tail empennage features a top-mounted horizontal stabilizer, cable-operated elevators with electric and/or manual pitch trim tabs, and a manually controlled rudder and rudder trim tab. A pneumatic rudder boost system is available for use in engine-out emergencies. (See the Flight Controls chapter.)

Figure 5A-11: T-tail Empennage

WingEach all-metal wing assembly consists of a conventional design, semi-monocoque box construction center section and two-panel outboard wing section. The center section, which contains an auxiliary fuel cell, is an integral part of the fuselage. The outboard wing section contains five interconnected fuel cells.

Figure 5A-12: All-Metal Wing



On aircraft S/N BT-1 through BT-30, BL 1 through BL-72, BN-1 through BN-4, and BB-2 through BB-1192, except BB-1158 and BB-1167, each wing panel attaches to the center section with four tension bolts. S/N BB1158 and BB-1167 and subsequent models have a three-element spar in the center wing section. In the three-spar configuration, each outboard wing attaches to the center section by three tension bolts and a shear bolt at the lower forward attach point.The leading edge assembly and main outboard wing assembly join together at the main spar with continuous hinges. Engine controls, anti-ice bleed air plumbing, and electrical wiring route through the wing root leading edge.An outboard aileron with trim tab, an outboard flap, and an inboard flap are on the wing trailing edge. A subspar forward of the rear spar provides a tunnel for control cables and shafts and serves as a fuel wall for the bladder fuel cells from the root rib to the nacelle.Landing gear structural supports in the nacelles are made of machined alloy plate. Sheet metal formers and stringers establish the nacelle fairing and a cavity for a bladder fuel cell above and forward of the wheel well. The main gear drag leg is bolted to an aluminum alloy forging attached to the main spar of the center section.

Fuel Tank SystemThe aircraft has 14 fuel cells with a total capacity of 544 U.S. gallons. Each engine’s separate fuel system interconnects through a crossfeed line and is subdivided into a main and an auxiliary system.The main fuel system consists of four interconnected, bladder fuel cells and one integral fuel cell in each wing and a fuel cell in each nacelle. The fuel filler port is near the wing tip. The auxiliary system consists of a center wing section auxiliary fuel cell with its own filler port.In addition, if two optional 53 U.S gallon wing tip tanks are installed, total fuel capacity increases to 650 U.S. gallons. See the Fuel System chapter for detailed description.

Aircraft Overview



Aircraft DimensionsRadome to Rudder

200/B200, 200T/B200T, 200CT/B200CT . . . . . . . . . . . . . . . . . 43´ 9˝ 200C/B200C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43´ 10˝

Nose Gear to Main Gear 200/B200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15´ 0˝200C/B200C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .14´ 111/2˝200T/B200T, 200CT/B200CT . . . . . . . . . . . . . . . . . . . . . .14´ 112/5˝Main Gear to Main Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17´ 2˝

Wing Tip to Wing Tip 200/B200, 200C/B200C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54´ 6˝200T/B200T, 200CT/B200CT . . . . . . . . . . . . . . . . . . . . . . . . . 56´ 7˝Horizontal Stabilizer, tip to tip . . . . . . . . . . . . . . . . . . . . . . . . . 18´ 5˝

Ground to Top of Vertical Stabilizer 200/B200 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15´ 0˝200C/B200C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14´ 10˝200T/B200T, 200CT/B200CT . . . . . . . . . . . . . . . . . . . . . . . . . 14´ 6˝

Pressurized Cabin Length, pressure bulkhead to pressure bulkhead . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22´ 0˝

Cabin Interior Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4´ 61/2˝Cabin Interior Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4´ 6˝Cabin Entry Airstair Door Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2´ 2˝Cabin Entry Airstair and Cargo Door Height . . . . . . . . . . . . . . . . . . 4´ 2˝Cargo Door Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4´ 4˝

1

2

3

4



Aircraft Dimensions

4

1

2

3

18' 5"

17' 2"

8' 2.5" DIA.

Aircraft Overview



Danger AreasRadar emissions from the radome, engine exhaust plume, and the propellers are the primary dangers around the aircraft.

RadarWhen the weather radar is operating, emissions are hazardous up to a 30 ft radius of the radome, depending on the radar equipment installed. For safety, do not operate radar on the ground except for brief systems tests; advise ground personnel in advance to stay clear of the radome during these tests.

Tngine Txhaust Plume/ Propeller WakeExhaust hazards lie in velocity and plume temperature. The engine exhaust danger area extends to approximately 18 ft aft of the exhaust stacks with engines at idle, or to 40 ft with engines at maximum power.Advise ground personnel of imminent engine starts. Do not start an engine without verifying that the immediate area aft and forward of the aircraft is clear of ground personnel, small articles, or sensitive equipment.

MAXPOWER53 KTS

30 FEET

RADIATION DANGERAREA

EXHAUSTDANGERAREA

PROPELLERDANGERAREA

40'18'0'

GROUNDIDLE10 KTS

390°C

EXHAUSTDANGERAREA

Figure 5A-13: Danger Areas Around the Aircraft




Aircraft Overview



Modifications and Service Instructions

Modifications200 Series: S/N BB-002 through 733, 735 through 792, 794 through 828, 830 through 853, 871 through 873, 892, 893, 895, 912, and 991.B200 Series: S/N BB-734, 793, 829, 854 through 870, 874 through 891, 894, 896 through 911, 913 through 990, 992 through 1313, 1135 and subsequent.200C Series (with cargo door): S/N BL-1 through 36.B200C Series: S/N BL-37 and subsequent, BP-64 and subsequent, BU-1 and subsequent, BV-1 and subsequent, BW-1 and subsequent.200T Series (with wing tip tanks): S/N BT-001 through 022 and 028.200CT Series (with cargo door and wing tip tanks): S/N BN-1.B200CT Series: S/N BN-002 and subsequent.

Selected Service InstructionsSI-0701-356: Electrical Battery Installation – Installation of an improved battery charge current detector.SI-1047: Landing Gear – Modification of the landing gear system. S/N BB-324 through BB-453.SI-1121-II: Flight Controls – Installation of improved flap switches. See Service Instructions 1121-II for effectivity.SI-0968-II: Air Conditioning – Installation of high pressure plumbing and a new high pressure switch. S/N BB-002 through 344, except 001, 003 through 005, 034, 123, 186, 203, and 270.

Airworthiness DirectivesAD 96-09-13: Severe Icing.AD 97-25-03: Power Levers.AD 98-20-38: Flight in Icing Conditions.




Avionics 5B

King Air 200 5B-1April 2011


ContentsAvionicsAvionics Equipment

Avionics Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-5SchematSc: Pitot/Static System Non-ADC Equipped 5B-7SchematSc: Pitot/Static System ADC Equipped 5B-8

Flight Environmental DataPitot/Static System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-10

Airspeed Indicator ..............................................................................5B-12Vertical Speed Indicator .....................................................................5B-12Altimeters ...........................................................................................5B-13Altitude Alerter ...................................................................................5B-14Radio Altimeter ..................................................................................5B-15Flight Profile Advisory ........................................................................5B-15Outside Air Temperature ....................................................................5B-16

CommunicationAudio Integrating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-19

Microphones/Headsets ......................................................................5B-19Audio Control Panels .........................................................................5B-20

Speech . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-20VHF Communications ........................................................................5B-21Audio Emergency ..............................................................................5B-21Emergency Operation ........................................................................5B-21HF Communication ............................................................................5B-22Radio Telephone ................................................................................5B-22

Static Discharging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-22Navigation

Attitude and Direction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-23Magnetic Compass ............................................................................5B-23Turn and Slip Indicators .....................................................................5B-24Gyro-Horizon/Vertical Gyro ................................................................5B-25Compass System ...............................................................................5B-26Radio Magnetic Indicators .................................................................5B-27

Position Determining . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-27

King Air 2005B-2April 2011


Instrument Landing ............................................................................5B-27VHF Navigation ..................................................................................5B-28Automatic Direction Finder ................................................................5B-28Distance Measuring ...........................................................................5B-29Area Navigation .................................................................................5B-29Transponder .......................................................................................5B-29Omega/VLF .......................................................................................5B-29LORAN ..............................................................................................5B-30Global Positioning System .................................................................5B-30Flight Management System ..............................................................5B-31Weather Radar ..................................................................................5B-31

Flight Control SystemAutopilot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-33

Autopilot Controller ............................................................................5B-34Autopilot Computer ............................................................................5B-35Servos ................................................................................................5B-36

Flight Director . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-36Flight Computer .................................................................................5B-36Mode Coupler ....................................................................................5B-36Mode Selector ....................................................................................5B-37Attitude Director Indicator ..................................................................5B-37Horizontal Situation Indicator .............................................................5B-38

Preflight and ProceduresPreflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5B-39Abnormal and Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . .5B-39

Emergency Procedures .....................................................................5B-39Pilot’s Static Air System .....................................................................5B-39Autopilot Malfunction .........................................................................5B-41

Avionics



AvionicsThis chapter provides a brief overview of the following: pitot/static system and instruments communication equipment navigation equipment flight control systems.

Besides these specific areas, this chapter includes instrumentation not addressed in other chapters.Cockpit panel art at the front of the chapter provides a ready reference to locate instruments and equipment addressed in this discussion.Several different EFIS installations are available. If your aircraft is so equipped, see appropriate supplement to AFM.For a detailed description and operating procedures for a particular piece of avionics equipment, refer to the applicable pilot’s guide.



Cockpit Forward Panel

ELEVTRIM

CABINPRESSDUMP

TEST

RUDDERBOOST

OFF

PRESS

HDGCRS

HDG CRS

TTGSPD

ET

ALT SEL ALT VS IAS

1/2 BANK HDG NAV APPR

TURB

DN

UP

TURN YD AP

DISENGAGED

ENGAGED

ABC1

DEF2

GHI3

JKL4

MNO5

POR6

STU7

VWX8

YZ*9

HOLD #0 BACK

MSC DIM

BAT

OFF

ON

NAV DATA FPL ALPHA

ENTER

SELF TEST

7

0 15

1311

10

98

12

34

56

1

35

3025

2321

1917

15

ACFT ALT1000 FT

1000FT

CABINALT

CABINALT

RATE

MAX

MIN

0

5

3

11

3

5

0

5

3

11

3

5

FLAP

UP

5

0

DN

DN

5

DN

ELEVA

TAB

UP

GO AROUND

POWER

LIFT

INCR

PROP

LANDINGAND

REVERSE

IDLE

REVERSE

FRICTIONLOCK

CAUTIONREVERSE

ONLY WITHENGINE

RUNNING

AILERON TABLEFT RIGHT

FRICTIONLOCK

UP

FLAP

APPROACH

DOWN

RUDDER TABLEFT RIGHT

CONDITION

HIGHRPM

HIGH IDLE

LOWIDLE

FUELCUT OFFFEATHER

WARNING DE-PRESSURIZE CABINBEFORE LANDING

NO

TRANSFER

NO

TRANSFER

PRESS TO TEST

STANDBY PUMPON

OFFAUX TRANSFER

OVER RIDE

AUTO

FUEL QUANTITYMAIN

AUXILIARY

CROSSFEEDFLOW

OFF

ENGINE ENGINE

SEE MANUAL FORFUEL CAPACITY

PRESS TO TEST

STANDBY PUMPON

OFFAUX TRANSFER

OVER RIDE

AUTO

5

FIREWALLVALVE

10

STANDBYPUMP

5

AUXTRANS

FERLEFT

5

QTYIND

5

PRESSWARN

5

CROSSFEED

5

PRESSWARN

5

QTYIND

5

AUXTRANS

FERRIGHT

10

STANDBYPUMP

5

FIREWALLVALVE

50 50 50 50 5 20 20 20 5 5 5 5 5 5

OPEN

CLOSED

FIREWALLSHUT OFF VALVE

OPEN

CLOSED

FIREWALLSHUT OFF VALVE

FUEL SYSTEM

NO. 3 NO. 4BUS FEEDERS PROP DE-ICE FLAP PROP IGNITOR START

CONTROL PROP

LEFT

PROP

RIGHT

MOTOR CONTROL POWERGOV C ONTROL

LEFT RIGHT LEFT RIGHT

0

2

4

6 810

12

14

LEFT

FUEL

MAIN TANKONLY

QTYLBS x 100

0

2

4

6 810

12

14

RIGHT

MAIN TANKONLY

FUEL

QTYLBS x 100

Avionics



Avionics EquipmentAvionics equipment includes those subjects covered by Air Transport Association (ATA) chapters 22 – Auto Flight, 23 – Communications, and 34 – Navigation.

Avionics PowerThe majority of the avionics equipment on the Beech Super King Air operates on 28 VDC supplied by the aircraft battery and the generators. Three avionics buses (Avionics No.1, Avionics No.2, and Avionics No. 3) distribute power from these sources through individual circuit breakers to the various equipment. A single AVIONICS MASTER PWR switch on the pilot’s instrument panel controls relays that supply power from L/R Generator buses. A 5A Circuit Breaker (CB), labelled AVIONICS MASTER sup-plies power from the No.1 Dual-Fed bus to the switch. The switch, in turn, controls the operation of three avionics power relays.

Figure 5B-1: AVIONICS MASTER PWR Switch

With the battery switch ON power flows to the three avionics buses and passes through the normally closed avionics relay. This power subsequently opens the relay to remove power from the avionics buses. Moving the AVIONICS MASTER switch to ON removes power from the relay; the relay closes and the avionics buses receive power.If the AVIONICS MASTER switch fails, pulling the AVIONICS MASTER CB restores power to the avionics buses.



The Ground Communication Electric Power Bus is offered as an option. This feature reduces battery requirements when obtaining an IFR clearance prior to engine start. Pressing the on the center instrument panel (typical location) supplies power through relays from the Hot Battery bus to the No. 2 COMM and the No. 2 audio panel.

Figure 5B-2: GND COMM Button

Another option for the aircraft is an Auxiliary DC bus system. This system provides essential power to some avionics equipment during electrical bus failure or when electrical load-shedding is a required action. After turning the generator and battery switches OFF, turning the AUX DC BUS switch ON supplies power directly from the Hot Battery bus through relays to the No. 2 COMM, No. 2 compass system, No. 2 audio panel, the ILS portion of the No. 2 NAV, and the glareshield floodlights.

Figure 5B-3: AUX DC BUS Switch

Avionics



Pitot/Static System Non-ADC Equipped

PITOTHEATSWITCH

ALTIMETER

VERTICALSPEEDINDICATOR

CABINDIFFERENTIALPRESSUREGAGE

ALTERNATESTATIC AIRSELECTORSWITCH

ALTERNATESTATIC AIR(PILOT ONLY)

PILOT'S PITOT

COPILOT'S PITOT

PILOT'S STATIC

COPILOT'S STATIC

ALTERNATE STATIC

STATICPORTS

STATICPORTS

N O.

1

DUAL

FED

BUS

AIRSPEEDINDICATOR

PILOT'SSIDE

COPILOT'SSIDE

AFT PRESSURE BULKHEAD

N O.

2

DUAL

FED

BUS

7.5A 7.5A



Pitot/Static System ADC Equipped

STATICPORTS


PITOTHEATSWITCH

AIRSPEEDINDICATOR

ALTIMETER


CABINDIFFERENTIALPRESSUREGAGE

ALTERNATESTATIC AIRSELECTORSWITCH

ALTERNATESTATIC AIR(PILOT ONLY)

PILOT'S PITOT

COPILOT'S PITOT

PILOT'S STATIC

COPILOT'S STATIC

ALTERNATE STATIC

AFCSAIR DATA

COMPUTER

AIRSPEEDINDICATOR

ALTIMETER


STATICPORTS

PILOT'SSIDE

COPILOT'SSIDE

N O.

1

DUAL

FED

BUS

N O.

2

DUAL

FED

BUS

7.5A 7.5A

Avionics



Flight Environmental DataFlight environmental data equipment senses environmental conditions to determine aircraft altitude and airspeed. This includes the pitot/static system and air data computer (if installed). The pitot/static system and air data computer, in turn, drive: altimeters altitude alerter airspeed indicators vertical speed indicators.

Although not part of the pitot/static system, the radio altimeter and flight profile advisory system (if installed) are part of the flight environmental data system. The radio altimeter determines altitude and the flight profile advisory system uses altitude information from the radio altimeter to provide verbal warnings of aircraft altitude and unsafe flight conditions. In addition, this section discusses the outside (free) air temperature gauge.



Pitot/Static SystemThe pitot/static systems provide ram air pressure and static pressure data from their respective pitot mast and static ports. The pitot masts have electrical heating elements that provide anti-icing; the static ports are unheated. The LEFT and RIGHT PITOT switches in the ICE PROTECTION switch quadrant control 28 VDC power to the pitot mast heating elements. Do not use pitot heat for extended periods on the ground as it may damage the heating elements.

Figure 5B-4: Pitot Tube

Figure 5B-6: LEFT and RIGHT PITOT Switches

Figure 5B-5: Static Ports

Avionics



To prevent yaw error, the pilot’s and copilot’s static system have a port on each side of the fuselage aft of the aft pressure bulkhead. Each pair of ports connects to a single line that carries static pressure to the front of the aircraft. The pilot’s static system has a selectable alternate static source on the lower part of the copilot’s CB panel.The pitot system supplies ram air pressure to the pilot’s and copilot’s -Airspeed Indicators (ASIs). On S/N BB-324 through BB-452 without SI 1047, a pressure differential switch supplied by the copilot’s pitot system; is part of the landing gear warning -system.The static system supplies ambient air pressure to the airspeed indicators, Vertical Speed Indicators (VSIs), altimeters, and the cabin differential pressure gauge. If the pilot’s static -system ports clog or a leak develops in the lines, moving the PILOT’S -STATIC AIR SOURCE valve handle aft to ALTERNATE selects the static port in the aircraft tailcone aft of the pressure bulkhead as the pilot’s static source. Because the static port is in the aircraft fuselage, the pilot’s altimeter and airspeed indicator require correction to obtain accurate indications. Refer to the applicable charts in the Aircraft Flight Manual Performance Section.

Figure 5B-7: PILOT’S STATIC AIR SOURCE Valve Handle

Aircraft equipped with a flight director have an Air Data Computer (ADC) that receives pitot/static pressure from the pilot’s system. The ADC electrically drives the pilot’s ASI, VSI, and altimeter. The ADC also supplies air data information to various other equipment including the transponder, altitude alerter, and the flight control system.Two drains for the pilot’s, pilot’s alternate, and copilot’s static system are behind an access door on the lower right cockpit wall. After the aircraft has been exposed to rain or high humidity, drain the static lines. The pitot system lines do not require drains as they are the lowest point in the pitot system; moisture drains out through the masts.



Atrsphhd IndtSmaorTypically, the pilot’s and copilot’s airspeed indicators are identical; both indicate airspeed from 40 to 300 Kts with a pointer moving over a circular scale. An aneroid driven pointer (barber pole) indicates the maximum allowable airspeed (VMO) depending on aircraft altitude. If the aircraft exceeds VMO, there is no audible warning.

Figure 5B-8: Airspeed Indicators

On aircraft with an ADC, the pilot’s airspeed indicator is a servo driven instrument that receives electrical driving signals from the ADC. The instrument displays Indicated Airspeed (IAS) and maximum allowable airspeed (VMO). If the unit loses power or the ADC fails, an OFF flag appears on the indicator.

VhratSml phhd IndtSmaorThe pilot’s and copilot’s Instantaneous Vertical Speed Indicators (IVSIs) use an internal accelerometer to reduce the time lag between pressure change (vertical acceleration) and instrument indication. Each IVSI displays aircraft vertical speed from 0 to 4,000 feet per minute (fpm), up or down.

Figure 5B-9: Instantaneous Vertical Speed Indicators

Avionics



On aircraft with an ADC, the pilot’s Instantaneous Vertical Speed Indicator (IVSI) receives driving signals from the ADC. The IVSI displays vertical speed from 0 to 6,000 fpm, up or down. An OFF flag appears on the instrument face if the unit loses power or the ADC malfunctions.

AlatehahrsThe pilot has a servo-driven encoding altimeter and the co-pilot has a conventional barometric (pneumatic) altimeter. Each displays aircraft altitude with three moving pointers or a moving pointer and drum counters.

Figure 5B-10: Servo-Driven Encoding

Altimeter

The pilot’s altimeter provides encoded altitude information as an electrical signal to the altitude alerter and to the transponder for Mode C/S altitude reporting. If the encoder loses power or fails, an OFF flag appears on the instrument face.The copilot’s altimeter has a 28 VDC powered internal vibrator that overcomes instrument friction and ensures instrument accuracy; the vibrator prevents the pointer(s) from sticking.On aircraft with an ADC, the pilot’s altimeter is a servo-driven instrument that receives electrical signals from the ADC. If the altimeter loses power, the ADC malfunctions, or the encoder loses power, an OFF flag appears on the instrument face.All altimeters have provisions for adjusting the instrument to local barometric pressure in inches of mercury (In Hg) and millibars. On encoding altimeters, adjusting the instrument to local pressure has no effect on the encoded altitude information; it is always relative to standard pressure (29.92 In Hg/1,013 millibars).

Figure 5B-11: Barometric Altimeter



Alataudh AlhrahrAn altitude alerter provides visual and aural warnings when the aircraft approaches or deviates from a selected altitude; these systems receive altitude information from the pilot’s encoding altimeter or the ADC. The altitude alerting unit has controls for setting a desired altitude, a display for the set altitude, and a visual warning light. Usually, the system allows the setting and alerting of an altitude from -900 to 37,000 ft. The unit provides visual warnings through a light on the unit and aural warnings through the cockpit speakers.

Figure 5B-12: Altitude Alerting Unit

Once the aircraft altitude is within 1,000 ft of the selected altitude, the altitude alerting light illuminates and the alert horn sounds for two seconds. The light extinguishes once the aircraft is within 300 ft of the selected altitude. If altitude deviates more than 300 ft from the set altitude, the light illuminates and the horn sounds. Correcting the error or selecting a new altitude extinguishes the light. If the unit loses power, an OFF flag appears.

Avionics



Rmdto AlatehahrRadio altimeters provide precise aircraft altitude Above Ground Level (AGL) during approach and landing or when the aircraft is below 2,000 or 2,500 ft AGL (depending on the system installed). The system consists of a transceiver, indicator, and transmit and receive antennas on the lower fuselage. The transceiver transmits a 4250 to 4350 (4.3 GHz ±5 MHz) signal toward the ground, receives the bounced signal, and computes altitude by the time delay between transmission and reception.

Figure 5B-13: Radio Altimeter

The system provides absolute altitude information from -20 to 2,000 or 2,500 ft AGL in 100 ft graduations down to 500 ft and in 10 ft graduations from 500 to -20 ft on an indicator on the pilot’s instrument panel and/or the attitude director indicator. The radio altimeter system also provides altitude information to the flight profile advisory system.The radio altimeter allows the pre-selection of a desired Decision Height (DH) for altitude alerting purposes. To preselect the DH, rotate the DH SET knob until the index aligns with the desired altitude. Once the aircraft reaches the desired altitude, the DH light illuminates.

Flight Profile AdvisoryWhen installed, the flight advisory system works with the radio altimeter and landing gear squat switches to provide a verbal warning of unsafe flight conditions.If the aircraft descends to 2,000 ft, the system provides a “RADIO ALTITUDE” warning. As the aircraft continues to descend past 1,000 ft radio altitude, the system announces “ONE THOUSAND” and provides a verbal announcement of radio altitude every one hundred feet. Setting a decision height on the radio altimeter provides a “MINIMUM” announcement once the aircraft reaches the preselected radio altitude.If the aircraft deviates more than one dot from the localizer, the system announces “LOCALIZER” three times; for a deviation of more than one dot from the glideslope, the system announces “GLIDESLOPE” three times. The warnings continue until the crew corrects the deviation from the localizer or glideslope path.



If the aircraft descends below 500 ft with the landing gear retracted, the system announces “CHECK GEAR” three times. These warnings repeat every 100 ft until gear extension.An FPA ADVSY switch and audio control for the flight profile advisory system are on the center instrument panel.

Figure 5B-14: FPA ADVSY Switch and Audio Control

Ouastdh Atr ThephrmaurhAn Outside Air Temperature (OAT) gauge, on the overhead panel (early aircraft) or on the left hand sidewall panel near the pilot’s seat consists of a temperature probe that protrudes through the aircraft skin and a gauge calibrated in degrees Fahrenheit and Centigrade. Post lights illuminate the gauge.

Figure 5B-15: OAT Gauge

Avionics



The digital outside air temperature system installed on BB-1439, 1444 and subsequent, BT-39 and subsequent, BL-139 and subsequent and BN-5 and subsequent consist of a probe remote processor mounted under the copilot’s floor and a digital display indicator mounted on the pilot’s left side panel.

Figure 5B-16: Digital Display Indicator




Avionics



CommunicationCommunications equipment on the King Air includes: audio integrating speech communication interphone static discharging.

Audio IntegratingAudio integrating includes the audio control panels, headsets, microphones, and cockpit loudspeakers.

MtSropconhs/HhmdshasThe microphone and headset jacks are on the pilot’s and copilot’s side panels. Each jack connects with the respective audio control panel. A microphone switch on the pilot’s and copilot’s subpanels selects the normal microphone or the oxygen mask microphone.

Figure 5B-17: Microphone and Headset Jacks



Audto Conarol PmnhlsAudio control panels for the pilot and copilot contain controls for audio source selection, microphone output selection, and volume control of the headsets and cockpit loudspeakers. Each panel accepts audio inputs from the communication and navigation receivers and provides individual switches for each audio source.

Figure 5B-18: Audio Control Panels

A master volume switch on each panel individually adjusts the volume of the combined audio sources to the headsets or loudspeakers. Depending on the audio panel installed, the cockpit loudspeakers and headphones are always operational (headphones plugged in) or are not operational at all times (either loudspeaker or headphones). One system has a switch that turns off the cockpit loudspeaker; the other system has a switch that selects either the headphones or the loudspeakers.Each audio panel has an emergency switch that bypasses the audio amplifiers to connect audio output directly from the source to the headsets if the audio panel malfunctions. If the emergency switch fails, pull the applicable audio control circuit breaker(s).Microphone output selection is made through a rotary selector switch on each panel that connects the microphone to the communications transmitters or the cabin speakers. A second microphone switch selects normal microphone operation or hot microphone operation. A hot microphone allows communication between the crew through headsets without -having to key the headset microphone.

SpeechSpeech communication includes: Very High Frequency Communication (VHF COM) transceivers High Frequency communication (HF) transceivers radio telephone.

Avionics



VHF CoeeuntSmatonsTypical VHF transceivers provide air-to-air, air-to-ground, and ground-to-ground communications. The unit operates in the 117.000 to 135.975 MHz frequency range with a frequency spacing of 25 kHz that provides 720 channels. Optional transceivers have an extended frequency range of 116 to 151.975; this provides 1,440 distinct channels.Each complete VHF transceiver installation consists of a control head, a transceiver, and an antenna. The control head contains controls for frequency selection and display, volume and power control, and squelch activation and testing. The transceivers are in the nose avionics compartment and a blade-type VHF antenna is on the top and bottom of the fuselage.

Figure 5B-19: VHF Transceivers Control Head

Audio EmergencyIn the event of failure of the pilot’s audio system, access to audio is available through the copilot’s speaker. If both systems have failed, place the emergency/normal switch in the EMER position.

Emergency OperationWith emergency/normal switch in EMER position, the following rules apply: all audio sources (COMM 1, NAV 2, ADF, etc.) are connected directly

to the headphones to eliminate any specific audio source, turn volume control down on

that source (e.g., NAV 1). This rule does not apply to COMM 1 and COMM 2.

volume control on microphone selector switch has no function in the emergency mode

if emergency/normal switch fails, the audio systems may be placed into emergency mode by pulling the two Circuit Breakers (CBs) labeled PLT AUDIO and COPLT AUDIO located directly beneath the avionics master CB on the main CB panel.



HF CoeeuntSmatonSome aircraft use High Frequency (HF) communications equipment to allow very long range communications. Typical systems operate in the 2.0000 to 29.9999 MHz range with frequency spacing of 100 Hz; this provides 280,000 distinct channels. Most HF transceivers provide Amplitude Modulation (AM) and Single Side Band (SSB) transmission modes.Each installation consists of a transceiver, control head, power amplifier/ antenna coupler, and a long wire antenna.

Rmdto ThlhpconhA radio-telephone allows the crew or passengers to communicate with ground stations through the public telephone system, with mobile telephones, or other aircraft with radio telephones on the High Frequency (HF) and Ultra-High Frequency (UHF) radio frequencies. The system also allows communication between the cockpit and passenger cabin.Typical systems consist of a transceiver, antenna, cockpit unit, and passenger cabin unit. Depending on the system installed in the aircraft, the cockpit and cabin units consists of a handset with panel-mounted controls or an integrated system; the integrated system has controls in the handset. With both systems, the controls consist of channel selector buttons, a combined power switch and volume control, and intercom and transmit indicator/buttons.

Static DischargingStatic discharging wicks on the aircraft structure and control surfaces minimize the effects of lightning strikes and static charges on avionics equipment. The static dischargers bleed off accumulated static charges to the atmosphere. Due to varying configurations, consult your MEL for number and position of static wicks.

Figure 5B-20: Static Discharging Wicks

Avionics



NavigationDuring the preflight inspection, check the security, condition and presence of the dischargers. The minimum required for dispatch is one on each control surface; the outboard discharger on each aileron must be present.Navigation equipment can be loosely grouped into equipment that provides aircraft direction and attitude information, equipment that determines aircraft position, and equipment that provides flight management.

Attitude and DirectionAttitude and direction equipment use inertial and magnetic forces to sense and display aircraft heading and attitude. This includes: magnetic compass turn and slip indicator gyro horizon/vertical gyro radio magnetic indicator vertical gyro system compass system.

MmgnhatS CoepmssA conventional, liquid filled magnetic compass on the windshield center post provides aircraft heading information. The compass contains provisions for maintenance personnel to adjust the unit to compensate for aircraft generated magnetic fields. A correction card near the unit provides a record of recent adjustments to the compass and compass deviation errors.Heading information from the compass is only accurate in level unaccelerated flight.

Figure 5B-21: Magnetic Compass



Turn mnd ltp IndtSmaorsThe pilot and copilot have a turn and slip (turn and bank) indicator on their lower instrument -panels. Each indicator provides an indication of aircraft rate-of-turn and direction and assists in proper turn coordination. The pilot’s indicator has an electrically-driven gyroscope and the copilot’s indicator uses a vacuum-driven gyroscope. An inclinometer at the bottom of each instrument uses a curved, liquid-filled tube with a steel ball to indicate aircraft slip (yaw). On some installations the turn and slip indicator provides yaw data for yaw dampening.

Figure 5B-22: Turn and Slip Indicators

The pilot’s turn-and-bank indicator operates on 28 VDC. If the unit loses power, a GYRO warning flag appears on the instrument face.

Avionics



Gyro-Horizon/Vertical GyroOn aircraft with single flight director systems, the pilot has an Attitude Director Indicator (ADI) driven by the flight director system and the copilot has a vacuum driven gyro-horizon. Aircraft with dual flight directors have an electrically-driven standby (emergency) gyro horizon on the pilot’s instrument panel. The gyro-horizon indicates aircraft attitude with a fixed airplane symbol over a moving sphere. The sphere has a horizon line and degree marks above and below the horizon line to indicate pitch angles. A knob on the instrument face adjusts the aircraft symbol to correct for variations in level flight attitudes.

Figure 5B-23: Horizon/Vertical Gyro

One (single flight director system) or two vertical gyros (dual flight director systems) provide aircraft pitch and roll information to the autopilot, flight instruments, flight director, and radar antenna stabilization system.Each unit consists of an electrically driven gyro rotating on its vertical axis. Gimbals within the unit limit the amount of freedom in the pitch and roll axes. The gyro is free to pitch approximately 80° up and down, and roll 360° (roll unlimited). The vertical gyro(s) in the nose compartment electrically drives servoed instruments on the instrument panel.A FAST ERECT switch on the instrument panel for each vertical gyro manually erects the vertical gyro.

Figure 5B-24: FAST ERECT Switch



Compass SystemTwo compass systems (directional gyros) provide 360° of magnetic heading information to the pilot’s and co-pilot’s Horizontal Situation Indicators (HSIs), Radio Magnetic Indicators (RMIs), autopilot, and flight director. The pilot’s system drives the copilot’s RMI and the copilot’s system drives the pilot’s RMI. Each directional gyro installation consists of a directional gyro, a flux valve, remote compensator and control switches. The directional gyros in the nose compartment electrically drive the RMIs, flight director system, and the horizontal situation indicator(s).Each gyro consists of an electrically-driven gyro with monitoring and control circuits. A flux valve in each outboard wing or horizontal stabilizer senses the strength and direction of the earth’s magnetic field and converts it into electrical signals for gyro compensation. The compensating signal applied to the gyro aligns the gyro with magnetic north. The directional gyro sees this input as an error signal that it compares to a reference signal. The difference between the error signal and the reference signal produces a signal that drives a slave torquer motor in the gyro. The gyro then precesses to align itself with the flux valve and magnetic north.The remote compensator uses adjust-able permanent magnets to counteract the effect of magnetic fields generated by direct current and ferrous materials in the aircraft on the flux valves.A GYRO/SLAVE/FREE switch for each directional gyro allows the selection of either slaved or free gyro operation. In SLAVE, the directional gyro follows signals provided by the flux valve. In FREE, the directional gyro operates independently from the flux valve; manual correction of the gyro is through the INCREASE/ DECREASE switch. To manually align the compass system, use the INCREASE/DECREASE until the compass card aligns with magnetic heading.

Figure 5B-25: GYRO/SLAVE/FREE Switch

Avionics



Rmdto MmgnhatS IndtSmaorsTwo radio magnetic indicators display aircraft heading information on a calibrated servo driven compass card. A pointer and compass card provide bearing indication to either VOR or ADF stations.

Figure 5B-26: Radio Magnetic Indicators

Position DeterminingPosition determining equipment includes systems that operate independently of ground stations or with ground stations to determine aircraft position. This includes: Instrument Landing System (ILS) Very High Frequency (VHF) navigation equipment Automatic Direction Finding (ADF) Distance Measuring Equipment (DME) transponder long range navigation equipment LORAN global positioning system flight management system weather radar.

Insaruehna LmndtngInstrument Landing Systems (ILS) combine outputs from the VHF navigation, UHF glideslope and marker beacon receivers to display ILS information on the attitude director indicator and the horizontal situation indicator.The system consists of a glideslope receiver operating in the 329.15 to 335.00 MHz frequency range, the VHF receiver in LOC mode operating in the 108.10 to 111.95 MHz frequency range, a glideslope antenna in the nose, and a LOC antenna on each side of the vertical stabilizer.



VHF NavigationVHF navigation receivers provide Very High Frequency Omni-Range (VOR), localizer (LOC), Glideslope (GS), and marker beacon navigation information to the flight crew.Each VHF NAV system receives 200 VHF frequencies from 108.00 to 117.95 with 50 MHz spacing, 40 paired glideslope frequencies from 329.15 to 335.00 MHz spaced at 150 kHz, and 40 LOC frequencies from 108.10 to 111.95 MHz. Automatic DME channeling is through the navigation receiver. Multiple outputs from the receivers drive the flight director, Radio Magnetic Indicators (RMIs), autopilot, course deviation indicators, and area navigation equipment (RNAV). The receiver supplies audio output to the audio control units.Receiver control, frequency selection, and frequency display are through control heads on the center instrument panel.

Figure 5B-27: Receiver Control Heads

As part of the VHF navigation receiver, a marker beacon receiver provides visual and aural indications of beacon passage. The system receives on 75 MHz and provides electrical outputs to two sets of three indicating lights on the instrument panel. The receiver also provides audio output to the audio control units for beacon passage notification.

AuaoematS DtrhSaton FtndhrAutomatic Direction Finder (ADF) systems consist of a receiver, control head, and a combined loop and sensing antenna. The receiver operates in the 190.0 to 1749.5 kHz frequency range with 0.5 kHz spacing that provides 3,120 distinct frequencies.ADF systems provide three basic modes of operation: antenna (ANT), Automatic Direction Finding (ADF), and tone (TONE). In antenna mode, the radio magnetic indicator (RMI) pointer parks and the system provides only audio output. ADF mode provides continuous relative bearing readings to low frequency radio beacons, and AM broadcast stations and audio outputs to the audio control panels. Tone mode provides a 1,000 Hz tone for identification of station identifiers.

Avionics



DtsamnSh MhmsurtngDistance Measuring Equipment (DME) computes and provides slant range distance between the aircraft and a VORTAC facility. The system transmits in the 1,025 to 1,150 MHz range and receives in the 962 to 1,213 MHz range. Pairing of DME channels with VHF navigation frequencies allows the automatic selection of DME channels by the VHF receiver.The DME system provides distance, speed, and time information to the Horizontal Situation Indicators (HSIs), and DME displays.

Figure 5B-28: DME Displays

Area NavigationArea navigation systems (RNAV) allow point-to-point navigation within the coverage of VHF navigation facilities (VOR/VORTAC/DME). Most systems allow the storage of flight plans containing multiple way-points for frequently flown routes.These systems utilize data provided by the VHF, DME, and localizer receivers to compute and display waypoint information.

TrmnspondhrTypical transponder systems with Mode C or Mode S capability provide identification and altitude reporting to surveillance radar installations. The system consists of a transceiver, control head, and a transmit/receive antenna. The system transmits on 1,090 MHz and receives on 1,030 MHz. The pilot’s encoding altimeter provides aircraft altitude information to the transponder system for transmission to ATC radar facilities.

Oehgm/VLFVery Low Frequency (VLF), VLF/ Omega, and Omega/VLF navigation systems provide great circle, point-to-point navigation on a world-wide basis. These systems utilize very low frequency transmissions from Omega and U.S. Navy facilities.



Typical systems consist of a control display unit and a receiver-computer unit. The control display unit contains a display and keyboard for data entry and navigational data selection. The display presents system initialization, self-test, flight plan, and navigational information.The receiver-computer unit processes all inputs and provides position co-ordinates, distance and deviation information, drift and track angle deviation, wind direction and speed, and ground speed to the control display unit. The receiver-computer also provides inputs to the Horizontal Situation Indicators (HSIs), autopilot, and flight director system. Loss of navigation facility signals causes the system to revert to dead reckoning based on aircraft heading, true airspeed, and last computed winds.Eight Omega stations operated by the U.S. Coast Guard broadcast between 9 to 14 kHz; 10 VLF stations operated by the U.S. Navy broadcast between 14 to 24 kHz. The positioning of the Omega and VLF stations and the characteristics of low-frequency radio signals provides world-wide coverage.Atmospheric conditions, solar activity, height of the ionosphere, and variations in the earth’s magnetic field all affect the accuracy of the Omega/VLF system. Because very low frequency signals bounce between the ionosphere and the earth’s surface, any changes in these factors affect signal accuracy, reduce signal strength, and increase background noise to result in unreliable system operation.

LORANLORAN (long range navigation) systems use chains of low frequency transmitting stations to determine aircraft position with a relatively high degree of accuracy. The LORAN C chains consist of a master transmitter (M) and up to four slave transmitters (W, X, Y, and Z). Each chain transmits in the 90 to 110 kHz frequency range with the slaves transmitting in a set order after the master. The location and arrangement of LORAN C transmitter chains provides coverage of North America, the Middle East, Asia, and the Pacific Rim.Each aircraft system consists of a computer controller unit and an antenna. The controller contains a receiver, computer processor, and a interface (keyboard/display). After entering the aircraft’s present position and destination, the system receives signals from the master and slave LORAN C transmitters. The computer/controller processes these signals and determines by recognizing the time difference between the master and slave transmitter signals. The computer/controller processes this information to establish a hyperbolic line of position; this provides aircraft longitude and latitude with an accuracy of approximately 60 ft.

Global Positioning SystemGlobal Positioning Systems (GPS) receive transmissions from a constellation of 24 satellites (21 operational/ three spares) orbiting the earth in six orbital planes. The arrangement of these orbits and the satellites within them ensures that six to ten satellites are visible from any point on the earth.

Avionics



Using transmissions from four satellites received by a GPS sensor, the NCU can compute aircraft longitude, latitude, and altitude. With only three satellites visible, the system can compute only longitude and latitude. Even when limited functions are available, the system can determine aircraft position within 100 meters (civilian use restriction).

Flight Management System Flight Management Systems (FMS) utilize position information from various navigation equipment to provide an integrated navigation display and control system.The FMS receives inputs from the VHF navigation equipment, computes the aircraft position, and provides outputs for the autopilot, flight director, RMIs, and Horizontal Situation Indicators (HSIs).

Whmachr RmdmrWeather radar systems consist of an antenna, receiver-transmitter, and radar display with system controls. The vertical gyro system provides aircraft attitude information to the radar system to stabilize the antenna. The system operates by transmitting a high frequency radio signal, receiving the bounced signal, and displaying the received signals on the display. Controls on and below the indicator select system mode, scan range, antenna tilt, and receiver gain (sensitivity). Typical systems provide: selectable scanning range ground mapping weather cell contouring adjustable antenna tilt and scan target alerting navigation information overlay (some systems).

Figure 5B-29: Weather Radar Systems



Radar power output and scanning area varies between equipment manufacturer, model, and radar capabilities. Hazard areas presented in this discussion come from various aircraft and component maintenance manuals and FAA Advisory Circular AC 2068B. Personnel hazard areas are the maximum recommended hazard area for radar operation on the ground.When operating radar on the ground, precautions should be taken to avoid injury to personnel, fuel ignition, or radar equipment damage. Avoid operating the radar during refueling or within 300 ft of refueling aircraft. Caution personnel to remain outside an area within 270° and 15 ft forward of the radome. Direct the nose of the aircraft so a 240° sector forward of the aircraft is free for a distance of 100 ft of large obstructions, hangars, and other buildings. Tilt the antenna up to its maximum angle.Automatic flight control systems combine an autopilot, flight director, air data system, controls, indicators, and displays to provide automatic control of high performance aircraft.Please refer to the applicable pilot’s guides, Aircraft Flight Manual Supple-ments and component maintenance manuals for a thorough dis-cussion of these systems and their operating procedures.Flight control system provide three modes of operation: manual, automatic, and semi-automatic. Manual operation allows the crew to fly the aircraft guided by cues provided by the flight director system. Automatic operation flies the aircraft through the autopilot coupled with the flight director; the crew monitors system operation. Semi-automatic operation flies the aircraft through crew interaction with the autopilot controller and controls on the control wheels.

Avionics



Flight Control System

AutopilotThe autopilot system provides automatic control and stabilization of the aircraft about the pitch, roll, and yaw axes. It positions the aircraft elevator, ailerons, and rudder in response to autopilot/flight computer steering commands. Selectable operating modes provide the ability to maintain automatically a desired altitude, pitch attitude or heading, and to capture automatically and track localizer, glideslope, and VOR signals. Systems certified on the Super King Air include: Collins AP-105 Collins AP-106 Collins APS-65 Collins APS-80 King KFC-300 King KFC-400 (B200 only).

A typical autopilot system on the Super King Air consists of: autopilot or flight control computer autopilot controller airspeed sensor or Air Data Computer (ADC) mode selector aileron, elevator, and rudder servo-actuators.

The autopilot system receives signals from the airspeed sensor or ADC, vertical accelerometer, vertical and directional gyros, and navigation receivers. With this data, the autopilot drives the servo-actuators to maintain a desired altitude, attitude, navigation path, or air-speed. A typical autopilot system provides: yaw damping roll rate and bank angle limiting automatic capture and track of VOR, ILS, and localizer heading, roll, airspeed, and altitude capture and hold heading select soft ride.



Auaoptloa ConarollhrAutopilot control and operation is through an autopilot controller or autopilot engage and manual controllers on the pedestal. Controls common to most autopilot systems include: autopilot and yaw damper engage levers or buttons pitch wheel turn knob or vertical trim switch soft ride (turbulence) mode button.

Figure 5B-30: Autopilot Controller

The autopilot engage lever or button works with the yaw damper control to engage the autopilot and yaw damper simultaneously. The autopilot does not engage without the yaw damper. Disengaging the autopilot does not disengage the yaw damper; the yaw damper can be engaged without engaging the autopilot.Through the turn knob and pitch wheel the pilot can manually “dialin” a desired roll or pitch angle. Each control initiates an aircraft attitude change in proportion and in the direction of control movement. With a lateral mode selected, moving the turn knob out of the center detent (wings level) disables the selected lateral mode. Depending on the system, moving the pitch wheel or vertical trim switch from the center detent disconnects the vertical mode selected (i.e., altitude hold, indicated airspeed, or vertical speed) or initiates a pitch change in the direction of control movement.The soft ride or turbulence button commands the system to reduce autopilot gains. Rather than react to every disturbance of aircraft attitude due to flight in turbulent air, the system reduces the control responses. This increases passenger comfort by reducing the continuous control movements the autopilot would command in-turbulence.

Avionics



Additional controls include an auto-pilot/yaw damper disconnect and Touch Control Steering (TCS) or Control Wheel Steering (CWS) button on the control wheel and a Go-Around (GA) button on the left throttle lever. Touch control steering allows manual control inputs without disconnecting the autopilot. After making attitude inputs, releasing the TCS button reengages the autopilot; the autopilot maintains the manually input aircraft attitude.

Figure 5B-31: Additional Autopilot Controller

Auaoptloa CoepuahrThe autopilot computer receives pitch, roll, and yaw signals from the navigation sensors, gyros, autopilot controller, and the flight director computer. The computer takes these signals, performs computations, and sends driving signals to the aileron, elevator, and rudder servos.Pitch axis signals come from flight director computer, airspeed sensor, autopilot controller, and vertical gyro. Roll axis signals come from the vertical gyro, directional gyro, flight director computer, and autopilot controller. Yaw axis signals come from a directional gyro that provides heading information.Altitude information comes from the Air Data Computer (ADC). The auto-pilot uses the altitude information to enable or disable torque switching. Torque switching provides two distinct rates of servo torque depending on aircraft altitude.The computer takes these signals and compares the aircraft’s actual attitude to the desired attitude. The computer then commands the servos that move the flight controls to reposition the aircraft to match the desired aircraft attitude.



ServosServos for the ailerons, elevator, and rudder consist of a DC motor-tachometer, clutch assembly, synchro, and power gear train. Signals from the autopilot computer drive the servo motor through cables to position the control surface. A feedback signal produced by the motor-tachometer relays control surface position information to the computer. Once the control surface reaches the commanded position, the computer signals the motor to stop. During autopilot engagement, the electromagnetic clutch assembly connects the servo motor to its output shaft. With the autopilot disengaged, the output shaft moves freely.The crew can overpower the auto-pilot through control wheel movement. If a mechanical failure occurs, the servo’s clutch slips to allow normal control.

Flight DirectorThe flight director system generates vertical and lateral steering commands for the Attitude Director Indicator (ADI) and the autopilot system. The ADI displays these commands as command bars; the autopilot uses them as steering commands. A typical flight director system consists of: flight guidance computer mode selector mode coupler attitude director indicator horizontal situation indicator.

Fltgca CoepuahrThe flight guidance computer provides pitch and roll commands to the flight control system and the autopilot. The computer uses signals supplied by the navigation receivers, air data computer (airspeed sensor), and vertical gyros to generate the pitch and roll commands. The computer also supplies warnings to the flight crew about the status of the flight control system.The mode coupler passes the pitch and roll commands to the ADI. The auto-pilot computer receives and passes the same commands to the servos.

Modh CouplhrThe mode coupler provides an interface between the units of the flight control system. It also filters, regulates, and protects the flight control system from DC power transients.

Avionics



Modh hlhSaorThe mode selector or flight guidance panel allows the selection of operating modes for the flight guidance computer.

Figure 5B-32: Mode Selector or Flight Guidance Panel

Separate vertical and lateral modes are selectable on the flight guidance panel. Vertical modes include altitude pre-select, altitude hold, vertical speed hold, indicated airspeed hold, and go-around. Lateral modes include heading select, NAV-LOC, and approach.

Aaataudh DtrhSaor IndtSmaorThe Attitude Director Indicator (ADI) provides a three-dimensional display of aircraft attitude and flight control system commands. The attitude director indicator displays: pitch and roll commands flight director steering commands localizer and glideslope deviation rate-of-turn radio altitude decision height speed deviation.

Figure 5B-33: Attitude Director Indicator



The relationship of a stationary aircraft symbol with respect to a moving horizon line represents aircraft attitude. The horizon line is servo-driven through the pitch and roll axes. The horizon line is a solid line with blue sky above and brown earth below. Degree bars on the sky and earth portions of the display indicate aircraft pitch angles in climb and descent. A scale at the top of the indicator displays aircraft roll angle.Two bars flanking the aircraft symbol or a “V” bar display steering commands from the flight director. The bars are servo-driven for combined pitch and roll commands. Numerous warning flags within the indicator alert the crew to invalid information received by the ADI.

Horizontal Situation IndicatorThe horizontal situation indicator displays: aircraft position and heading selected heading and selected course distance to or from a DME station or waypoint deviation from selected VOR, localizer (LOC), or other navigation

aid vertical deviation from glideslope, and TO/FROM and bearing/track

pointer information elapsed time or time-to-go.

Figure 5B-34: Horizontal Situation Indicator

An aircraft symbol on the Horizontal Situation Indicator (HSI) shows airplane position and heading in relation to an azimuth card, lateral deviation bar, and selected heading. The azimuth card displays heading information from a gyro-stabilized magnetic compass. Heading is read on the card beneath the lubber line at the top center of the indicator.The Heading (HDG) knob allows the crew to set a marker to a desired heading as read on the azimuth card. This allows the crew to set a heading on the azimuth card for display as a steering command on the ADI. A course knob rotates the course arrow on the indicator to a bearing as read on the azimuth card.Numerous warning flags appear to alert the crew to invalid data and system and component failures.

Avionics



Preflight and Procedures

PreflightDuring the exterior preflight inspection, check the condition of the pitot masts and static ports. All should be clean and free from obstructions. Check the security and condition of the antennas on the underside of the fuselage. Visually inspect the antennas on the fuselage top and vertical stabilizer. Check the static wicks for security and condition.Check that the ELT is armed and that its antenna is secure (if visible). If suspected, drain moisture from the static lines with the static system drain valves on the right side of the cockpit. Preflight procedures for the autopilot vary with the system installed. Always refer to the applicable AFM Supplement. Generally, turn the battery and avionics master switches on. Engage the autopilot and yaw damper. Check that control wheel and rudder pedal movement overpowers the autopilot; there should be resistance to control movement. Check that the control wheel disconnect switch disconnects the autopilot and yaw damper.

Abnormal and Emergency Procedures

Emergency ProceduresEmergency procedures for the pitot/static system and autopilot are limited. They include: pilot’s static air system autopilot malfunction or failure.

Pilot’s Static Air SystemIf the pilot’s and copilot’s airspeed indicator, vertical speed indicator, and/or altimeter disagree or the pilot’s vertical speed indicator appears sluggish, there may be a partial or complete obstruction of the pilot’s static ports. Move the PILOT’S STATIC AIR SOURCE switch to ALTERNATE and observe instrument indications. If this procedure corrects the problem, the pilot’s primary static ports are obstructed.With ALTERNATE selected, the pilot’s airspeed indicator and altimeter are in error because of the location of the pilot’s alternate static port. Use the Airspeed Calibration and Altimeter Correction charts in the AFM Performance Section to correct the indications.



Autopilot MalfunctionNormally, if the autopilot malfunctions or fails, the autopilot automatically disconnects. If the autopilot fails to disconnect, disconnect the autopilot by: pressing the control wheel disconnect switch pressing the go-around button pressing the TEST button on the controller pulling the autopilot circuit breaker turning off the avionics master switch.

Typically, most autopilot systems disconnect if there is a power interruption or failure.If AP TRIM and MASTER WARNING annunciators illuminate with the autopilot engaged, immediately disconnect the autopilot while restraining the control wheel; there may be a severe nose down pitching movement. During autopilot malfunction, there may be an altitude loss before regaining positive control of the aircraft.Refer to Table 5B-1, for a partial list of demonstrated altitude losses observed during autopilot malfunction.

’Configuration Collins Collins Collins King Sperry SperryClimb 200 450 450 -40 250 130Cruise 400 440 440 170 410 310

Maneuvering 120 150 150 50 100 140Descent 430 450 450 420 500 510

Approach ILS 80 150 100 50 80 90Single Engine 90 — — 80 50 80

Table 5B-1: Altitude Loss During Autopilot Malfunction Testing

Avionics 5B

King Air 200 (EFIS) 5B-1December 2011


ContentsAvionicsAvionics Equipment

Avionics Power ......................................................................................5B-11Flight Environment Data

SchematSc: Pitot/Static System (Non-ADC Equipped) ...............5B-14SchematSc: Pitot/Static System (ADC Equipped) .......................5B-15

Pitot/Static System ................................................................................5B-16Pitot Heads ........................................................................................5B-17Static Sources ....................................................................................5B-17Outside Air Temperature ....................................................................5B-18Air Data Computers (ADC) ................................................................5B-19Altimeter ALI-80A ...............................................................................5B-20Airspeed Indicator ..............................................................................5B-21Vertical Speed Indicator .....................................................................5B-21Controls .............................................................................................5B-26Altitude Preselector/ Alerter ...............................................................5B-27Altitude Alert Light/ PUSH TO CANCEL/Test Switch .........................5B-27Radio Altimeter ..................................................................................5B-28Emergency LocatorTransmitter ..........................................................5B-28Turn and Slip Indicators .....................................................................5B-29Cockpit Voice Recorder .....................................................................5B-30

Communication SystemsAudio Integrating ...................................................................................5B-31

Microphones/Headsets ......................................................................5B-31Audio Control Panels .........................................................................5B-32

Speech Communication ........................................................................5B-32VHF Communications ........................................................................5B-33Audio Emergency Operation Switch ..................................................5B-33Emergency Operation ........................................................................5B-33HF Communication ............................................................................5B-34Radio Telephone ................................................................................5B-34

Static Discharging .................................................................................5B-34

King Air 200 (EFIS)5B-2December 2011


NavigationAttitude and Direction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-35

Magnetic Compass ...........................................................................5B-35Turn and Slip Indicators ....................................................................5B-36Gyro-Horizon/Vertical Gyro ...............................................................5B-37Compass System ..............................................................................5B-38Radio Magnetic Indicators ................................................................5B-39

Position Determining . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-39Instrument Landing System ..............................................................5B-39VHF Navigation .................................................................................5B-40Automatic Direction Finding ..............................................................5B-40Distance Measuring ..........................................................................5B-41Area Navigation ................................................................................5B-41Transponder ......................................................................................5B-41Omega/VLF ......................................................................................5B-41LORAN .............................................................................................5B-42Global Positioning System ................................................................5B-42Flight Management System .............................................................5B-43Weather Radar .................................................................................5B-43

Flight Control SystemFlight Management System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-45

Universal UNS-1K .............................................................................5B-45Turbulence Weather Radar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-46Flight DirectorSystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-47

Flight Director System Flight Guidance Computer (FGC) ................5B-47Flight Control Panel ..........................................................................5B-48Lateral Modes ...................................................................................5B-51Vertical Modes ..................................................................................5B-52Autopilot ............................................................................................5B-53Autopilot Controller ...........................................................................5B-53Autopilot Computer ...........................................................................5B-56Servos ...............................................................................................5B-56

EFIS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-56Overview ...........................................................................................5B-56Description ........................................................................................5B-56Display Select Panel .........................................................................5B-56Display Processor Unit .....................................................................5B-57Multifunction Processor Unit .............................................................5B-58EADI .................................................................................................5B-58

Avionics



EHSI . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-59Multifunction Display .........................................................................5B-61

TCAS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-62Description ........................................................................................5B-62Advisory Type ...................................................................................5B-63ATC Control ......................................................................................5B-63

Preflight and ProceduresAvionics Related Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-67

Autopilot Annunciators ......................................................................5B-67EFIS Annunciators ............................................................................5B-67Global Positioning System ................................................................5B-67

Preflight and Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5B-67Preflight .............................................................................................5B-67

Abnormal and Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . 5B-68Emergency Procedures ....................................................................5B-68Pilot’s Static Air System ....................................................................5B-68Autopilot Malfunction ........................................................................5B-68



Tcts pmgh tnahnatonmlly lhfa blmnk.

Avionics



AvionicsThis chapter provides a brief overview of the following: pitot/static system and instruments communication equipment navigation equipment flight control systems.

Besides these specific areas, this chapter includes instrumentation not addressed in other chapters.Cockpit panel art at the front of the chapter provides a ready reference to locate instruments and equipment addressed in this discussion.Several different EFIS installations are available. If your aircraft is so equipped, see appropriate supplement to AFM.For a detailed description and operating procedures for a -particular piece of avionics equipment, refer to the applicable pilot’s guide.




King Air 200 (EFIS)December 2011

5B-7For Training Purposes Only

Avionics

Cockpit Instrument Panel

- - -

TO TEST

PR

18 12

27

33

30

2421

3

0

6 9

16

18 12

27

33

30

2421

3

0

6 9

16

MASTERWARNING

MASTERCAUTION

TO RESET

33

24

21

30

15

12 6

3

DIM

VOL

F A S T

E R E C T

33

3024

21 15

126

3

GS

HDGNAV

PRESS

TO RESET

PRESS

MASTERWARNING

MASTERCAUTION

TO RESET

PRESS

TO RESET

PRESS

2

0

1 2

0

1

8050

70 8050

70

DOORL ENG FIRE INVERTER ALT WARN R ENG FIREUNLOCKEDL FUEL R FUEL

ESSPRESS PRESS

EXTINGUISHEREXTINGUISHERPUSHPUSH L BL AIR R BL AIR

FAIL FAILDISCHARGEDDISCHARGED

START START AUDIO COMM NA MV KR BCN DME COMM NA MV KR BC DN ME AUDIO 9 12 9 12 COMM 1 2 1 2 1 2 21 ADF 21 1 2 1 2 1 A2 DF COMM

EFIS AUX

DRIVEITT ITT MFD MPU EFD DPU COMPARE PUSH TOFAN FANOFF8 2 8 2 OFF RESETAUX FAIL FAILTERR GPWS GPWS PILOT AUDIO OFF PUSH COPILOT AUDIO OFF GPSDRIVEGPWS GPWS GPSINOP TERR INOP FLAP INTEGON/OFF HOT AUDIOP/TEST 4 4 AUDIO AUDIOOVRD G/S INTEG °C x 100 °C x 100 PAGING INTPHTERR G/S SPKR7 7 EMER VOICE VOICE INTPH-PRESS- BSPKR B-PRESS- P/CANCELCANCLDP/OVRD 5 56 6 O GND OCOMM 2 COMM 2T TCOMM VOLVOLCOMM 1 CABIN HAFX AFX OFF NORM RANGE PWR RANGE H OFF OFF COMM 1 CABIN

MKR BCNCollins DME 1 2

1 & 2 VOLALTITUDE SET HI 300 40 040 0300 PUSH TOVOL VOL160 ALTS VS 550 HDG LOC1 9 60 9 1TORQUE TORQUE0 0

ALT 30 , 00 0VOL

ALERT 26 26280 280KNOTS KNOTS2 2GS ALT0 0 700 LO80 8024 244 4AIR SPEED AIR SPEED2AP/L 8 ENCD 8 222 22 CANCEL OANN SIDE-SIDE- INTPHINTPH260 260 VG FAST6 6 ALTM FPUSHFTLB x 100YD FTLB x 10020 20 TONETONE SENS FSENS ERECT8 820 1100 100 PUSH BRT 100 0 718 1810 1016 12 16 1214 14 ON240 7 2403 1000 FT

100 FT7 32120 120 IN HG10 2 9 . 9 2 10 10 2992220220

140 COMM 1 COMM 2 140200 200 6 4Collins5 10TEST 10Collins Collins160 1600 0 520180 180 20BARO5 5 PWR INT A2 P3 ROP 2 P3 ROP A XFR XFRL R CC10 22 22 T USTB GCS TRU 1030

T

RDR G+2 10 1021 21

MEM MEMTFC20 2020 RPM x 100 13 RPM x 100 13 MEM MEM0 1 19 4 1 19 418 180

1 17 16 5 1 17 16 5 SQ COM RDR SQ COMDH 200 1 2 ON OFF 100 ON OFF 1 2NAVSTO OFF STOD C OFF .5 4 .5 4V V

RMTTEST TEST0 ACT ACTVERT SPEED 0 L R1010 VERT SPEED 2020Collins 0 6 0 6A A A APGENAV 1 NAV 2D D D D

F3030GYRO F SLAVING

F FCollins A-2.5 Collinsx1000 FPM 0x1000 FPM100 40 100 40 WX+T 50MAG1DIST MSG LIN .5 4 EMG040 ATGT XFR XFR NNN 4A .590 PERCENT

RPMPERCENT C18.4H 90 CRPM T D CT1 2 60CID 60 1 2N M RSLEW MODE NAV NAV

DG NAV NAV M RMEM MEM

DATAMEM MEM V O

W

E

NAV NAVO R ON HLD ON HLD DHRCL SKP CLROFF STO OFF STO6 61 5 5 DH

GYROV VFUEL FLOW FUEL FLOWW

E

1 2 3 WPT NM HLD KT MIN ID SLAVING4 4V TTEESSTT TTEESSTTEFIS 1ACT ACT S

OAUX POWER 1 2 3 WPT NM HLD KT MIN ID

3CH SEL PWR R X1 00 FT3ON HORN 22 TRANSPONDER ADFH PPH X 100 PPH X 100Collins SLEW MODE CH SEL PWRCollins Collins

COLLINS20 DG

3

S

BRG A XFRA 1 TEST SILENCE ADF 1 CRS C 15 RAD ALTC

T090 T 410 5OI OL IL 2 MEM

140 20 10 40 200AADDFF TEST

100 15 10 00 150 ATCON ALT ADF TONE STBY IDENT NORMAL NORMAL ANT STO OFF

60 10 60 0 100 CMPST MFD DRIVE HDG OFF 20 5 20 0 50MIC INVERTE ER NG AUTO

NORMAL NO I1 GNITION 00 PREAVIONICS PROP SYNC 00 TTEESSTT TTEESSTT ACTMASTER POWER -2 -0 20ONARM °C PS °I C PSI

OOFF MICF

NORMALF 10OFF 4 5OXYGEN NO 2OFF LEFT RIGHT 63 PSIMASK

LIGHTS ENVIROMENTAL OFF 0 20GE EN NGINE ANTI-ICE LIGHTS CABINGYRORESET LANDING TAXI IC NE AV RECOG MANUA VL ENT OXYGE PN NUEMATICLEFT RIGHT BEACO SN TROBE FC UR NN O SMOK ME AN AUTO SUCTIONON GEN - UP TO RESET

TEM BP LOWER INCR PRESSUREMASTER SWITCH BRIGHTOO N MASK& FSB INCHES OF MERCURYHEATDOWN - OF OF ON N INC HR IGHLDG GEAR CONTROL F OF F LO CABIN/OFF OFF E F

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STARTER ONLY TEST MANUA LL EFT RIGHT RELAY SILENC TE EST OF OF FF OFFOFFPARKING BRAKE OFF

20000


5B-8 For Training Purposes Only



5B-9For Training Purposes Only

Avionics

Pedestal

MAP

PAGE ADV

ATC IDENT

MAP

PAGE ADVATC IDENT

5

3 1

3

5

3 1

5

1 3

1

5

TI

RM

NOSE UP

NOSE DN

HT

CI

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NOSE DN

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PRESSTO

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L DC GEN

L ENG ICEFAIL

L ENGANTI-ICE

L IGNITIONON

L CHIPDETECT

L BL AIROFF

BATTERYCHG

LDG/TAXILIGHT

HYD FLUIDLOW RVS NOT READY

DUCT OVERTEMP

EXT PWR

AIR CONDN1 LOW

PASS OXYON

FUELCROSSFEED

R BL AIROFF

R DC GEN

R IGNITIONON

R CHIPDETECT

R ENG ICEFAIL

R ENGANTI-ICE

UP 020 2

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QUARTZ CHRONOMETER DN ONLY WITH ENGINES 20 4

5 RUNNING REVERSE FLAP 18 6

FRICTION LOCK UP 16 8 BRT

ST-SPMODERST

LC ETZU FT DCAILERON TAB

LEFT RIGHT FLAP 1214 10

U S A

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SET ADVAPPROACH

TDIM

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O2 ARM MAN OVRD

PULL ON PASSENGER SYSTEM READY MANUAL DROP OUTOXYGEN

HDG CRS UNIVERSAL

HDG CRS NAV CRS

FPL MENU 1/2DTA CTL

ET MSG NAVNORMAL STORE FPL Collins ENGINES AVIONICS

DATA DTO HSI

CLEARANCE XFILL FPL DH SET ARC MAP SEL/RNG PROP LEFT FIRE LEFT LEFT LEFT LEFT LEFT LEFT AVIONICS

FUEL PPOS TO WPT RAIM PRED FPL CRSRDR ACT 5 5 5 5 71/2 5 5 2 5 5

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ON FLOW ON PREV DELETE FPL RETURN LIST SYNC

CHIP VANE CONTROL OIL OIL TORQUE FUEL CRS XFRRA CRS PRE BRG DET CONTROL HEAT TEMP PRESS METER FLOWDIMTST AUTO AVIONICS PILOT CABIN COPILOTNEXT A B C D E F G 1 2 3 MENU

PWR 5 5 5 71/2 5 5 2 5 30 15 71/2 2 OFF OFF

Collins

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LIGHTS WARNINGS WEATHERO P Q R S T U 6 7 8 AUX TRANSFER AUX TRANSFER

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8 108 10 5 71/2 5 5 5 5 5 71/2 5 5 30 71/2 2 21212 14 EFIS POWER EFISSEE MANUAL FOR 6

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AUTO 2 AUTO AVIONICS NO SMK AIRCollins PRCSRYAW DIS AP DIS T AP E RA 1/2Ø SR TRIM TRIM GEAR VENT2 18 18 READING & ENG OVHD & & FSB WARN WSHLD AVIONICS COMM NAV XPNDR

ADC0 L R0 20 FUEL QUANTITY 20 TEST 5 5 5 5 5 5 10 30 2 2 71/25 5QTY MAIN QTY DNSR LBS X 100 LBS X 100 OFF OFF INSTR CONSOLE CABIN RIGHT IND IND RIGHT WIPER NO. 2 NO. 2 NO. 2 NO.2YAW AP ENVIRONMENTAL FLIGHT ELECTRICALENGRIGHTLEFT ENG 1/2Ø

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5 5 5 5 5 5 3 1 5 10 50 50 3 2 10

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OPEN FIRE STANDYBY AUX QTY PRESS CROSS PRESS QTY AUX STANDYB FIRE OPEN 5 5 3 1 5 5 10 50 50 3 2 5WALL PUMP TRANS IND WARN FEED WARN IND TRANS YPUMP WALL

FIREWALL VALVE FER FER VALVE FIREWALL CONTROL RIGHT COPLT ENCD ALTM BOOST NO .2 RIGHT NO. 2 NO. 1 NO. 1 NO. 2CABINSHUT OFF SHUT OFF RUDDER DUMP PRESS FURNISHINGVALVE 5 10 5 5 5 5 5 5 5 10 5 VALVE BOOST

P TEST ERASE MASTER CIGAR AP FCS EHSI HDG CLOSED CLOSED R

LEFT RIGHT ELEVFUEL SYSTEM E 10 5 2 2 2 5TRIMS

TEST OFF FLAP ENGINE INSTRUMENTS

S HEADSET POWER LIGHTER SERVO POWER PRCSR

NO 3 LEFT LEFT LEFT NO 3 LEFT COCKPIT VOICE RECORDER

5 RADAR ADF RADIO50 25 20 5 5 5 50 5 5 5 5 5 5 UNIVERSAL

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PERFORM-I A ANCE SECTION FOR

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OFF

TEST


5B-10 For Training Purposes Only


Avionics

5B-11December 2011

For Training Purposes OnlyKing Air 200 (EFIS)

Avionics EquipmentAvionics equipment includes those subjects covered by Air Transport Association (ATA) chapters 22 – Auto Flight, 23 – Communications, and 34 – Navigation.

Avionics PowerThe majority of the avionics equip-ment on the Beech Super King Air operates on 28 VDC supplied by the aircraft battery and the generators. Three avionics buses (Avionics No.1, Avionics No.2, and Avionics No. 3) distribute power from these sources through individual circuit breakers to the various equipment. A single AVIONICS MASTER PWR switch on the pilot’s instrument panel controls relays that supply power from L/R Generator buses. A 5A Circuit Breaker (CB), labelled AVIONICS MASTER sup-plies power from the No.1 Dual-Fed bus to the switch. The switch, in turn, controls the operation of three avionics power relays.

Figure 5B-1: AVIONICS MASTER PWR Switch

With the battery switch ON power flows to the three avionics buses and passes through the normally closed avionics relay. This power subse-quently opens the relay to remove power from the avionics buses. Moving the AVIONICS MASTER switch to ON removes power from the relay; the relay closes and the avionics buses receive power.If the AVIONICS MASTER switch fails, pulling the AVIONICS MAS-TER CB restores power to the avion-ics buses.

5B-12December 2011

For Training Purposes Only King Air 200 (EFIS)

The Ground Communication Electric Power Bus is offered as an option. This feature reduces battery requirements when obtaining an IFR clearance prior to engine start. Pressing the on the center instrument panel (typical location) supplies power through relays from the Hot Battery bus to the No. 2 COMM and the No. 2 audio panel.

Figure 5B-2: GND COMM Button

Another option for the aircraft is an Auxiliary DC bus system. This system provides essential power to some avionics equipment during electrical bus failure or when electrical load-shedding is a required action. After turning the generator and battery switches OFF, turning the AUX DC BUS switch ON supplies power directly from the Hot Battery bus through relays to the No. 2 COMM, No. 2 compass system, No. 2 audio panel, the ILS portion of the No. 2 NAV, and the glareshield floodlights.

Figure 5B-3: AUX DC BUS Switch

Avionics

5B-13December 2011


Flight Environment DataFlight environmental data equipment senses environmental conditions to determine aircraft altitude and air speed. This includes the pitot/static system and air data computer (if installed). The pitot/static system and air data computer, in turn, drive: Altimeters Altitude Alerter Airspeed Indicators Vertical Speed Indicators.

Although not part of the pitot/static system, the radio altimeter and flight profile advisory system (if installed) are part of the flight environmental data system. The radio altimeter determines altitude and the flight pro file advisory system uses altitude information from the radio altimeter to provide verbal warnings of aircraft altitude and unsafe flight conditions.In addition, this section discusses the outside (free) air temperature gauge.

5B-14December 2011


Pitot/Static System (Non-ADC Equipped)

Avionics

5B-15December 2011


Pitot/Static System (ADC Equipped)

5B-16December 2011


Pitot/Static SystemThe pitot/static systems provide ram air pressure and static pressure data from their respective pitot mast and static ports. The pitot masts have electrical heating elements that provide anti-icing; the static ports are unheated. The LEFT and RIGHT PITOT switches in the ICE PROTECTION switch quadrant control 28 VDC power to the pitot mast heating elements. Do not use pitot heat for extended periods on the ground as it may damage the heating elements.To prevent yaw error, the pilot’s and copilot’s static systems have a port on each side of the fuselage aft of the aft pressure bulkhead. Each pair of ports connects to a single line that carries static pressure to the front of the air craft. The pilot’s static system has a selectable alternate static source on the lower part of the copilot’s CB panel. The pitot system supplies ram air pres sure to the pilot’s and copilot’s Airspeed Indicators (ASI). A pressure differen tial switch connected to the pitot sys tem provides input to the overspeed warning system.

Figure 5B-6: LEFT and RIGHT PITOT Switches

Figure 5B-4: Pitot Tube Figure 5B-5: Static Ports

Avionics

5B-17December 2011


Ptaoa HhmdsThe LH and RH pitot heads are heated to prevent ice from forming on them, resulting in erratic readings. The left pitot feeds the pilot’s Airspeed Indica tor and Air Data Computer and the right pitot feeds the copilot’s Airspeed Indicator and speed-warning switch. The pitot head is used as a drainage point for the pitot system.

amatS ourShsThe static system supplies ambient air pressure to the airspeed indicators, Vertical Speed Indicators (VSI), altime ters, and the cabin differential pressure gauge. If the pilot’s static system ports clog or a leak develops in the lines, moving the PILOT’S STATIC AIR SOURCE valve handle aft to ALTERNATE selects the static port in the aircraft tailcone aft of the pressure bulkhead as the pilot’s static source. Because the static port is in the aircraft fuselage, the pilot’s altime ter and airspeed indicator require cor rection to obtain accurate indications. Refer to the applicable charts in the Aircraft Flight Manual Performance Section.

Figure 5B-7: PILOT’S STATIC AIR SOURCE Valve Handle

Two drains for the pilot’s, pilot’s alter nate, and copilot’s static system are behind an access door on the lower right cockpit wall. After the aircraft has been exposed to rain or high humidity, drain the static lines. The pitot system lines do not require drains as they are the lowest point in the pitot system; moisture drains out through the masts.

5B-18December 2011


Ouastdh Atr ThephrmaurhAn Outside Air Temperature (OAT) gauge, on the overhead panel (early aircraft) or on the left hand sidewall panel near the pilot’s seat consists of a temperature probe that protrudes through the aircraft skin and a gauge calibrated in degrees Fahr enheit and Centigrade. Post lights illu minate the gauge. The digital outside air temperature system installed on BB-1439, 1444 and subsequent, BT-39 and sub-sequent, BL-139 and subsequent and BN-5 and subsequent con sist of a probe remote processor mounted under the copilot’s floor and a digital display indicator mounted on the pilot’s left side panel.

Figure 5B-8: OAT gauge Figure 5B-9: Digital Display Indicator

Avionics

5B-19December 2011


Atr Dmam Coepuahrs (ADC)The ADC-85 Air Data Computer computes air data parameters for all displays, flight con trol systems and navigation systems outputs. The air data computer receives its primary inputs from the pilot’s pitot and static sources. These inputs are converted to electrical sig nals and are used for computation. Other inputs are from the outside air temperature, the remote test switch, and a safety switch that enables the self-test only on the ground. The ADC has an air data module that is pro grammed with the air data parameters applicable to the aircraft type and con figuration.

Figure 5B-10: ADC

Computed results are converted into digital format and transmitted to the cockpit displays (altimeter, TCAS/VSI, airspeed indicator, EFIS, flight control and navigation system).

Test Switch Press and hold the ADC TEST switch on the pedestal to initiate the

computer self-test. While holding the ADC TEST switch, the following should occur:

the amber fault light on the ADC will illuminate; ADC TEST will be displayed on the EADI the fault light will extinguish within 1/2 second and the valid light

illu minates if the system is okay; ADC TEST will extinguish; Altimeter flag will be displayed and indication will go to the clos est

250 ft mark.Release the ADC TEST switch and the following should occur: all lamps illuminated during test should extinguish; altimeter indication should be restored and flag should go out of view.

5B-20December 2011


Alatehahr ALI-80AThe altimeter receives digital data from the air data computer to display barometric altitude using a 5-digit numerical readout and a pointer. The altimeter supplies baro-corrected data to the altitude preselector/alerter. Altitude information is also sent to the transponder for the Mode C altitude reporting.

Figure 5B-11: Altimeter

The numerical display indicates altitude from –1,000 to +55,000 ft; the drum counter is set in 100 ft increments.The pointer indicates altitude from 0 to 1,000 ft in 20 ft increments. The barometric setting is displayed below the center of the display and can be set using the BARO knob. Pulling on the BARO knob will switch the display from Hg to MB mode. The correction range is 22.00 to 32.00 InHg.In the event of lost data from the ADC, a warning flag will come into view and the baro pointer will go to the 250 increment.Pressing and holding the TEST button on the altimeter will exercise the self-test of the altimeter. While in self-test mode, the flag will be displayed for 1/ 2 second, the pointer will go to 750, and the transponder output will be inhibited. Releasing the TEST switch will return the indicator to normal operation.The pilot’s altimeter provides encoded altitude information as an electrical signal to the altitude alerter and to the transponder for Mode C/S altitude reporting. If the encoder loses power or fails, an OFF flag appears on the instrument face.The copilot’s altimeter has a 28 VDC powered internal vibrator that overcomes instrument friction and ensures instrument accuracy; the vibrator prevents the pointer(s) from sticking.The pilot’s altimeter is a servo-driven instrument that receives electrical signals from the ADC. If the altimeter loses power, the ADC malfunctions, or the encoder loses power, an OFF flag appears on the instrument face.

Avionics

5B-21December 2011


All altimeters have provisions for adjusting the instrument to local barometric pressure in inches of mercury (InHg) and millibars. On encoding altimeters, adjusting the instrument to local pressure has no effect on the encoded altitude information; it is always relative to standard pressure (29.92 InHg/1,013 millibars).

Atrsphhd IndtSmaorTypically, the pilot’s and copilot’s airspeed indicators are identical; both indicate airspeed from 40 to 300 Kts with a pointer moving over a circular scale. An aneroid-driven pointer (barber pole) indicates the maximum allowable airspeed (VMO) depending on aircraft altitude. There is an audible warning if the aircraft exceeds MMO.

Figure 5B-12: Airspeed Indicators

VhratSml phhd IndtSmaorThe pilot and copilot Vertical Speed Indicator (VSI) is a TVI-920D, which combines the functions of a vertical speed indicator, a TCAS display, and a resolution advisory display to show all TCAS information on one indicator.The TVI-920D has a color Liquid Crystal Display (LCD) and two push-buttons; one for selection of the traffic display mode and the other for selection of the traffic display range. The TVI-920D receives TCAS data from the TCAS control and vertical speed data from the air data systems.

5B-22December 2011


DisplayThe TVI-920D liquid crystal display indicator shows: a vertical speed scale and pointer (VSI); TCAS resolution advisories and a TCAS traffic display; TCAS mode annunciators and warning flags.

Figure 5B-13: TVI-920D Liquid Crystal Display Indicator

VSI DisplayThe VSI display shows climb and descent speeds on a white vertical speed scale. The scale is non-linear to optimize readability at low vertical speeds. Each VSI displays aircraft vertical speed from 0 to 6,000 ft per minute (fpm) up and down.

Resolution AdvisoriesTCAS Resolution Advisories (RA) show as red and green bands around the vertical speed scale. To comply with an RA, fly vertical speeds in the green arc areas, and avoid vertical speeds in the red arc.

Traffic DisplayThe traffic display shows nearby traffic with Mode C or Mode S transponders that reply to TCAS interrogations. The traffic display information aids pilots in visually acquiring the nearby traffic.The following represents the traffic display: an airplane symbol in white in the lower center of the display a white-dotted 2/20 NM range ring around the airplane symbol,

depending on the range (20 or 40 NM) selected (units that can show a 40 NMI range show a 20 NMI dotted range ring when set to the 40 NMI range);

four types of TCAS traffic symbols; and TCAS annunciators and flags.

OTTc: Units that can show a 40 NM range show a 20 NM range ring instead of the 2 NM ring when set to the 40 NM range.

Avionics

5B-23December 2011


Airplane SymbolThe “own aircraft” airplane symbol shows in white in the lower middle of the display. The airplane symbol shows your aircraft’s position with respect to the traffic shown on the display. The position of the symbol is such that the distance of the display range shown below the symbol is approximately half the distance shown above the symbol.

Range RingsThe fixed 2 NM range ring shows as dots around the airplane symbol. The dots show the 12 clock positions or bearings around the aircraft. The vertical speed scale is the maximum traffic display range ring for a selected display range.TCAS traffic symbols show for nearby traffic that is within the selected range and altitude mode to reply to TCAS interrogations. The symbols on the traffic display show the bearing, distance, relative altitude (if available), and, as appropriate, the vertical speed of the traffic. Each of the four types of TCAS traffic shows a unique symbol on the traffic display.

Resolution Advisory (RA)

Traffic Advisory (TA)

Proximate Traffic (PT)

Other Traffic (OT)

The relative altitude and the direction of vertical speed show in the same color as the associated traffic symbol. It is displayed in hundreds of ft preceded with “+” or “–”. The “+” indicates altitude above and a “–” indicates altitude below your aircraft. The vertical speed direction is displayed with an up or down arrow, depending on the traffic vertical direction. No arrow is displayed if the traffic vertical speed is less than 500 fpm.Only RA and TA traffic show partial symbols for traffic beyond the selected range of the traffic display. The partial symbols show at the appropriate bearing along the maximum range ring (VSI scale).The altitude data is displayed above the traffic symbol for traffic at an altitude above your aircraft and below the symbol for traffic below your aircraft. Vertical speed direction arrows are displayed to the right of the traffic symbols.Relative altitude data is displayed as two numbers in hundreds of feet preceded by a “+” or “–”sign. The number on the left shows thousands of feet of altitude and the number on the right shows hundreds of feet of altitude (i.e., +22 = 2200 ft above the aircraft and 02 = 200 ft below the aircraft).

5B-24December 2011


The vertical speed direction is displayed with an up or down arrow, depending on the traffic arrows shown for traffic with an actual (not relative) vertical speed equal to or greater than 500 fpm. An upward pointing arrow (↑) shows for climbing traffic and a downward pointing arrow (↓) shows for descending traffic. No arrow shows for traffic climbing or descending at an actual vertical speed of less than 500 fpm.Other traffic and proximate traffic do not show partial symbols.

Annunciators and FlagsThe TVI-920D LCD display shows annunciators and flags for the various TCAS operating modes and system failures (see Table 5B-1 and 2).

Annunciation/ Flag DescriptionABV The above mode annunciator shows the letters ABV in white,

with an amber background, in the upper right corner of the display. In the above mode, nearby traffic from 2,700 ft below, to 9,900 ft above the aircraft, with transponders that reply to TCAS interrogations, show on the traffic display.

BLW The below mode annunciator shows the white letters BLW, on an amber background, in the upper right corner of the display. In the below mode, nearby traffic from 2,700 ft above, to 9,900 ft below the aircraft, with transponders that reply to TCAS interrogations, show on the traffic display.

Range The range annunciator (i.e. 3NM, 5NM, 10NM, etc.) shows in white, with an amber background, in the upper right corner of the display. This annunciator shows the maximum range for traffic shown on the display. The TVI920D has two versions of range modes. One version shows 6 and 12 NMI ranges and the other shows 3, 5, 10,20, and 40 NMI ranges.

ONLY TA The ONLY TA annunciator shows in the upper left corner of the display. If TCAS does not show any TA traffic on the display, the annunciator letters show in white in a white box. If TCAS detects TA traffic, the letters show in amber in an amber box. The TA or TA ONLY mode showsthis annunciator. TCAS does not give resolution advisories when this annunciator shows on the display.

Table 5B-1: Annunciators and Flags

Avionics

5B-25December 2011


Annunciation/ Flag DescriptionTA/RA No Bearing TA and RA traffic, for which TCAS cannot calculate a bearing,

show on a two line annunciator field at the bottom middle of the display. The annunciator shows the type (TA or RA), range and relative or absolute altitude ofthe traffic. The annunciator shows in amber for TA traffic and in red for RA traffic. Range is displayed as a two-digit readout in nautical miles with a resolution of one-tenth of a nautical mile. The altitude data is displayed the same way as the traffic symbols.This annunciator has two numbers and, in the relative mode, a preceding "+" or "-" sign.The number on the left shows thousands of feet of altitude and the number on the right shows hundreds of feet of altitude (i.e., +22 = 2200 ft above the aircraft and –02 = 200 ft below the aircraft). An upward or downward pointing arrow shows with the altitude data for traffic climbing or descending at an actual vertical speed of 500 fpm or greater.

VIS The vertical speed flag shows the black letters VIS, on anamber background, in the lower left corner of the display.When this flag shows, vertical speed data is missing or invalid.

RA The resolution advisory flag shows the black letters RA, on an amber background, in the upper left corner of the display. This flag indicates a failure of the RA function of the VSI. TCAS does not show resolution advisories with this flag displayed.

TCAS The TCAS failure flag shows the black letters TCAS, on an amber background in the upper left corner of the display. This annunciation indicates a TCAS failure, no traffic or resolution advisory will be displayed on the Traffic Vertical Indicator (TVI). This flag is displayed when a TCAS system failure occurs. No traffic information is displayed and no resolution advisories are given when this flag is showing.

TCAS OFF The TCAS OFF annunciator shows in white letters with a white box around the letters in the upper right corner of the display. This annunciator is displayed when TCAS is set to the standby mode or turned off. No traffic information and no resolution advisories data will be displayed on the TVI.

TEST The TEST annunciator shows in white letters with a whitebox around the letters in the bottom middle of the display. This annunciator is displayed while the TCAS system self-test operates.

Table 5B-2: Annunciators and Flags (Continued)

Diagnostic DisplaySome versions of the TVI-920D show a diagnostic failure list of TCAS-related systems that fail or do not provide valid data to the TCAS system. A 10-second (approximate) or longer push of the TCAS self-test button turns on the diagnostic list display. When all TCAS related systems operate properly, only the part number of the TCAS computer software shows on the list. TCAS system failures show below the part number.

OTTc: In some TVI-920D installations the self-test and diagnostic functions operate only when the aircraft is on the round.

5B-26December 2011


ConarolsRange Button (R)Push the R button to select the range of the traffic display. The range annunciator is displayed in the top right corner of the indicator. Successive pushes of the R button cycles through the available ranges.

Display Mode (M)Push the M button to alternately select the display mode of the TVI920D. There are two modes available, a part time threat only traffic display and full-time traffic display.

Part-time Threat Only Traffic DisplayThe part-time mode will display the vertical speed data only until a TA or RA within selected range and altitude mode is detected. When all TA and RA traffic is downgraded to proximate or other traffic, the display changes back to VSI only.

Full-time Traffic DisplayThe full-time traffic display shows the VSI and the traffic display at all times. In this mode, all types of nearby traffic within the selected range and altitude mode, which reply to TCAS interrogations, show on the display at all times.

A/B (“OT” Altitude Mode) Button (TVI with A/B and R buttons)Push the A/B button to select the altitude volume, relative to your own aircraft, for which TCAS shows “Other Traffic”. Successive pushes of the button cycles through the available modes. “Other Traffic” outside of the selected altitude volume does not show on the traffic display. Proximate, TA and RA traffic are not affected and always show on the display, regardless of the selected altitude volume (except RA traffic does not show on the display in the ONLY TA mode). The selected mode is annunciated on the display. The available selections are:No Annunciation - This is normal altitude mode. “Other Traffic” from 2,700 ft below, to 2,700 ft above the aircraft, with transponders that reply to TCAS interrogations, show on the traffic display.ABV - “Other Traffic” from 2,700 ft below, to 9,900 ft above the aircraft, with transponders that reply to TCAS interrogations, show on the traffic display.BLW - “Other Traffic” from 2,700 ft above, to 9,900 ft below the aircraft, with transponders that reply to TCAS interrogations, show on the traffic display.

Avionics

5B-27December 2011


Alataudh PrhshlhSaor/ AlhrahrThe PRE-80C Altitude Preselector/ Alerter is used to preselect an altitude to be captured by the flight control system and provide warning to the pilot when approaching or deviating from the selected altitude. The altitude is displayed in 5-digit numerical format. The altitude preselector/alerter receives input from the altimeter and compares it to actual settings. Depending on the mode setting, the system will generate proper annunciation.

Figure 5B-14: PRE-80C Altitude Preselector/Alerter

The two modes of operation are acquisition (capture) and deviation (after capture).The altitude is displayed in 100 ft increments, from 0 to 49,000 ft, and is selected using the altitude select knob. Pressing and turning the altitude select the knob to select the 100 digit and with the knob out turing the altitude select knob will select the thousand.

Warning FlagThe red warning flag will come on if there is a fault detected or if incoming data is lost. If a fault is detected in the aural warning system, the ALT ALERT will flash at the same time as the alert.

Alataudh Alhra Ltgca/ PUH TO CACTL/Thsa wtaScWhen the system is in the acquisition mode, the ALT ALERT light illuminates within a 1,000 ft of the preselected altitude. After the aircraft is within 300 ft of preselected altitude, the system switches to deviation mode. If the aircraft exceeds the preselected altitude by 300 ft the ALT ALERT light will flash. Pressing the PUSH TO CANCEL switch will extinguish any ALT ALERT indication. Selection of a new altitude will set the system in the acquisition mode. An aural warning tone is generated at acquisition or deviation.

5B-28December 2011


Rmdto AlatehahrCollins ALT-50 radio altimeters provide precise aircraft altitude Above Ground Level (AGL) during approach and landing or when the aircraft is below 2,000 ft AGL. The system consists of a transceiver, indicator, and transmit and receive antennas on the lower fuselage. The transceiver transmits a signal toward the ground, receives the bounced signal, and computes altitude by the time delay between transmission and reception.The system provides absolute altitude information from -20 to 2,000 ft AGL in 100 ft graduations down to 500 ft and in 10 ft graduations from 500 to 20 ft on an indicator on the pilot’s instrument panel and/or the attitude director indicator. The radio altimeter system also provides altitude information to the flight profile advisory system.The radio altimeter allows the preselection of a desired Decision Height (DH) for altitude alerting purposes. To preselect the DH, rotate the DH SET knob until the index aligns with the desired altitude. Once the aircraft reaches the desired altitude, the DH light illuminates.

Figure 5B-15: ALT-50 Radio Altimeters

TehrghnSy LoSmaorTrmnsetaahrThe airplane is equipped with an Emergency Locator Transmitter (ELT), to assist in tracking and recovery of the airplane and crew in the event of a crash or emergency landing. The ELT is mounted on the right side approximately four feet aft of the pressure bulkhead. An access panel is located adjacent to the transmitter and an antenna is mounted on the fuselage aft of the unit. The output frequencies are 121.5 and 243 MHz, simultaneously and its maximum range is line of sight. An impact switch activates the transmitter if the switches sustain excessive G force.

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Magnetic CompassA conventional, liquid-filled magnetic compass on the windshield center post provides aircraft heading information. The compass contains provisions for maintenance personnel to adjust the unit to compensate for aircraft generated magnetic fields. A correction card near the unit provides a record of recent adjustments to the compass and compass deviation errors.Heading information from the compass is accurate only in level, unaccelerated flight.

Turn mnd ltp IndtSmaorsThe pilot and copilot have a turn and slip (turn and bank) indicator on their lower instrument panels. Each indicator provides an indication of aircraft rate-of-turn and direction and assists in proper turn coordination. The pilot’s indicator has an electrically-driven gyroscope and the copilot’s indicator uses a vacuum-driven gyroscope. An inclinometer at the bottom of each instrument uses a curved, liquid-filled tube with a steel ball to indicate aircraft slip (yaw).

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CoSkpta VotSh RhSordhrThe Cockpit Voice Recorder (CVR) is on any time the airplane power is on. The CVR stores the last 30 minutes of conversation at all times. The CVR-30B is also capable of storing other data such as: Greenwich Mean Time (GMT) and CVR internal Built In Test Equipment (BITE) status. There is an impact switch, which will shut off the recorder on impact. This prevents the recorder from recording over the last conversation before impact. The cockpit voice recorder microphone is mounted in the glareshield to the right of the centerline in front of the copilot. The control unit for the cockpit voice recorder is mounted in the pedestal extension. The control unit consists of a headset jack, test switch, erase switch and test circuit meter. With the airplane on the ground, the erase switch may be used to erase what is stored in the recorder. The cockpit voice recorder and impact switch are located on the avionics shelves, aft of the aft pressure bulkhead. Any time the impact switch has been activated, the indicator light on the switch should illuminate to indicate that the reset button needs to be depressed. Attached to the front of the CVR is an Underwater Locator Beacon (ULB). The ULB is water-activated. This ULB is battery operated with a service life as indicated in Chapter 5-11-00 of the AMM. The actual operating life of the batteries is at least 30 days. The expiration date for the batteries is noted on the front panel. The ULB must be functionally tested as indicated in Chapter 5-11-00 of the AMM to ensure proper operation. Whether in the airplane or the shop, a portable test set should be used to test the ULB. For ULB model DK100, if the battery life has expired or other maintenance is indicated, the ULB must be removed from the CVR and returned to Dukane for service. No maintenance on the ULB is required or allowed in the field. ULB model DK120 has a user replaceable battery; for other maintenance the ULB must be removed from the CVR and returned to Dukane for service. (Loral will not service the ULB). For more information refer to the Beech King Air Series Component Maintenance Manual.

Figure 5B-16: Cockpit Voice Recorder (CVR)

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Communication SystemsCommunications equipment on the King Air includes: audio integrating speech communication interphone static discharging.

Audio IntegratingAudio integrating includes the audio control panels, headsets, microphones, and cockpit loudspeakers.

MtSropconhs/HhmdshasThe microphone and headset jacks are on the pilot’s and copilot’s side panels. Each jack connects with the respective audio control panel. A microphone switch on the pilot’s and copilot’s subpanels selects the normal microphone or the oxygen mask microphone.

Figure 5B-17: Microphone and Headset Jacks

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Audto Conarol PmnhlsAudio Control Panels for the pilot and copilot contain controls for audio source selection, microphone output selection, and volume control of the headsets and cockpit loudspeakers. Each panel accepts audio inputs from the communication and navigation receivers and provides individual switches for each audio source.

Figure 5B-18: Audio Control Panels

A master volume switch on each panel individually adjusts the volume of the combined audio sources to the headsets or loudspeakers. Depending on the audio panel installed, the cockpit loudspeakers and headphones are always operational (headphones plugged in) or are not operational at all times (either loudspeaker or headphones). One system has a switch that turns off the cockpit loudspeaker; the other system has a switch that selects either the headphones or the loudspeakers.Each audio panel has an emergency switch that bypasses the audio amplifiers to connect audio output directly from the source to the headsets if the audio panel malfunctions. If the emergency switch fails, pull the applicable audio control circuit breaker(s).Microphone output selection is made through a rotary selector switch on each panel that connects the microphone to the communications transmitters or the cabin speakers. A second microphone switch selects normal microphone operation or hot microphone operation. A hot microphone allows communication between the crew through headsets without -having to key the headset microphone.

Speech CommunicationSpeech communication includes: Very High Frequency Communication Transceivers (VHF COM) High Frequency communication transceivers (HF) radio telephone.

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VHF CoeeuntSmatonsTypical VHF transceivers provide air-to-air, air-to-ground, and ground-to- ground communications. The unit operates in the 118.000 to 135.975 or 136.975 MHz frequency range with a frequency spacing of 25 kHz that provides 720 or 760 channels. Optional transceivers have an extended frequency range of 118 to 151.975; this provides 1,360 distinct channels.Each complete VHF transceiver installation consists of a control head, a transceiver, and an antenna. The control head contains controls for frequency selection, frequency storing function, light sensor and display, volume and power control, and squelch activation, a transfer memory selection and testing.Up to six preset frequencies can be stored in the nonvolatile memory. The transceivers are in the nose avionics compartment and a blade-type VHF antenna is on the top and bottom of the fuselage.

Figure 5B-19: VHF Transceivers Control Head

Audto TehrghnSy Ophrmaton wtaScIn the event of failure of the pilot’s audio system, access to audio is available through the copilot’s speaker. If both systems have failed, place the emergency/normal switch in the EMER position.

TehrghnSy OphrmatonWith emergency/normal switch in EMER position, the following rules apply: all audio sources (COMM 1, NAV 2, ADF, etc.) are connected directly

to the headphones to eliminate any specific audio source, turn volume control down on that

source (e.g., NAV 1). This rule does not apply to COMM 1 and COMM 2. volume control on microphone selector switch has no function in the

emergency mode if emergency/normal switch fails, the audio systems may be placed

into emergency mode by pulling the two Circuit Breakers (CBs) labeled PLT AUDIO and COPLT AUDIO located directly beneath the avionics master CB on the main CB panel.

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HF CoeeuntSmatonSome aircraft use High Frequency (HF) communications equipment to allow very long range communications. Typical systems operate in the 2.0000 to 29.9999 MHz range with frequency spacing of 100 Hz; this provides 280,000 distinct channels. Most HF transceivers provide Amplitude Modulation (AM) and Single Side Band (SSB) transmission modes.Each installation consists of a transceiver, control head, power amplifier/ antenna coupler, and a long wire antenna.

Rmdto ThlhpconhA radio-telephone allows the crew or passengers to communicate with ground stations through the public telephone system, with mobile telephones, or other aircraft with radio telephones on the HF and Ultra-High Frequency (UHF) radio frequencies. The system also allows communication between the cockpit and passenger cabin.Typical systems consist of a transceiver, antenna, cockpit unit, and passenger cabin unit. Depending on the system installed in the aircraft, the cockpit and cabin units consists of a handset with panel-mounted controls or an integrated system; the integrated system has controls in the handset. With both systems, the controls consist of channel selector buttons, a combined power switch and volume control, and intercom and transmit indicator/buttons.

Static DischargingStatic discharging wicks on the aircraft structure and control surfaces minimize the effects of lightning strikes and static charges on avionics equipment. The static dischargers bleed off accumulated static charges to the atmosphere. Due to varying configurations, consult your MEL for number and position of static wicks.During the preflight inspection, check the security, condition and presence of the dischargers. The minimum required for dispatch is one on each control surface; the outboard discharger on each aileron must be present.

Figure 5B-20: Static Discharging Wicks

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NavigationNavigation equipment can be loosely grouped into equipment that provides aircraft direction and attitude information, equipment that determines aircraft position, and equipment that provides flight management.

Attitude and DirectionAttitude and direction equipment use inertial and magnetic forces to sense and display aircraft heading and attitude. This includes: Magnetic Compass Turn And Slip Indicator Gyro Horizon/Vertical Gyro Radio Magnetic Indicator Vertical Gyro System Compass System.

MmgnhatS CoepmssA conventional, liquid filled magnetic compass on the windshield center post provides aircraft heading information. The compass contains provisions for maintenance personnel to adjust the unit to compensate for aircraft generated magnetic fields. A correction card near the unit provides a record of recent adjustments to the compass and compass deviation errors.Heading information from the compass is only accurate in level unaccelerated flight.

Figure 5B-21: Magnetic Compass

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Turn mnd ltp IndtSmaorsThe pilot and copilot have a Turn And Slip (turn and bank) indicator on their lower instrument -panels.Each indicator provides an indication of aircraft rate-of-turn and direction and assists in proper turn coordination. The pilot’s indicator has an electrically-driven gyroscope and the copilot’s indicator uses a vacuum-driven gyroscope. An inclinometer at the bottom of each instrument uses a curved, liquid-filled tube with a steel ball to indicate aircraft slip (yaw). On some installations the turn and slip indicator provides yaw data for yaw dampening.The pilot’s turn-and-bank indicator operates on 28 VDC. If the unit loses power, a GYRO warning flag appears on the instrument face.

Figure 5B-22: Turn and Slip Indicators

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Gyro-Hortzon/VhratSml GyroOn aircraft with single flight director systems, the pilot has an Attitude Director Indicator (ADI) driven by the flight director system and the copilot has a vacuum-driven gyro-horizon. Aircraft with dual flight directors have an electrically-driven standby (emergency) gyro horizon on the pilot’s instrument panel. The gyro-horizon indicates aircraft attitude with a fixed airplane symbol over a moving sphere. The sphere has a horizon line and degree marks above and below the horizon line to indicate pitch angles. A knob on the instrument face adjusts the aircraft symbol to correct for variations in level flight attitudes.

Figure 5B-23: Horizon/Vertical Gyro

One (single flight director system) or two vertical gyros (dual flight director systems) provide aircraft pitch and roll information to the autopilot, flight instruments, flight director, and radar antenna stabilization system.Each unit consists of an electrically driven gyro rotating on its vertical axis. Gimbals within the unit limit the amount of freedom in the pitch and roll axes. The gyro is free to pitch approximately 80° up and down, and roll 360° (roll unlimited). The vertical gyro(s) in the nose compartment electrically drives servoed instruments on the instrument panel.A FAST ERECT switch on the instrument panel for each vertical gyro manually erects the vertical gyro.

Figure 5B-24: FAST ERECT Switch

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Coepmss ysaheTwo compass systems (directional gyros) provide 360° of magnetic heading information to the pilot’s and co-pilot’s Horizontal Situation Indicators (HSIs), Radio Magnetic Indicators (RMIs), autopilot, and flight director. The pilot’s system drives the copilot’s RMI and the copilot’s system drives the pilot’s RMI. Each directional gyro installation consists of a directional gyro, a flux valve, remote compensator and control switches. The directional gyros in the nose compartment electrically drive the RMIs, flight director system, and the horizontal situation indicator (s).Each gyro consists of an electrically-driven gyro with monitoring and control circuits. A flux valve in each outboard wing or horizontal stabilizer senses the strength and direction of the earth’s magnetic field and converts it into electrical signals for gyro compensation. The compensating signal applied to the gyro aligns the gyro with magnetic north. The directional gyro sees this input as an error signal that it compares to a reference signal. The difference between the error signal and the reference signal produces a signal that drives a slave torquer motor in the gyro. The gyro then precesses to align itself with the flux valve and magnetic north.The remote compensator uses adjustable permanent magnets to counteract the effect of magnetic fields generated by direct current and ferrous materials in the aircraft on the flux valves.A GYRO/SLAVE/FREE switch for each directional gyro allows the selection of either slaved or free gyro operation. In SLAVE, the directional gyro follows signals provided by the flux valve. In FREE, the directional gyro operates independently from the flux valve; manual correction of the gyro is through the INCREASE/ DECREASE switch.To manually align the compass system, use the INCREASE/DECREASE until the compass card aligns with magnetic heading.

Figure 5B-25: GYRO/SLAVE/FREE Switch

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Rmdto MmgnhatS IndtSmaorsTwo radio magnetic indicators display aircraft heading information on a calibrated servo-driven compass card. A pointer and compass card provide bearing indication to either VOR or ADF stations.

Figure 5B-26: Radio Magnetic Indicators

Position DeterminingPosition determining equipment includes systems that operate independently of ground stations or with ground stations to determine aircraft position. This includes: Instrument Landing System (ILS) Very High Frequency (VHF) navigation equipment Automatic Direction Finding (ADF) Distance Measuring Equipment (DME) transponder long range navigation equipment LORAN global positioning system flight management system weather radar.

Insaruehna Lmndtng ysaheInstrument Landing Systems (ILS) combine outputs from the VHF navigation, UHF glideslope and marker beacon receivers to display ILS information on the attitude director indicator and the horizontal situation indicator.The system consists of a glideslope receiver operating in the 329.15 to 335.00 MHz frequency range, the VHF receiver in LOC mode operating in the 108.10 to 111.95 MHz frequency range, a glideslope antenna in the nose, and a LOC antenna on each side of the vertical stabilizer.

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VHF mvtgmatonVHF navigation receivers provide Very High Frequency Omni-Range (VOR), Localizer (LOC), Glideslope (GS), and marker beacon navigation information to the flight crew.Each VHF NAV system receives 200 VHF frequencies from 108.00 to 117.95 with 50 MHz spacing, 40 paired glideslope frequencies from 329.15 to 335.00 MHz spaced at 150 kHz, and 40 LOC frequencies from 108.10 to 111.95 MHz. Automatic DME channeling is through the navigation receiver. Multiple outputs from the receivers drive the flight director, Radio Magnetic Indicators (RMIs), autopilot, course deviation indicators, and area navigation equipment (RNAV). The receiver supplies audio output to the audio control units.Receiver control, frequency selection, and frequency display are through control heads on the center instrument panel.As part of the VHF navigation receiver, a marker beacon receiver provides visual and aural indications of beacon passage. The system receives on 75 MHz and provides electrical outputs to two sets of three indicating lights on the instrument panel. The receiver also provides audio output to the audio control units for beacon passage notification.

Figure 5B-27: Receiver Control Heads

AuaoematS DtrhSaton FtndtngAutomatic Direction Finder (ADF) systems consist of a receiver, control head, and a combined loop and sensing antenna. The receiver operates in the 190.0 to 1749.5 kHz frequency range with 0.5 kHz spacing that provides 3,120 distinct frequencies.ADF systems provide three basic modes of operation: Antenna (ANT), Automatic Direction Finding (ADF), and Tone (TONE). In antenna mode, the RMI pointer parks and the system provides only audio output. ADF mode provides continuous relative bearing readings to low frequency radio beacons, and AM broadcast stations and audio outputs to the audio control panels.Tone mode provides a 1000 Hz tone for identification of station identifiers.

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DtsamnSh MhmsurtngDistance Measuring Equipment (DME) computes and provides slant range distance between the aircraft and a VORTAC facility. The system transmits in the 1025 to 1150 MHz range and receives in the 962 to 1213 MHz range. Pairing of DME channels with VHF navigation frequencies allows the automatic selection of DME channels by the VHF receiver.The DME system provides distance, speed, and time information to the Horizontal Situation Indicators (HSIs), and DME displays (Figure 5B-28).

Figure 5B-28: DME Displays

Arhm mvtgmatonArea navigation systems (RNAV) allow point-to-point navigation within the coverage of VHF navigation facilities (VOR/VORTAC/DME). Most systems allow the storage of flight plans containing multiple waypoints for frequently flown routes.These systems utilize data provided by the VHF, DME, and localizer receivers to compute and display waypoint information.

TrmnspondhrTypical transponder systems with Mode C or Mode S capability provide identification and altitude reporting to surveillance radar installations. The system consists of a transceiver, control head, and a transmit/receive antenna. The system transmits on 1090 MHz and receives on 1030 MHz. The pilot’s encoding altimeter provides aircraft altitude information to the transponder system for transmission to ATC radar facilities.

Oehgm/VLFVery Low Frequency (VLF), VLF/Omega, and Omega/VLF navigation systems provide great circle, point-topoint navigation on a world-wide basis. These systems utilize very low frequency transmissions from Omega and U.S. Navy facilities.

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Typical systems consist of a control display unit and a receiver-computer unit. The control display unit contains a display and keyboard for data entry and navigational data selection. The display presents system initialization, self-test, flight plan, and navigational information.The receiver-computer unit processes all inputs and provides position co-ordinates, distance and deviation information, drift and track angle deviation, wind direction and speed, and ground speed to the control display unit. The receiver-computer also provides inputs to the Horizontal Situation Indicators (HSIs), autopilot, and flight director system. Loss of navigation facility signals causes the system to revert to dead reckoning based on aircraft heading, true airspeed, and last computed winds.Eight Omega stations operated by the U.S. Coast Guard broadcast between 9 to 14 kHz;10 VLF stations operated by the U.S. Navy broadcast between 14 to 24 kHz. The positioning of the Omega and VLF stations and the characteristics of low-frequency radio signals provides world-wide coverage.Atmospheric conditions, solar activity, height of the ionosphere, and variations in the earth’s magnetic field all affect the accuracy of the Omega/VLF system. Because very low frequency signals bounce between the ionosphere and the earth’s surface, any changes in these factors affect signal accuracy, reduce signal strength, and increase background noise to result in unreliable system operation.

LORALORAN (long range navigation) systems use chains of low frequency transmitting stations to determine aircraft position with a relatively high degree of accuracy. The LORAN C chains consist of a master transmitter (M) and up to four slave transmitters (W, X, Y, and Z). Each chain transmits in the 90 to 110 kHz frequency range with the slaves transmitting in a set order after the master. The location and arrangement of LORAN C transmitter chains provides coverage of North America, the Middle East, Asia, and the Pacific Rim.Each aircraft system consists of a computer controller unit and an antenna. The controller contains a receiver, computer processor, and a interface (keyboard/display). After entering the aircraft’s present position and destination, the system receives signals from the master and slave LORAN C transmitters. The computer/controller processes these signals and determines by recognizing the time difference between the master and slave transmitter signals. The computer/controller processes this information to establish a hyperbolic line of position; this provides aircraft longitude and latitude with an accuracy of approximately 60 ft.

Globml Postatontng ysaheGlobal positioning systems (GPS) receive transmissions from a constellation of 24 satellites (21 operational/ three spares) orbiting the earth in six orbital planes. The arrangement of these orbits and the satellites within them ensures that six to ten satellites are visible from any point on the earth.

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Using transmissions from four satellites received by a GPS sensor, the NCU can compute aircraft longitude, latitude, and altitude. With only three satellites visible, the system can compute only longitude and latitude. Even when limited functions are available, the system can determine aircraft position within 100 meters (civilian use restriction).

Fltgca Mmnmghehna ysahe Flight Management Systems (FMS) utilize position information from various navigation equipment to provide an integrated navigation display and control system.The FMS receives inputs from the VHF navigation equipment, computes the aircraft position, and provides outputs for the autopilot, flight director, RMIs, and Horizontal Situation Indicators (HSIs).

Whmachr RmdmrWeather radar systems consist of an antenna, receiver-transmitter, and radar display with system controls. The vertical gyro system provides aircraft attitude information to the radar system to stabilize the antenna. The system operates by transmitting a high frequency radio signal, receiving the bounced signal, and displaying the received signals on the display. Controls on and below the indicator select system mode, scan range, antenna tilt, and receiver gain (sensitivity). Typical systems provide: selectable scanning range ground mapping weather cell contouring adjustable antenna tilt and scan target alerting navigation information overlay (some systems).

Figure 5B-29: Weather Radar Systems

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Radar power output and scanning area varies between equipment manufacturer, model, and radar capabilities. Hazard areas presented in this discussion come from various aircraft and component maintenance manuals and FAA Advisory Circular AC 2068B. Personnel hazard areas are the maximum recommended hazard area for radar operation on the ground.When operating radar on the ground, precautions should be taken to avoid injury to personnel, fuel ignition, or radar equipment damage. Avoid operating the radar during refueling or within 300 ft of refueling aircraft. Caution personnel to remain outside an area within 270° and 15 ft forward of the radome. Direct the nose of the aircraft so a 240° sector forward of the aircraft is free for a distance of 100 ft of large obstructions, hangars, and other buildings. Tilt the antenna up to its maximum angle.

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Flight Control System

Flight Management System

Untvhrsml U-1KThe UNS-1K is a multi-sensor flight management system that consists of one Flat Panel Control Display Unit (FPCDU), a Navigation Computer Unit (NCU) and a Global Positioning System (GPS) antenna.

Figure 5B-30: Flat Panel Control Display Unit

The UNS-1K is a fully integrated navigation management system designed to provide the pilot with centralized control for the airplane’s navigation sensors, computer based flight planning, and fuel management.The FMS accepts primary position information from short- and long-range navigation sensors. The primary position data received from the sensors is filtered within the FMS to derive a Best Computed Position (BCP). It accomplishes these computations and advises the flight crew of components or systems requiring attention, as well as other irregularities, such as loss of enough sensors to compute a valid position. In the latter situation, if sensor loss endures over a set length of time, the system will enter Dead Reckoning (DR) mode and so, inform the pilot through a message on the Control Display Unit (CDU) and display a red-boxed “FMS” on the MFD.The UNS-1K provides lateral steering information to the pilot through the EHSI. When connected to the autopilot, it provides roll steering commands. The VNAV function provides vertical steering information displayed on the UNS-1K FPCDU. VNAV guidance is not provided to the flight director or autopilot. The NAV computer additionally computes fuel quantity information, providing a current fuel status and airplane gross weight throughout the flight, if the fuel and gross weight are updated prior to takeoff.The UNS-1K database incorporates Standard Instrument Approaches (SID), Standard Terminal Arrivals (STAR), and approaches including GPS approaches.

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These procedures may be flown coupled to the autopilot or flight director. The internal database must be updated to the latest revision every 28 days; updating to be accomplished with the universal avionics update disk or equivalent.The internal 12-channel GPS receiver provides real-time and predictive Receiver Autonomous Integrity Monitoring (RAIM), automatic Fault Detection And Exclusion (FOE), step detection, and manual satellite deselection capabilities. GPS can be used to fly in all phases of flight as a stand-alone. GPS stand-alone and GPS-overlay approaches are contained in the database complete with transitions and missed approach procedures.

Turbulence Weather RadarTurbulence Weather radar systems consist of an antenna and receiver-transmitter; radar display is done through the EFIS system with system control panel. The vertical gyro system provides aircraft attitude information to the radar system to stabilize the antenna. The system operates by transmitting a high frequency radio signal, receiving the bounced signal, and displaying the received signals on the display. Controls on the weather radar panel and on the MFD select system mode, scan range, antenna tilt, receiver gain (sensitivity), and slave operation. Typical systems provide: selectable scanning range weather display hold ground mapping weather cell contouring adjustable antenna tilt, auto-tilt and scan sector scan turbulence weather target alerting ground clutter suppression path attenuation correction and alert navigation information overlay.

Figure 5B-31: Turbulence Weather Radar Systems

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Radar power output and scanning area varies between equipment manufacturer, model, and radar capabilities. Hazard areas presented in this discussion come from various aircraft and component maintenance manuals and FAA Advisory Circular AC 2068B. Personnel hazard areas are the maximum recommended hazard area for radar operation on the ground.When operating radar on the ground, precautions should be taken to avoid injury to personnel, fuel ignition, or radar equipment damage. Avoid operating the radar during refueling or within 300 ft of refueling aircraft. Caution personnel to remain outside an area within 270° and 30 ft forward of the radome. Direct the nose of the aircraft so a 240° sector forward of the aircraft is free for a distance of 100 ft of large obstructions, hangars, and other buildings. Tilt the antenna up to its maximum angle.

Flight Director SystemThe King Air 350 has a Collins FCS65 Flight Control System (FCS). The system is a combination of autopilot, guidance displays and sensors. The system provides three axes (roll, yaw and pitch), flight control with elevator trim, yaw damping and rudder boost. The system combines the following sub-systems into a flight control system: autopilot system flight director system electronic flight instrumentation system magnetic compass system air data system.

The flight control system provide three modes of operation: manual, automatic, and semi-automatic. Manual operation allows the crew to fly the aircraft guided by cues provided by the flight director system. Automatic operation flies the aircraft through the autopilot coupled with the flight director; the crew monitors system operation. Semi-automatic operation flies the aircraft through crew interaction with the autopilot controller and controls on the control wheels. Refer to the applicable pilot’s guides, Aircraft Flight Manual Supplements and component maintenance manuals for a thorough discussion of these systems and their operating procedures.

Fltgca DtrhSaor ysahe Fltgca GutdmnSh Coepuahr (FGC)The flight guidance computer receives inputs from the flight control panel, vertical gyro, air data system, compass system, and radio navigation systems. These inputs are processed to provide vertical and lateral steering commands for the flight instrument system command bars or cross-pointers, system annunciation, monitoring, and diagnostic testing.

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The FGC is matched to each airplane type by a configuration module that is plugged into the top of the unit. This module establishes system gains and provides special operating instructions that may be unique to each airplane type. The configuration module is electrically interlocked in the airplane’s wiring to ensure that the correct computer is always installed.

Fltgca Conarol PmnhlThe Flight Control Panel includes mode annunciators, a test button and mode select buttons to select the following operating modes of the Flight Guidance System: roll hold Navigation (NAV) approach pitch hold Indicated Airspeed Hold (IAS) Vertical Speed hold (VS) Altitude Preselect (ALT SEL) Descent (DSC) Vertical Navigation (VNAV)

Mode selection is annunciated on the flight control panel (Table 5B-3 and 4).

Figure 5B-32: Flight Control Panel

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Annunciator Indication Color AP Autopilot is engaged. Green DIS (AP) Autopilot is disengaged. DIS portion of annunciator will flash for 5 seconds, then become

steady. DIS also illuminated when autopilot is engaged and the SYNC button is pressed. Amber

YAW Yaw damper is engaged. Green DIS (YAW) Yaw damper is disengaged. DIS portion of annunciator will flash for 5 seconds, then become

steady. DIS also illuminated when yaw damper is engaged and the SYNC button is pressed. Amber

HDG Selection of heading mode. Also illuminates if NAV or APPR is selected but capture has not occurred. Green NAV After NAV mode is selected and after capture occurs. Green ARM (NAV) After NAV mode is selected and before capture occurs. *Amber or WhiteDR System is operating in dead reckoning mode. Green APPR After APPR mode is selected and after capture occurs. Green ARM (APPR) After APPR mode is selected, after an ILS, VOR, or LOC frequency is tuned, and before capture

occurs. *Amber or White

B/C or REV After APPR and B/C (back-course) mode are selected. ARM (APPR) also illuminates when a LOC frequency is tuned.

Green

ALT Selection of altitude hold mode or after ALT SEL is selected and after altitude capture occurs. Green ALT SEL Selection of altitude preselect mode. Green VS Selection of vertical speed hold mode. Green GS After glideslope capture in APPR mode on a front-course approach. Green ARM (GS) After APPR mode is selected, a LOC frequency is tuned, and a glideslope valid is received on

a front-course approach. *Amber or White

IAS Selection of indicated airspeed hold mode. Green GA Selection of go-around mode. Green AP Autopilot failure. Red TRIM Trim system failure. Red TRIM Trim system operating (trim in motion). *Amber or WhiteCLM Selection of climb mode. Green DSC Selection of descent mode. Green 1/21/> Selection of half-bank mode. Green

* Annunciator color depends upon the CPN of the FCP-65 and the MAP-65Table 5B-3: Flight Control Panel

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Mode selection is interlocked to ensure that only compatible modes can be selected at the same time. When no lateral modes are selected, the command bars on the PFD or ADI are removed from view except during and after selection of go-around mode.The TEST button is a momentary-action pushbutton that selects the system diagnostic mode. This mode can be entered on the ground or in the air. In diagnostic mode, a lamp test of the system annunciators is performed, followed by a series of test routines to determine the operational status of the system. Both airborne and ground test modes provide the same diagnostic capability except in the area of assessing servo operation. In airborne test mode, servo related fault annunciators indicate the actual operational status of the servos. In ground test mode, the diagnostics do not provide valid indications of servo operation.The flight control system is designed to have smooth transition between operating modes. Whether engaged or disengaged, the system provides steering commands. These same commands are used by the autopilot when engaged.

Flashing Annunciator

Indication Color

SR Selection of soft-ride mode. GreenOM (MAP only) Passage of outer marker. Blue MM (MAP only) Passage of middle marker. AmberIM (MAP only) Passage of inner marker. White VNAV Selection of vertical NAV mode. GreenARM (VNAV) System is armed to capture VNAV path. White ALT ARM System is armed to capture a preselected altitude. HDG Loss of attitude or compass monitor. GreenNAV Loss of attitude, compass or NAV monitor. GreenAPPR Loss of attitude, compass or NAV monitor. GreenDIS (YAW) Flashes for 5 seconds, then steady indication after yaw disengagement. AmberALT Loss of attitude or air data sensor validity. GreenALT SEL Loss of attitude or preselector flag, or air data sensor validity. GreenVS Loss of attitude flag or air data sensor validity. GreenGS Loss of attitude or glideslope flag. GreenIAS Loss of attitude or air data sensor validity. GreenDSC Loss of attitude flag or air data sensor validity. GreenCLM Loss of attitude flag or air data sensor validity. GreenGA Loss of attitude flag. GreenDIS (AP) Flashes for 5 seconds, then steady indication after AP disengagement. Amber

Table 5B-4: Flight Control Panel (continued)

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Lmahrml ModhsRoll Hold ModeSelected with the Turn Knob, Manual Roll Hold mode is the basic lateral operating mode when no other lateral modes are selected. The selected roll attitude (provided it is greater than 5°) can be held if the autopilot is engaged. The SYNC button on the control yoke, or the Turn Knob can be used to change the bank angle. Bank angle is limited to 30°.When the system is in Manual Roll Hold mode, the steering command (Vbars) on the ADI are biased out of view.With the Turn Knob in the detent, or when the SYNC button is released, the system will switch to a wings level attitude hold provided the bank angle is less than 5°.During ILS approach mode, interlocks prevent taking command by the Turn Knob, or SYNC buttons.During other lateral mode operations, the SYNC button disengages the autopilot servos, allowing the pilot to maneuver the aircraft manually until the button is released. After button release, the system returns to the previously selected lateral mode.

Navigation ModeCapture and track selected navigation aid from either VOR, ILS or LOC. Pressing the NAV button once with course selected will put the system in a NAV ARM submode, HDG and NAV are annunciated and system will respond to heading from the heading knob until the track is captured. When the flight path is within limit of course deviation, the system captures the course and NAV ARM change to NAV and all steering commands are done through change of radio signal.

Approach ModeApproach mode is used any time an approach to a runway is executed and consists of the following:VOR – Selected APPR mode with VOR frequency selected the system will arm and provide guidance to intercept the course. Once the course has been captured, guidance is given to maintain approach until an input is given. Glideslope is locked out and the system displays HDG, APPR, ARM APPR and DR (at station passage) at the appropriate stage of the approach. RNAV – This is similar to VOR but with more accuracy. ILS – With ILS station selected the system captures the localizer,

the GS ARM annunciator illuminates and is ready to capture the glideslope beam. Capture can be done from any vertical mode; it occurs before indication on the EHSI or EADI. At capture indication changes to GS and other vertical modes are canceled.

Back Course – Back course is a localizer approach done by selecting B/C button. This is similar to the ILS except the B/C annunciator illuminates and once the course is captured, the approach must be switched to HDG and manual flow.

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GO AROUND – The GO AROUND switch located on the left hand throttle will disengage the autopilot and illuminate the GA annunciator. Can be selected from any mode and it is cancelled by selecting any lateral mode.

VhratSml ModhsPitch hold ModeWhen the autopilot is engaged, or another lateral mode is selected, the Pitch Hold mode can be selected provided no other vertical modes are active. The system responds to the selected lateral mode and to the pitch attitude at the time the lateral mode was selected, or the autopilot was engaged.The steering display is in view provided a lateral mode is selected.The pitch attitude can be changed by either the SYNC button on the control wheel, or the vertical control on the autopilot panel. Pressing a vertical mode momentarily provides a 0.5 ° pitch change. Holding a pitch command for longer than 1.0 seconds, provides a 1° per second slew rate until released.

Altitude Hold Mode (ALT)This mode will maintain the altitude at the time the mode was selected, if autopilot is engaged or a lateral mode selected. Altitude mode should be selected as aircraft reaches desired altitude and rate of climb is less the 500 ft per minute. Change can be made with the vertical switch and canceled by selecting IAS or VS mode. Pressing sync button during glideslope operation will cancel vertical hold and activate pitch hold mode.

Indicated Airspeed Hold Mode (IAS)Selecting IAS will maintain the airspeed at the time the system was initiated, if autopilot is engaged or a lateral mode selected. Speed reference can be changed in one-knot increments using the vertical control and hold mode can be cancelled by selecting ALT or VSl mode. Pressing the sync button during glideslope operation will cancel indicated airspeed hold mode and activate pitch hold mode.

Vertical Speed Mode (VS)Selecting IAS will maintain the vertical speed at the time the system was initiated, if autopilot is engaged or a lateral mode selected. Speed reference can be changed using the vertical control and mode can be canceled by selecting ALT or IAS mode. Pressing sync button during glideslope operation will cancel vertical speed hold and activate pitch hold mode.

Altitude Preselect Mode (ALT SEL)With an altitude selected with the altitude preselector/alerter and other vertical mode selected, pressing the ALT SEL, CLIMB or DSC will arm the altitude preselect mode. Guidance is given to capture the selected altitude; after altitude is captured the system will switch to altitude hold mode.

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Descent (DSC) and Climb (CLB) ModesThese modes provide a smooth transition between present and selected altitude at a preprogrammed vertical speed. Once the system is armed, a DSC or CLM and an ALT ARM annunciator will be displayed. Selecting ALT ARM will cancel climb or descent mode while the ALT ARM annunciator is illuminated.Pressing the vertical control button in the descent (DSC) mode, provides a ±200 ft per minute incremental change each time the button is pressed. In climb (CLB) mode, the rate of climb is changed by ±50 ft per minute.

AuaoptloaThe autopilot system provides automatic control and stabilization of the aircraft about the pitch, roll, and yaw axes. It positions the aircraft elevator, ailerons, and rudder in response to autopilot/flight computer steering commands. Selectable operating modes provide the ability to automatically maintain a desired altitude, pitch attitude or heading, and to capture automatically and track localizer, glideslope, and VOR signals.The autopilot system on the Super King Air consists of: autopilot computer, autopilot panel, flight control panel, altitude preselector/alerter; normal accelerometer; and primary servos and servo mounts.The autopilot system receives signals from the ADC, vertical accelerometer, attitude heading system, and navigation receivers. With this data, the autopilot drives the servo-actuators to maintain a desired altitude, attitude, navigation path, or airspeed. A typical autopilot system provides: yaw damping roll rate and bank angle limiting automatic capture and track of VOR, ILS, and localizer heading, roll, airspeed, and altitude capture and hold heading select soft-ride.

Auaoptloa ConarollhrAutopilot control and operation is through an autopilot controller or autopilot engage and manual controllers on the pedestal. Controls common to most autopilot systems include: yaw engage button autopilot engage button turn knob vertical control soft-ride mode select button ½ mode select button.

Pushing the YAW ENG buttons once actuates the yaw channel of the autopilot and engages the rudder servo.

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Pushing a second time disengages the rudder servo. Pushing the AP ENG button once engages all autopilot servos; pushing a second time disengages all servos except the rudder servo, which disengages by pushing the YAW ENG button or the disconnect switch on the control wheel.Each time the autopilot or yaw system is engaged, the computer performs a pre-engage test routine during which the AP and TRIM failure annunciators may illuminate (depends upon the APC, refer to the Airplane Flight Manual Supplement for specific operation on your aircraft). Autopilot annunciator display selected modes and faults are shown in (see Table 5B-5).Through the turn knob and vertical control the pilot can manually “dialin” a desired roll or pitch angle. Each control initiates an aircraft attitude change in proportion to and in the direction of control movement.The control has a center detent position at wings level position and it is not spring loaded. The control will remain at any position between the end stops when released. Operation of the control cancels any previously selected lateral modes except the APPR mode.

OTTc: When engaging the system, do not push the YAW ENG and then the AP ENG buttons in rapid succession. If the autopilot engage occurs in the middle of the yaw pre-engage test, the autopilot will fail and engagement will not occur.

OTTc: The turn knob becomes inactive if it is out of the detent position when the SYNC button is pushed.

Annunciator Indication Color AP Autopilot is engaged. Green DIS (AP) Autopilot is disengaged. DIS portion of annunciator will flash for 5

seconds, then become steady. DIS also illuminated when autopilot is engaged and the SYNC button is pressed.

Amber

YAW Yaw damper is engaged. Green DIS (YAW) Yaw damper is disengaged. DIS portion of annunciator will flash for 5

seconds, then become steady. DIS also illuminated when yaw damper is engaged and the SYNC button is pressed.

Amber

SR Selection of soft-ride (turbulence) mode. Green 1/2 ø Selection of half-bank mode. Green T Trim servo monitor. Amber E Elevator servo monitor. Amber A Aileron servo monitor. Amber R Rudder servo monitor. Amber AP Autopilot failure. Red TRIM Trim system failure. Red TRIM Trim system operating (trim in motion). *Amber orWhite

*Color of TRIM annunciator depends upon the CPN of the APP-85A.Table 5B-5: Autopilot Annunciator Display Selected Modes and Faults

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Early versions of the APC will not allow the system to engage if the turn knob is out of detent. Later versions will allow the system to engage but the turn knob is inactive until after it has been returned to the center detent position. Refer to the Airplane Flight Manual Supplement for specific operation in your aircraft. The UP/DN vertical control is a center-off, spring-loaded rocker switch that provides manual control of the elevator channel when the autopilot is engaged and no vertical modes are selected. The vertical control also provides incremental changes to the air data reference if any vertical mode except DSC (descent) mode is selected in systems.The soft-ride or turbulence button commands the system to reduce autopilot gains. Rather than react to every disturbance of aircraft attitude due to flight in turbulent air, the system reduces the control responses. This increases passenger comfort by reducing the continuous control movements the autopilot would command in turbulence.The soft-ride mode limits the maximum commanded bank angle to approximately 12.5°. Mode can be selected with any lateral mode except APPR mode or when the APP turn knob is used. 1/2 1/1 mode is automatically cancelled when the VOR or localizer signal is captured during the NAV or APPR modes. However, 1/2 1/1 mode can be reselected after capture but will have no effect since these modes have a track bank limit of 10°.An autopilot disconnect switch is normally located on each control wheel to provide a convenient means to disengage both the autopilot and the yaw damper. Operation of the disconnect switch causes all autopilot servos to disengage and all selected vertical modes to cancel if no lateral modes are selected. The vertical modes can be reselected after the autopilot has disengaged and a lateral mode has been selected. Any selected lateral and vertical modes are retained and can be used in the flight director mode.The SYNC button, located on the control wheel, is used to manually maneuver the aircraft without completely disengaging the autopilot. Pressing the SYNC button disengages the primary servos and may cancel any selected vertical modes without interrupting computer computations for the selected modes. (Canceling the vertical modes with SYNC is optional in some APCs. Refer to the Flight Manual Supplement for the particular operation in your aircraft.) The airplane can be maneuvered to any desired attitude while the SYNC button is pressed. When the SYNC button is released, the primary servos reengage, the APC synchronizes to the new attitude and maintains it, and control is returned to the previously selected modes. Return to the selected modes is restrained to prevent any rapid maneuvers when the aircraft attitude varies from that commanded by the autopilot.After glideslope capture in APPR mode, pressing the SYNC button has no effect on vertical commands (will not cancel the vertical mode). However, the SYNC button does still select the SYNC mode if the autopilot is engaged. When the SYNC button is released, commands are generated to return the airplane to the center of the glideslope.A GO-AROUND switch is located on the left throttle handle. The altitude preselector/alerter panel selects and displays altitude.

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Auaoptloa CoepuahrThe Autopilot Computer is a Flight Guidance computer with four servo control circuit cards installed; one each for roll, pitch, yaw axis, and one for elevator trim. (The yaw channel provides only yaw damper capabilities; it does not provide rudder trim.) This single computer concept allows all of the modes of operation in the flight director to be flown automatically by the autopilot. In addition to the inputs used by the FGC, the APC receives inputs from the autopilot panel, yaw rate sensor, and an optional slip/skid sensor. These inputs are processed by the APC to provide outputs to drive the autopilot servos as well as the command bars or crosspointers in the electronic flight instrument system.

hrvosThe four servos (roll, pitch, yaw, and trim) position the airplane control surfaces in response to commands from the autopilot computer.

EFIS

OvhrvthwThe EFIS-85B Electronic Flight Instrument Systems combine electronic ADI/HSIs with a center panel Multifunction Display (MFD) to provide a fully integrated flight deck display system. The electronic ADI has a 50% larger earth/sky representation, terminal and enroute format than the conventional five inch electromechanical instrument. The electronic HSI combines multiple display formats with weather radar and built-in NAV source selection for simplified coordination of all navigation and weather avoidance tasks. The MFD augments the electronic ADI/HSI displays by providing the pilot with additional navigation, weather avoidance, traffic warning, data management and reversionary functions.The only physical difference between the EFIS display is the addition of the inclinometer on the bottom of the EADI.

DhsSrtpatonThe EFIS system is comprised of a family of components that can be configured to meet various flight deck display applications. The system components are:

Dtsplmy hlhSa PmnhlProvides pilot and copilot inputs of selected heading and course for the EFIS and provide control for the various formats and functions for the EHSI and EADI. Format switch (FORMAT) – Select the EHSI display (ROSE,

APPR, and ENR). Range knob (RNG) – Controls the range of the EHSI when APPR or

ENR is selected.

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Course Switch (CRS) – Select the navigation sensor to be connected to the EHSI active course arrow.

Bearing Switch (BRG) – Select the nav sensor to be connected to the EHSI bearing pointer.

Decision Height Set Knob (DH SET) – Set the decision height on the EADI.

Radio Test Altimeter Button (TST) – Initiate the test function for the radio altimeter.

Navigation Data Button (NAV DTA) – Toggle information between time-to go, ground speed, elapsed time, or wind information displayed on the upper right corner of the EHSI.

Second Course Button (2ND CRS) – Allow to display information from a second nav sensor on the EHSI.

Course Control Button (CRS CTL) – Allow the course knob to control the active selected or the preset second course arrow.

Weather Button (WX) – Press button to display weather information on the EHSI when APPR or ENR format is selected.

Heading Select Knob (HDG) Position the cursor for the on-side EHSI. Heading Sync Button (PUSH HDG SYNC) – When pressed, causes

the heading cursor to match the aircraft heading. Course Select Knob (CRS) – Position the active selected course

arrow for the on-side EHSI. Course Direct To Button (PUSH CRS DIRECT) – When pressed with

VOR as the nav sensor, the selected course and course arrow rotate towards the station until VOR deviation is zero.

Dtsplmy ProShssor UntaProcesses sensor and control inputs to provide necessary deflection and video signals for up to two EFD. A single DPU is able to drive the EHSI and EADI. Software may vary between aircraft; consult the AFM supplement for latest software version and feature.

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MulatfunSaton ProShssor UntaProcesses information from sensors and control inputs to provide necessary deflection and video signals for the MFD. The MPU can also drive the EHSI and EADI displays using cross-side data reversion mode in the event of a DPU failure. Software may vary between aircraft; consult the AFM supplement for latest software version and feature.

Figure 5B-33: Multifunction Processor Unit

TADIThe EADI display provides an ADI presentation that can include the following (with defined sensor inputs): pitch and roll attitude flight control system mode annunciation lateral deviation (VOR, localizer, or LNAV) vertical deviation (glideslope, MPG, or VNAV) autopilot engage annunciation attitude comparator warnings decision height and set annunciation.

Figure 5B-34: EADI Display

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In addition, an EADI shows pilot selected information for specific functions or operating mode such as: flight control system steering commands altitude sensor source annunciation marker beacon annunciation indicated airspeed with reference airspeed and airspeed trend vector fast/slow or angle-of-attack deviation mach speed altitude deviation radio altitude approach mode comparator warnings Cat II excessive deviation for ILS or MLS and airspeed.

EHSIThe EHSI displays present navigation information in three different display formats: ROSE (full compass rose), APPR (approach) and ENR (enroute). ROSE mode presents compass rose similar to existing electromechanical instruments, APPR displays a segment or sector of an expanded view of the compass, and ENR displays a map-like presentation of the horizontal navigation situation within a compass arc. Selection of the format is done with the format switch on the EFIS control panel.

Figure 5B-35: EHSI Displays

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Rose Format Displays and Annunciators bearing pointer LNAV annunciation compass display and heading sensor type selected heading active selected course to/from arrow TGT and TRB weather annunciation data preset course vertical deviation lateral deviation LIN/ANG/XTK and B/C annunciation waypoint or DME identifier and waypoint alert distance APPR format display and annunciator compass sector display selected heading display active selected course display bearing pointer display vertical deviation display lateral deviation bar and scale display range display to/from display weather radar display preset course display peripheral display APPR format display and annunciator compass sector display active selected course display preset course display.

Warning Flags Distance – Displays invalid distance data detected from DME or LNAV. Heading – Heading system failure. DPU or MPU – Displays failureof the processor unit; if displayed

more then five seconds the display will go blank except for the DPU or MPU flag.

Data – Invalid time to go or ground speed. Vertical Deviation – Failure of the glideslope receiver or LNAV when

in FGS or PGS mode. Bearing pointer – Bearing pointer sensor failure. Cross-Side data – Indicates failure of the cross-side data bus.

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Display Select panel – Control panel failure. Active selected course – Navigation sensor failure while selected for

the active course. Preset course flag – Navigation sensor while it is selected for the

preset course.

MulatfunSaton DtsplmyProvides a multicolor CRT instrument display of weather radar, pictorial navigation map, page data, waypoint definition, display of multiple waypoints and remote data.

TCASDisplays a pictorial presentation of traffic around the aircraft. The traffic’s type, range, bearing, vertical speed, direction of vertical speed and relative or absolute altitude is displayed in this format.

Weather Radar Weather data – Displays weather information from the turbulence

weather radar. Mode, range and other parameters selection is done on the TWR-850

control panel.

Navigation InformationFour display formats can be used: heading up; north up with aircraft centered; north up max view; and map plan.

Remote DataDisplays remote data for the selected source and extended data.

Extended data:Contains pages of data useful to pilots and maintenance personnel: VOR/DME CRUISE APPROACH CAT II DIAGNOSTICS.

Page and Emergency ModesContains 100 pages of user information may be stored in the nonvolatile memory in the MFD. Pages may be modified by the user and can be arranged in chapter format. Pages can be used for emergency or data and can be divided into any ratio for a maximum of 100. The top line of each page is reserved for weather radar mode and/or target alert annunciators.

Additional Function Waypoint definition – Permits selection of a waypoint with the cursor

and transfer data for that waypoint to a compatible LNAV system.

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Remote data – Displays LNAV navigation information transfer to the EFIS, displayed in remote format.

Warning FlagsNAV – Navigation sensor failure, the annunciator change to for the nav sensor change color to red.Heading – Displayed if there is a failure of the heading sensor.Heading Comparator – Displayed when the MPU detects a heading comparison error.MFD Controls Power Button (PWR) – Turns on the MFD. Intensity Control (INT) – Controls MFD display brightness. Traffic Button (TFC) – Turns on the traffic display mode Radar and Navigation Button (RDR/NAV) – Turns on or off the radar

and navigation display modes. Remote Button (RMT) – Shows remote data selectable from the line

keys. Page (PGE) and Emergency (EMG) buttons – Displays page or

emergency data information. Display Line Keys – selects options data adjacent to the line keys. Joystick – Used to select waypoint or slew through data pages. Data Jack – Input for remote programming. Line Reverse Button (_) – Changes display range or moves cursor

up in PGE or EMG modes. Line Advance Button (_) – Changes display range or moves cursor

down in PGE or EMG modes. Recall Button (RCL) – View previous lines in PGE or EMG modes. Skip Button (SKP) – Skip lines in PGE or EMG modes. Clear Button (CLR) – Reset all lines selected in PGE or EMG

modes.

TCAS

DhsSrtpatonTCAS is designed to protect the airspace around the aircraft, using modes C and S signal of nearby aircraft and display surrounding traffic bearing, altitude and rate of climb. Aural and visual indication is given for all Resolution Advisory (RA) and Traffic Advisory (TA) if the traffic is within 15 to 48 seconds of impact. The system consists of receiver/transmitter, control, directional antenna, VSI/TA/RA indicator, MFD display, mode S transponder and ATC control panel. Traffic can be displayed on the VSI and on the MFD.The TCAS computer analyzes traffic information and divides the traffic into four categories: other traffic, proximity traffic, traffic advisory and resolution advisory.

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RA resolution is given in a form of vertical maneuver to provide pilot with proper evasion maneuvers (green display shows proper vertical speed to be selected and red must be avoided.If the other TCAS equipped aircraft has the same equipment, the steering command will be coordinated prior to RA being issued, to be least disruptive to maneuver.

Advtsory TyphOther Traffic – Non-threatening traffic, display range and altitude.Proximity Traffic – Non-threatening aircraft traffic within 1,200 ft relative altitude or 6 miles radius. Indication display range and altitude, this indication is used by the pilot to take the appropriate TA or RA evasive maneuvers.Traffic Advisory – Non-threatening traffic, display to aid the pilot acquire surrounding traffic. This traffic type may be upgraded to RA depending on own and other aircraft flight path. Reaction time to Close Point Of Approach (CPA) is given in seconds (see Table 5B-6).Resolution Advisory – Threatening traffic, display range and altitude and issue an aural and visual advisory to resolve threat situation. There are two types of RA: Preventive: No evasive action is required, but issue an aural command

to “monitor vertical speed” and display range to be avoided. Corrective: The computer determines what action the pilot should

take to avoid the threat traffic. The pilot must follow the aural annunciation and visual warning to avoid collision. Reaction time to CPA is given in seconds (see Table 5B-7).

The self-test routine tests the receiver/transmitter, mode S transponder, TCAS antennas, radio altimeter validity, heading data validity and TCAS display. If TCAS failed during self-testan aural warning will be annunciated and displayed on the VSI.

ATC Conarol Function Knob – Selects mode of operation. ATC Code Set Knob – Sets the aircraft squawk codes. Mode S/Mode A/C Switch –Selects TCAS or ATC mode. IDENT Button – Transmits an identifier code. PRE Button – Selects a stored ATC squawk code. TEST Button – Initiates self-test of the TCAS-94 (including transponder).

The transponder control displays ATC squawk codes, encoded altitude, annunciators and diagnostic codes. A function knob sets the operating mode; ATC dial selects ATC codes; PRE button selects stored ATC squawk codes; IDENT button transmits the aircraft identifier code.

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Altitude Seconds to CPAUp to 1,000 20 1,000-2,350 25 2,350-5,000 30 5,000-10,000 40

10,000-20,000 45 20,000-35,000 48

Table 5B-6: Traffic Advisory Response Time

Altitude Seconds to CPA1,000-2,350 15 2,350-5,000 20 5,000-10,000 25

10,000-20,000 30 20,000-35,000 35

Table 5B-7: lResolution Advisory Response Time

Enhanced Ground Proximity Warning System (EGPWS) The EGPWS provides the crew with aural and visual warning if projected flight path could result in impact with the terrain. The system consists of a computer, warning lights and switches. The system utilizes aircraft sensor input (ADC, radio altimeter, compass system, GPS, flap and gear switch, attitude information and glideslope deviation) and computes it to provide warnings for: excessive rate of descent terrain closure rate alert to descend after takeoff insufficient terrain clearance descent below glideslope advisory callouts windshear detection.

Operation of the EGPWS is automatic when avionics power switches are in the ON position and all related systems are valid. The system is operational from altitudes of 50 to 2,000 ft (4.57 to 609.6 meters) above ground level as determined by the radio altimeter. The EGPWS has seven operational modes all of which are valid for ground proximity warning. Visual warnings are indicated on instrument panel mounted annunciators and on the EHSI, and MFD. Voice warning messages are announced on the aural warning system.

Mode 1Excessive Sink Rate. Sink rate envelope is measured barometrically and registers in a flight envelope beginning at approximately 2,500 ft per minute at 2,500 ft above ground level. The warning will go on between 1,100 and 700 ft AGL. If this flight envelope is entered, the aural warning, SINK RATE, will be heard. Warning will stop after sync rate is reduce or climb is initiated.

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Mode 2Excessive Terrain Closure Rate. Terrain closure rate, during cruise operation (flaps up), is sensed by the radio altimeter. The aural message is, TERRAIN, TERRAIN, followed by, PULL UP, if the airplane remains in the mode envelope. Warning will stop after a climb or 45 seconds after overflowing the peak.

Mode 3Altitude Loss After Takeoff. After takeoff, a rate-of-climb for a specific altitude loss will trigger an aural warning of, DON’T SINK. The amount of altitude loss varies from 25 at 50 ft at 300 ft AGL (100 to 150 ft for aircraft at 1,200 ft AGL). Warning stops once a positive rate of climb has been established.

Mode 4Proximity To Terrain and Airplane Not In Landing Configuration. There are three conditions and messages in this mode. If the airplane descends below 500 ft above ground level at a speed less than 150 Kts and the landing gear is not lowered, an aural warning of, TOO LOW – GEAR, will be repeated every 0.75 seconds, until the situation is corrected or the airplane is flown out of the envelope. If the airplane descends below 245 ft ground level at a speed of less that 150 Kts, the landing gear is down, and the flaps are not in landing position, an aural warning TOO LOW – FLAPS will be heard. The PULL UP – FLAPS aural warning may be deactivated by placing the GPWS NORM/GPWS FLAP OVRD switch in the GPWS FLAP OVRD position. While in a descend of 500 FPM at a speed of approximately 250 Kts, at 900-1,100 ft AGL increasing linearly, a warning TOO LOW – TERRAIN will be heard.

Mode 5Inadvertent Descent Below Glideslope. Repeated aural warnings of GLIDESLOPE and illumination of the amber BELOW G/S annunciator are initiated, if the airplane descends slightly more than 1.3 dot below the instrument landing system glideslope. This soft warning can be silenced by returning to the glideslope. When the airplane is more than two dots below the glideslope and is between 300 ft (91.44 meters) and 150 ft (45.72 meters) above ground level, the warning becomes hard as evidenced by the warning GLIDESLOPE being repeated louder and faster. The hard warning can be silenced only by a positive pull up of the aircraft. Mode 5 operation can be inhibited by the BELOW G/S annunciator, while in the soft warning area. Climbing to a radio altitude of above 1,000 ft (304.80 meters) or descending below 50 ft (15.24 meters) will reset Mode 5 if it was cancelled. If the pilot desires to purposely descend below the glideslope, Mode 5 operation can be inhibited. If the BELOW G/S annunciator is depressed at an altitude below 1,000 ft (304.80 meters) above ground level and still in the soft mode of operation.

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Mode 6Altitude Callouts. A “FIVE HUNDRED” audible message, calling altitude above the DH altitude, is repeated twice, when the airplane is below 1,000 ft AGL above ground level and the radio altimeter passed through the altitude set in the radio altimeter DH window. No other system warning is provided. However, the DH annunciator on the pilot’s attitude director indicator will also illuminate from inputs received from the radio altimeter.Bank Angle. A “BANK ANGLE” audible message, calling for excessive bank angle, is repeated twice.

Mode 7Windshear Detection. Windshear system calculates both the rotational rate and force of the acceleration vector, alerting the pilot to apply go-around power and increase the angle-of-attack to exit the potentially hazardous situation.

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Avionics Related Annunciators

Auaoptloa AnnunStmaors AP – autopilot failed TRIM – trim system failed

TFI AnnunStmaors EFD/DPU FAN FAIL MFD/MPU FAN FAIL

Globml Postatontng ysahe GPS INTEG – system integrity failure.


PreflightThe FMS, EFIS, TCAS, and EGPWS should be functionally checked as part of the daily inspection in accordance with the associated pilot’s guide.During the exterior preflight inspection, check the condition of the pitot masts and static ports. All should be clean and free from obstructions. Check the security and condition of the antennas on the underside of the fuselage. Visually inspect the antennas on the fuselage top and vertical stabilizer. Check the static wicks for security and condition.Check that the ELT is armed and that its antenna is secure (if visible). If suspected, drain moisture from the static lines with the static system drain valves on the right side of the cockpit. Preflight procedures for the autopilot vary with the system installed. Always refer to the applicable AFM Supplement. Generally, turn the battery and avionics master switches on. Engage the autopilot and yaw damper. Check that control wheel and rudder pedal movement overpowers the autopilot; there should be resistance to control movement. Check that the control wheel disconnect switch disconnects the autopilot and yaw damper.

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Abnormal and Emergency Procedures

TehrghnSy ProShdurhsEmergency procedures for the pitot/static system and autopilot are limited. They include: pilot’s static air system autopilot malfunction or failure.

Ptloa’s amatS Atr ysaheIf the pilot’s and copilot’s airspeed indicator, vertical speed indicator, and/or altimeter disagree or the pilot’s vertical speed indicator appears sluggish, there may be a partial or complete obstruction of the pilot’s static ports. Move the PILOT’S STATIC AIR SOURCE switch to ALTERNATE and observe instrument indications. If this procedure corrects the problem, the pilot’s primary static ports are obstructed. With ALTERNATE selected, the pilot’s airspeed indicator and altimeter are in error because of the location of the pilot’s alternate static port. Use the Airspeed Calibration and Altimeter Correction charts in the AFM Performance Section to correct the indications.

Auaoptloa MmlfunSatonNormally, if the autopilot malfunctions or fails, the autopilot automatically disconnects. If the autopilot fails to disconnect, disconnect the autopilot by: pressing the control wheel disconnect switch pressing the go-around button pressing the TEST button on the controller pulling the autopilot circuit breaker turning off the avionics master switch.

Typically, most autopilot systems disconnect if there is a power interruption or failure.If AP TRIM and MASTER WARNING annunciators illuminate with the autopilot engaged, immediately disconnect the autopilot while restraining the control wheel; there may be a severe nose down pitching movement. During autopilot malfunction, there may be an altitude loss before regaining positive control of the aircraft.

Electrical System 5C



ContentsElectrical

SchematSc: DC Electrical System (King Air 200) .........................5C-5SchematSc: DC Electrical System ...............................................5C-7SchematSc: DC Electrical System (King Air B200) .......................5C-9SchematSc: DC Electrical System .............................................. 5C-11SchematSc: DC Electrical System (With EFIS) ........................... 5C-13SchematSc: DC Electrical System (With EFIS) ........................... 5C-15SchematSc: Circuit Breaker Panels ............................................. 5C-17

DC Electrical SystemBattery .................................................................................................... 5C-19

Battery Switch .................................................................................... 5C-20Thermal Runaway .............................................................................. 5C-21Battery Monitor System ..................................................................... 5C-21

Starter/Generators ................................................................................. 5C-22Generator Switches ........................................................................... 5C-23Optional Generator Overheat Monitoring System ............................. 5C-23Volt/Loadmeter .................................................................................. 5C-23

Generator Voltage Regulation Systems .............................................. 5C-24Voltage Regulation System ............................................................... 5C-24Generator Control Panels .................................................................. 5C-25

External Power ....................................................................................... 5C-26DC Distribution System ........................................................................ 5C-28

Hot Battery Bus .................................................................................. 5C-28Main Battery Bus ............................................................................... 5C-29Isolation Bus ...................................................................................... 5C-29Generator Buses ................................................................................ 5C-29Dual-Fed Buses ................................................................................. 5C-30DC Avionics Buses ............................................................................ 5C-31Avionics Master Power ...................................................................... 5C-31

Circuit Breaker System ......................................................................... 5C-32Engine Starting and Ignition............................................................... 5C-33Auto Ignition ....................................................................................... 5C-35

SchematSc: AC Electrical System ............................................... 5C-36



AC Electrical SystemAC Inverters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-37INVERTER Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-38Annunciators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-38AC Voltmeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-39Inverter Circuit Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-39

Preflight and ProceduresPreflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-41Abnormal Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-41BATTERY CHARGE Annunciator Illuminated on the Ground . . . . . 5C-41BATTERY CHARGE Annunciator Illuminated in Flight . . . . . . . . . . 5C-42Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-42Electrical System Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-42

Generator Inoperative (DC GEN Annunciator ON) ...........................5C-42Current Limiter Check With One Generator Off Line . . . . . . . . . . . . 5C-43Bus Feeder Circuit Breaker Tripped in Flight . . . . . . . . . . . . . . . . . . 5C-43Inverter Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-43

Lighting SystemsLighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-45Interior Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-45Cockpit Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-45Cabin Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-46Exterior Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-47Power Sources . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-48

Data SummariesElectrical System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-49DC System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-49AC System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5C-50

Electrical System



ElectricalThis section details the electrical and lighting systems of the Super King Air 200/B200 series.The aircraft 28 VDC electrical system consists of a primary DC system that powers most of the aircraft’s electrical components through a Dual-Fed bus system and a secondary AC system that powers avionics, yaw rate, and some engine instruments. The negative lead of each power source is grounded to the main aircraft structure.The standard DC electrical system power sources are one 24V, 34 amp-hour battery, two 250A, 28.25 VDC starter/generators, and an external power unit. These sources supply power to their respective Hot Battery and Generator buses.Two Generator buses supply DC power to their respective inverters, which provide 400 Hz, 115 VAC and 26 VAC to equipment and systems requiring alternating current.The aircraft external power circuitry can accept from an external power source 400 Amps continuously and up to 1,000 Amps in current surges of short duration (0.1 second).The aircraft lighting system includes pilot controlled interior and exterior lights. The passengers can also control individual cabin lights.





5C-5For Training Purposes Only

Electrical System

DC Electrical System King Air 200

CAE SimuFlite

DC Electrical SystemKing Air 200

L FUEL FLOW BB225 AND AFTER

LEFT PITOT HEAT SW

BEACON LIGHTS SW

AVIONICS MASTER CONTROL

LEFT OIL TEMP INDICATOR

LEFT ICE VANE

PROP SYNCHROPHASER

CABIN PRESSURE CONTROL

FLIGHT INSTRUMENT LIGHTS

OVERHEAD AND SIDE PANEL LIGHTS

LEFT BLEED AIR WARNING

ANNUNCIATOR POWER

PNEUMATIC SURFACE DEICE

PILOT’S TURN AND SLIP INDICATOR

NO. 1 INVERTER CONTROL

FUEL DRAIN COLLECTER PUMPS

LEFT LANDING LIGHT SW

PROP DEICE AUTO HEAT SW

STROBE LIGHTS

LEFT OIL PRESSURE INDICATOR

LEFT ENGINE FUEL CONTROL HEAT

FIRE DETECTION

LEFT BLEED AIR CONTROL

LOGO LIGHT

INSTRUMENT DIRECT LIGHT

STALL WARNING SYSTEM

LANDING GEAR WARNING HORN

LEFT FUEL VENT HEATER

OPTIONAL ALTIMETER

TRIM TAB

LEFT GEN CONTROL

RIGHT OIL PRESSURE INDICATOR

RIGHT ENGINE FUEL CONTROL HEAT

CHIP DETECTOR

RIGHT BLEED AIR CONTROL

RADIO & ENGINE INSTRUMENT LIGHTS

CABIN FASTEN SEAT BELT & NO SMOKING

SIGN AND CHIMES

LANDING GEAR POSITION INDICATOR

RIGHT FUEL VENT HEATER

COPILOT’S TURN & SLIP INDICATOR

RUDDER BOOST SYSTEM

RIGHT GEN CONTROL

RIGHT OIL TEMP INDICATOR

RIGHT ICE VANE

AUTOFEATHER

CABIN TEMP CONTROL

SUBPANEL AND CONSOLE LIGHTS

RIGHTBLEED AIR WARNING

ANNUNCIATOR INDICATOR

WINDSHIELD WIPER

VACUUM DAMPER SYSTEM

NO. 2 INVERTER CONTROL

LEFT FUEL QUANTITY

LEFT FIREWALL VALVE

LEFT STANDBY FUEL PUMP

LEFT AUX FUEL QUANTITY

WARNING AND TRANSFER

LEFT FUEL PRESSURE WARNING

LEFT STARTER CONTROL

LEFT IGNITOR POWER

FLAP CONTROL & INDICATOR

FLAP MOTOR

LEFT MANUAL PROP DEICE

STALL WARNING HEAT

LANDING GEAR CONTROL

ICE LIGHT SW

RIGHT LANDING LIGHT SW

MANUAL PROP DEICE CONTROL

RIGHT MANUAL PROP DEICE

RIGHT FIREWALL VALVE

RIGHT STANDBY FUEL PUMP

RIGHT FUEL PRESSURE WARNING

RIGHT FUEL QUANTITY

TAXI LIGHT SW

NAV LIGHT SW

RIGHT PITOT HEAT SW

PROPELLER GOVERNOR

RIGHT IGNITOR POWER

RIGHT STARTER CONTROL

RIGHT AUX FUEL QUANTITY

WARNING AND TRANSFER

FUEL CROSSFEED

0

20 40 60

80

100 0

20 30

10

DC LOAD

0

20 40 60

80

100 0

20 30

10

DC LOAD

REVERSE CURRENT

PROTECTION

LEFT GEN CONTROL

+ -LEFT START RELAY

REVERSE CURRENT

PROTECTION

RIGHT GEN CONTROL

RIGHT START RELAY

EXTERNAL POWER PLUG-ENGAGED

SENSOR

EXT PWR

+ -ISOLATION LIMITER 325A

RIGHT FIRE EXT

RIGHT FIREWALL SHUTOFF VALVE

LEFT FIREWALL SHUTOFF VALVE

THRESHOLD LT


LEFT STANDBY FUEL PUMP

+- +

+-

60A 50A

60A 50A

60A 50A

60A 50A

60A 50A

60A 50A60A

50A SHUNT

EXT POWER CONNECTION ISOLATION LIMITER 325A

EXT POWER RELAY

BATTERY RELAY

BATTERY SW

BATTERY RELAY FUSEBB364 AND AFTER

BATTER

SHUNT

BATTERY CHARGE SENSOR

RIGHT STARTER GEN

LEFT STARTER GEN

SHUNT

LEFT FIRE EXT

LEFT GENERATOR BUS

RIGHT GENERATOR BUS

N O 4

D U A L

F E D

B U S

N O 3

D U A L

F E D

B U S

N O 2

D U A L

F E D

B U S

N O 1

D U A L

F E D

B U S

I S O L A T I O N

B U S

M A I N

B A T T E R Y

B U S

H O T

B A T T E R Y

B U S

OVER VOLT PROT

BB364 AND

AFTER 5A

YAW DAMP

60A 50A 15A

15A

5A

5A 5A

5A 5A

5A

5A

5A

5A

5A

15A 15A

R FUEL FLOW BB225 AND AFTER

RECOGNITION LIGHT SW

5C-6 Developed for Training Purposes King Air 200October 2001


5C-6 For Training Purposes Only




Electrical System

DC Electrical System King Air 200

VENT BLOWER POWER

AFT EVAPORATOR BLOWER POWER

AIR CONDITIONER CLUTCH

COPILOT’S WINDSHIELD

ANTI-ICE

RIGHT RADIANT HEAT

DC TEST JACK

LANDING GEAR MOTOR CONTROLLER AND DYNAMIC BRAKE

PILOT’S WINDSHIELD

ANTI-ICE

LEFT RADIANT HEAT

CONDENSER BLOWER

MOD OPTIONAL EQUIPMENT

AVIONICS NO. 2 POWER RELAY



AVIONICS MASTER

CONTROL

ON

OFF

AVIONICS SWITCH

60A

40A

40A

5A

60A

LDG GEAR CONTROL

FROM NO. 2 DUAL FED BUS

LANDING GEAR

MOTOR

30A 30A

30A

5A

TO AC POWER

TO AC POWER

FROM NO.1 DUAL FED BUS

50A

30A

7.5A

50A

20A

15A

60A 50A

40A 50A

50A

20A

AVIONICS BUS NO. 1

AVIONICS BUS NO. 2

AVIONICS BUS NO. 3

RIGHT GENERATOR BUS

LEFT GENERATOR BUS

King Air 200 Developed for Training Purposes 5C-7 October 2001






Electrical System

DC Electrical System King Air B200

CAE SimuFlite

DC Electrical System King Air B200

RIGHT GEN 40 60 80 RIGHT GENERATOR BUSRIGHT CONTROL

20

10 20 100 300

0

STARTER GEN

50A15A 50A 50ADC LOAD 60A 50A 60A 60A15A 60A ISOLATION L EFT LANDINGREVERSE LEFT PITOTLIMITER LIGHT + -CURRENT MANUAL PROPHEAT325A TAXI LIGHTPROTECTION RIGHT LANDING DEICE CONTROLRIGHT START RELAYSHUNT SWITCHPROP AUTOMATIC LIGHT SWITCH LEFT MANUAL HEAT SWITCH PROP DEICE RIGHT MANUAL

PROP DEICEICE LIGHTSEXT PWR NAV LIGHTS FLAP MOTORBEACON LIGHTSEXTERNAL POWER PROPELLERSWITCH LANDING GEARRECOGNITIONPLUG-ENGAGED GOVERNORCONTROLSENSOR LIGHTS FLAP CONTROLSTROBE LIGHTS TAIL FLOOD & INDICATORSWITCH RIGHT IGNITORSHUNT LIGHTS SWITCH STALL WARNING POWER

RIGHT STARTER

RIGHT PITOT HEAT HEAT LEFT IGNITOR POWER

I

MA

N

BATTERY

BATTER CONTROLBATTERY BATTERY SW AVIONICS LEFT STARTER MASTER CONTROL

LEFT FUEL FLOW RIGHT FUEL CONTROL FLOW INDICATOR RIGHT OIL PRESS

LEFT OIL PRESSURE INDICATOR

CHARGE SENSOR 5AR NAV

MEMORY (OPT) 5A

BAT

5A

LEFT OIL PRESSURE WARNING (OPT.) RIGHT OIL PRESS RIGHT OIL TEMP N

O4

I NO3

LEFT ICE VANE WARNING (OPT.) LEFT FUELTEMPERATURE PRESSURECONT LEFT ENGINE FUEL WARNINGCONT HEAT RIGHT ENGINE FUEL

SOLAT

R ICE VANE CONTLEFT GEN CONTROL HEATSTEREO (OPT.)

ENTRY LTS & 5A

HOT

OVERHEAT (OPT.) LEFT FUELFIRE DETECTION DUAL

RIGHT FIREWALL VALVE

DUAL

RIGHT FIRE RIGHT GEN QUANTITYPROP DETECTION OVERHEAT (OPT.)BATTERY I ON

SYNCHROPHASER LEFT CHIPCLOCK LTS & RELAY NO1

NO2

DETECTOR LEFT AUX FUEL AUTOFEATHER QUANTITY

EXT PWR SENSE BUS

RIGHT CHIPFLIGHT & GYRO5A BATTERY

WARNING ANDDETECTORINSTRUMENT LIGHTS INSTRUMENTRIGHT STANDBY TRANSFERAVIONICS & ENGINE5A INDIRECT LIGHTSFUEL PUMP INSTRUMENT LTS FED

BUS

FED


OVERHEAD AND CABIN READING2 5A SIDE PANEL LIGHTS LIGHTSSTALL WARNING SYSTEM

DUAL

DUAL

COPILOT’S TURN & SLIP

LEFT STANDBY 5A LEFT BLEEDFUEL PUMP RIGHT AUX FUEL QUANTITY

EXT POWER OVERHEAD,2 AIR WARNING LANDING GEAR SUBPANEL5A RELAY WARNING HORN FLOURESCENT LTS WARNING ANDAND BUS

THRESHOLD LT BUS

ANNUNCIATOR TRANSFERNO SMK, FSB, AND+-BUS

POWER LEFT FUEL CABIN LEFT STANDBYRIGHT BLEED AIR VENT HEATER5A FUEL PUMPLEFT FIREWALL WARNINGFED

FED

RIGHT FUELPNEUMATIC SURFACE DEICE

SHUTOFF VALVE QUANTITYLANDING GEAR POSITION IND BRAKE DEICE ANNUNCIATOR5A (OPTIONAL)RIGHT FIREWALL INDICATOR RIGHT FUELAUTOMATICSHUTOFF VALVE PRESSUREOXYGEN CONTROL RIGHT FUELOVER WARNINGYAW DAMPER5A WINDSHIELD WIPER LEFT FIREWALLVENT HEATB

US

BUS

VOLTLEFT FIRE EXT LEFT BLEED AIR PROT CONTROL 5A VALVE

FUEL CROSSFEED

PILOT’S TURN & SLIP CABIN TEMP CABIN PRESSURERIGHT FIRE EXT 5A ENCODER CONTROL LOSS (OPT.)

ALTIMETER (OPT.) NO. 1 INVERTER + CONTROL NO. 2 INVERTER RIGHT BLEED AIR

PITCH TRIM CONTROL CONTROLISOLATION +-1

EXT POWER LEFT OILLIMITERCONNECTION LEFT GEN CONTROL TEMPERATURE RUDDER BOOSTFURNISHINGS325A CONTROLMASTER CONTROL CABIN PRESS

CONTROL RIGHT GENERATORSHUNT LEFT START RELAY 50A CONTROLREVERSE 50A50A 50A + - 60ACURRENT 60A60A CIGARETTE 60APROTECTION LIGHTER 40 6015A

8015A 20

10 20 100 0

0LEFT 30

STARTER LEFT GEN DC LOAD GEN CONTROL LEFT GENERATOR BUS

ON BB1096, 1098 AND SUBSEQUENT CIRCUIT 2 ON BB1097, 1095 AND PRIOR1 BREAKERS REPLACE FUSES STBY FUEL PUMPS ARE NO LONGER POWERED

5C-8 Developed for Training Purposes King Air 200October 2001






Electrical System

DC Electrical System King Air B200Electrical System CAE SimuFlite

DC Electrical System

(With EFIS)

VENT BLOWER POWER

AFT EVAPORATOR BLOWER POWER


COPILOT’S WINDSHIELD

ANTI-ICE

RIGHT RADIANT HEAT

DC TEST JACK


PILOT’S WINDSHIELD

ANTI-ICE

LEFT RADIANT HEAT

CONDENSER BLOWER





AVIONICS MASTER

CONTROL

ON

OFF

AVIONICS SWITCH

60A

40A

40A

5A

60A

LDG GEAR CONTROL


LANDING GEAR

MOTOR

30A 30A

30A

5A

TO AC POWER

TO AC POWER


50A

30A

7.5A

50A

20A

15A

60A 50A

40A 50A

50A

20A

AVIONICS BUS NO. 1

AVIONICS BUS NO. 2

AVIONICS BUS NO. 3

RIGHT GENERATOR BUS

LEFT GENERATOR BUS

King Air 200 Developed for Training Purposes 5C-9 5C-10 May 2006






Electrical System

DC Electrical System (With EFIS)

RIGHT GEN 40 80

60 RIGHT GENERATOR BUSRIGHT CONTROL 20

10 20 100 300

0

STARTER GEN

50A15A 50A 50ADC LOAD 60A 50A 60A 60A15A 60A ISOLATION REVERSE LIMITER OAT+ -CURRENT ALT ALERT MANUAL PROP COPILOT TURN & SLIP325A INVERTER CONT. NO. 2PROTECTION DEICE CONTROLRIGHT START RELAY SHUNT AURAL WARNING PILOT EADI

1 4

5 LEFT MANUAL CIGAR LIGHTERCOPILOT AUDIO PROP DEICE RIGHT MANUALAVIONICS PILOT EHSI PROP DEICE

PROPELLER

MASTER CONTROLEXT PWR FURNISHINGS MASTER CONTROL

CABIN AUDIO FLAP MOTOR PILOT TURN & SLIPEXTERNAL POWER 4 DISPLAY PRCSR

INSTRUMENT INDIRECT LIGHTS

4 1

PLUG-ENGAGED RIGHT MULTI GOVERNOR LANDING GEAR RELAYFCTN PRCSRSENSOR FLAP CONTROL & INDICATOR LEFT MUTLI FUNCTION4 RIGHT IGNITOR

PRCSR INSTRUMENT INDIRECTRIGHT GENERATOR POWERSHUNT 2 2CONTROL LIGHTSCABIN LIGHTS LEFT IGNITOR 1 INVERTER CONT. NO. 1 POWER RIGHT STARTER2 1MA

AVIONICS & ENGINE CABIN LIGHTSAUTOFEATHER CONTROLINSTRUMENT LIGHTSPILOT ALTM AIR DATA 3BATTERY LEFT STARTERBATTERY SW BATTERY 1AVIONICS & ENGINEIN

CONTROLRIGHT ENGINE FUELPILOT FLIGHT INSTRUMENTS LIGHTSCHARGE SENSOR CONTROL HEATPILOT AUDIO INSTRUMENT LIGHTS

CABIN READING LIGHTSRIGHT CHIP DETECTCVR COPILOT FLIGHT INSTRUMENT LIGHTS

LEFT GEN CONTROL

1BATTERY

NO4

I NO3

2RIGHT MAIN ENG COPILOT FLIGHT LEFT FUELSOLAT

ANTI-ICE INSTRUMENT LIGHTS21 PRESSURESIDE PANEL LIGHTS 35AR NAV WARNING LEFT CHIP DETECT MEMORY (OPT) BATTERY OVERHEAD 2 RIGHT STANDBY ENG SUBPANEL & CONSOLE

ANTI-ICEFLOOD LIGHTS OVERHEAD LIGHTSLEFT ENGINE FUELHOT

RELAY LEFT FUEL5A CONTROL HEAT DUAL

DUAL

RIGHT FIREWALL VALVEQUANTITYBAT RELAY BEACONS TAXI LIGHTSRIGHT ENG INSTR

POWER I 3 FIRE DETECTION ON

NO1

NO2

STROBES

TAIL FLOOD LIGHTS

ICE LIGHTS LEFT AUX FUEL QUANTITYENTRY LTS & LEFT MAIN ENG5A B

US

RIGHT FUEL FLOWANTI-ICECLOCK LTS & EXT PWR SENSE

THRESHOLD LT

BATTERY

NAV LIGHTS WARNING AND TRANSFER RIGHT OIL PRESSLEFT STANDBY ENG LEFT LANDING LIGHT RIGHT LANDING F

ED

BUS

FED


ANTI-ICEEXT POWER RELAY

5A LIGHT DUAL

DUAL

AVIONICS RIGHT OIL TEMP ANNUNCIATORS3 LEFT ENG

RECOG LIGHTINSTR POWER+ - RIGHTAUX FUEL35A RIGHT N1LEFT FIREWALL QUANTITYPROP AUTO HEATLEFT FUEL FLOW ANNUNCIATORSHUTOFF VALVE WARNING AND INDICATOR BUS

3 BUS

RIGHT N2 TRANSFERPROP SYNCLEFT OIL PRESS5ARIGHT FIREWALL LEFT STANDBYRIGHT BLEED AIR3SHUTOFF VALVE

LEFT FIRE EXT 5A

BUS

WARNING FUEL PUMPRIGHT TORQUEFED

FED

ANNUNCIATOR RIGHT FUELLEFT OIL TEMP POWER

AVIONICS

QUANTITY3 3OVER LANDING GEAR POS INDICATORS LEFT N1 RIGHT ITT

VOLT RIGHT FUEL ANNUNCIATORPROT5A 3 LEFT N2 PRESSURERIGHT BLEED AIR NO SMOKING & FASTEN 1RIGHT FIRE EXT WARNING CONTROL SEAT BELTS SIGNS3 LANDING GEAR LEFT FIREWALLBUS

BUS

5A LEFT TORQUE WARNING HORN

LEFT OIL PRESS

VALVE RIGHT OIL PRESS FUELCABIN TEMP CONTROL3 LEFT ITT WARNING CROSSFEED++ - WARNING

RUDDER BOOSTISOLATION LEFT BLEED AIR RIGHT FUELEXT POWER CONTROL

5

4

23

3

LIMITER CONTROL NO SMOKING & FASTEN SEAT BELTS SIGNS325A

VENT HEAT CONNECTION PILOT EADIAUTO OXYGEN WINDSHIELD WIPER

CONTROL LEFT STALL WARNING SYSTEM N1 STALL WARNING HEATSHUNT PILOT EHSILEFT START RELAY REVERSE PRESS CONTROL LEFT FUEL VENT HEATCURRENT + - RIGHT PITOT HEATCOPILOT ENCODINGPROTECTION PITCH TRIM ALTIMETER

PNEUMATIC 6015A 40

80 SURFACE DEICE 50A 50A15A 20

10 20 100 50A30 60A60ALEFT 0 0

BRAKE DEICE 50A60ASTARTER GEN

LEFT GEN 60ADC LOAD CONTROL LEFT PITOT HEAT

LEFT GENERATOR BUS BB-1439, BB-1444 thru BB-1448, BB-1450 thru BB-1457

4 EFIS-85 Installations BB-1449, BB-1458 thru BB-1462, BB-1464 thru BB-1485,

5

6

EFIS-84 Installations

BB--1802 thru BB-1842, except BB-1834 except BB-1484; BL-139 and BL-140 BB-1484, BB-1486 thru BB-1842, except BB-1834; BL-141 thru BL-147; BW-1 thru BW-29

1

2

3






Electrical System

DC Electrical System (With EFIS)

VENT BLOWER POWER


COPILOT'S WINDSHIELD

ANTI-ICE

DC TEST JACK


PILOT'S WINDSHIELD

ANTI-ICE

LEFT RADIANT HEAT

CONDENSER BLOWER





AVIONICS MASTER

CONTROL

ON

OFF

AVIONICS SWITCH

60A

40A

40A

5A

60A

LDG GEAR CONTROL


LANDING GEAR

MOTOR

30A 30A

30A

5A

TO AC POWER

TO AC POWER


60A 50A

40A 50A

50A

20A

AVIONICS BUS NO. 1

AVIONICS BUS NO. 2

AVIONICS BUS NO. 3

RIGHT GENERATOR BUS

LEFT GENERATOR BUS

AFT ELECTRIC HEAT

AFT ELECTRIC NO. 2

INVERTER

115 VAC

26 VAC

NO. 2 INVERTER

115 VAC

26 VAC






Electrical System

Circuit Breaker PanelsS/Ns BB-1439, BB-1444 through BB-1485, except BB-1436, and BB-1484; BL-139 and BL-140

ENGINES AVIONICS

PROP LEFT FIRE LEFT LEFT LEFT LEFT LEFT LEFT AVIONICS

5 5 5 5 71 2 5 5 2 5 5

AUTO

SYNC CHIP DET

DET FUEL CONTROL

HEAT

ICE VANE

CONTROL OIL

TEMP OIL

PRESS TORQUE METER

FUEL FLOW

AVIONICS OAT

PROBE

MASTER GND

COMM GND

AUDIO

5 5 5 71 2 5 5 2 5 30 15 71

2 2 FEATHER RIGHT

LIGHTS RIGHT RIGHTRIGHT

WARNINGS RIGHT RIGHTRIGHT

WEATHER NO. 3 HEA

AVIONICS FLIGHT INSTR SIDE STAL LEFT WAR POWER LEFT SURF BRAKE AVIONICS COMM NA PILOT

5 71 2 5 5 5 5 5 71

2 5 5 5 30 71 2 2 2

ANN

READING

INSTR INDIRECT PANE

AVIONICS & ENG OVHDL

NO SMK & FSB

WAR BLEED

AIR WARN

LANDING GEAR

ANN FUEL VENT

WSHLD

DE-ICE DE-ICE

AVIONICS

NO. 1

RADIO

NO. 1

NA

NO. 1

VLF STEREO

AUDIO

5 5 5 5 5 5 5 5 10 30 2 2 71 2 4

INSTR CONSOLE CABIN RIGHT IND IND RIGHT WIPER NO. 2 RELA NO. 2 ENVIRONMENTAL FLIGHT ELECTRICAL

OXYGEN PRESS LEFT YA PILOT PILOT PITCH AL NO. 1 LEFT NO. 1 XPONDER DME AP AP FL T

5 5 5 5 5 3 5 1 5 10 50 50 3 2 10 2 5 CONTROL CONTROL BLEED

AIR CONTROLTEMP

TURN & SLIP

AIR SPEED

DAMP ALTM AIR

DATA

TRIM

RUDDER

ALERT INV

CONTROL GEN

CONTROL BUS

FEEDERS

NO. 1

XPONDER RMI RADAR

POWER

RADAR RADAR

ENGAGE DIR

5 5 3 1 5 5 10 50 50 3 2 5 5 3 CONTROL RIGHT & AIR DATA SPEED BOOST NO. 2 RIGHT NO. 2 NO. 2 NO. 1 CMPTR DISPLA

FURNISHING FLT MASTER CIGAR ADF RMI RADIO PRFL RADIO

10 5 2 2 2 1 4 POWER LIGHTER NO. 2 ALT ADVSY PHONE

AUX AUX AUX INSTR AVIONICS COMM COPILOT COMPASS ILS VNA

2 20 71 2 2 3 2 1

INDR LT NO. 2 AUDIO NO. 2 NO. 2

FUEL SYSTEM OPEN OPEN

5 10 5 5 5 5 5 5 5 10 5

CLOSED FIRE STANDBY AUX QTY PRESS CROSS PRESS QTY AUX STANDBY FIRE CLOSED VALV FER FER VALV WALL PUMP TRANS IND WARN FEED WARN IND TRANS PUMP WALL

FIREWALL FIREWALLLEFT RIGHT SHUT OFF VALVE SHUT OFF VALVE

IGNITOR STAR NO. 3 NO. 4

PROPBUS FEEDERS FLAPPROP DE-ICE CONTROL PROP PROP MOTOR CONTROL GOV POWER CONTROL

5 5 5 5 20 20 20 5 5 550 50 50 50 LEFT RIGHT LEFT RIGHT LEFT RIGHT

Circuit Breaker PanelsS/Ns BB-1484, BB-1486 and Subsequent, BL-141 and Subsequent, BW-1 and Subsequent

ENGINES AVIONICS

PROP LEFT FIRE LEFT LEFT LEFT LEFT AVIONICS

5 5 5 5 71 2 5 5 5

AUTO

SYNC CHIP DET

DET FUEL CONTROL

HEAT STBY ENG ANTI-ICE

MN ENG ANTI-ICE

ENG INSTR

POWER AURAL PILOT

MASTER

CABIN COPILOT

5 5 5 71 2 5 5 30 15 71

2 2 FEATHER RIGHT

LIGHTS RIGHT RIGHT RIGHT

WARNINGS RIGHT

WEATHER WAR AUDIO AUDIO AUDIO

AVIONICS COPL FLT INSTR PLT FLT STAL LEFT WAR POWER LEFT SURF AVIONICS COMM NA XPNDR

5 71 2 5 5 5 5 5 71

2 5 5 30 71 2 2 2

ANN

READING

INSTR INDIRECT SIDE PNL

AVIONICS & ENG OVHDL

NO SMK & FSB

WAR BLEED AIR

WARN LANDING

GEAR ANN FUEL

VENT WSHLD

DE-ICE NO. 1

AVIONICS

NO. 1

COMM

NO. 1

NA

NO. 1

XPNDR

5 5 5 5 5 5 5 5 10 30 2 2 71 2

INSTR CONSOLE CABIN RIGHT IND IND RIGHT WIPER NO. 2 NO. 2 NO. 2 NO. 2 ENVIRONMENTAL FLIGHT ELECTRICAL

OXYGEN PRESS LEFT ALT PILOT PILOT PITCH OUTSIDE NO. 1 LEFT NO. 1 RMI DME COMPASS

5 5 5 5 5 3 5 1 5 10 50 50 3 2 10 CONTROL CONTROL

TEMP

BLEED AIR

CONTROL

ALER TURN & SLIP

ALTM AIR

COPLT

TRIM

RUDDER

AIR TEMP INV

CONTROL GEN

CONTROL BUS

FEEDERS

NO. 2

RMI

NO. 1

DME

NO. 1

COMPASS

5 5 3 1 5 5 10 50 50 3 2 5 CONTROL RIGHT COPLT ENCD BOOST NO. 2 RIGHT NO. 2 NO. 2 NO. 1 NO. 2

FURNISHING MASTER CIGAR AP FCS EHSI HDG

10 5 2 2 2 1 POWER LIGHTER SERVO POWER PRCSR

RADAR ADF RADIO

71 2 2 3

ALT

FIRE AUX AUX FIRE OPEN WALL STANDBY TRANS QTY PRESS CROSS PRESS QTY TRANS STANDBY WALL OPEN

VALV PUMP FER IND WARN FEED WARN IND FER PUMP VALV FIREWALL FIREWALL SHUT OFF SHUT OFF5 10 5 5 5 5 5 5 5 10 5VALVE VALVE

CLOSED CLOSED LEFT RIGHTFUEL SYSTEM

ENGINE INSTRUMENTS FLAP LEFT

50 25 20 5 5 5 50 5 5 5 5 5 5 5

NO. 3 LEFT L LEFT EFT NO. 3

MOTOR CONTROL BUS PROP IGNITOR START BUS PROP TURBINE FUEL OIL OIL

FEEDERS DEICE PROP POWER CONTROL FEEDERS ITT TORQUE TACH TACH FLOW PRESS TEMP

DEICE

50 25 5 5 5 5 50 5 5 5 5 5 5 5 NO. 4 RIGHT CONTROL GOV RIGHT RIGHT NO. 4 RIGHT




Electrical System



DC Electrical SystemThe Super King Air Direct Current (DC) electrical system provides and distributes 28 VDC power via a Dual-Fed bus system. At least two power sources feed the major buses so backup power is continuously available to most aircraft electrical systems. One 24V, 34 amp-hour, 20-cell nickel-cadmium (Ni-Cad) battery and two 28V, 250A parallel-connected starter/generators provide power to the aircraft’s DC electrical bus system. In addition, an external power receptacle allows connection of an auxiliary power unit for ground operation.

BatteryThe 24V Ni-Cad battery is in the right wing center section, forward of the main spar.A thermostat-controlled valve at the air inlet under the battery opens and closes for air-cooling. At 80°F, the valve is fully open; at 30°F, the valve is fully closed. The air-cooling thermostat, battery relay, and charge monitor shunt are mounted either under or on the battery box.The battery supplies electrical power for starting and for emergency operation of essential equipment powered from the Hot Battery bus. (See DC Power Distribution, this chapter.)

Figure 5C-1: 24V Ni-Cad Battery Figure 5C-2: Thermostat-Controlled Valve



Bmaahry wtaScThe battery ON/OFF switch on the pilot’s left subpanel controls the battery relay. With the battery switch on, battery power from the Hot Battery bus routes through the battery relay to the Main Battery bus.

Figure 5C-3: Battery ON/OFF Switch

The Main Battery bus feeds the Isolation bus, which connects the left and right Generator buses together via two 325A current limiters. The Main Battery bus also feeds two starter circuits controlled by starter switches and starter relays. Reverse-current protection prevents the generators from absorbing power from the Generator buses if the generators are not operating, or when generator voltage is less than bus voltage.

MAIN BU S

L E F T

G E N

B U S

HOT BUS

BAT STARTER/

GENERATOR STARTER/

GENERATOR

325 CURRENT LIMITER

BAT RELAY

START RELAY

START RELAY

ISOLATION BUS

R I G H T

G E N

B U S

Figure 5C-4: Main Battery Bus

Electrical System



Tchreml RunmwmyA Ni-Cad battery overheat condition causes battery charge current to increase. Thermal runaway is a self-sustaining reaction caused by excessive heat buildup. As cell heat increases, internal current flow increases; the increased current flow causes further cell heat increase. The cycle becomes self-sustaining, even without a charging source.

Bmaahry Montaor ysaheA battery charge monitor system alerts the crew to a possible thermal runaway. On S/N BB-037 and subsequent, BL-001 and subsequent, and earlier models with SI-0701-356, a battery charge current detector continuously monitors the battery condition.

Figure 5C-5: Battery Monitor System

The detector illuminates a yellow BATTERY CHARGE annunciator if the battery charge current exceeds 7A for six seconds or more. In addition, both MASTER CAUTION annunciators flash until they are manually reset.The BATTERY CHARGE annunciator may illuminate for short intervals when heavy loads switch off. Following a battery-powered engine start when the battery recharge current is very high, the current detector illuminates the BATTERY CHARGE annunciator; this provides an automatic self-test of the detector and the battery.As the battery approaches a full charge and the charge current decreases to a satisfactory level, the annunciator extinguishes. This normally occurs within a few minutes after an engine start; however, the annunciator may remain illuminated for a longer period if the battery has a low state of charge, low charge voltage per cell, or low battery temperature.If the BATTERY CHARGE annunciator illuminates in flight, turn off the battery; if the loadmeter change is greater than 2.5% of the loadmeter scale, leave the battery off for the remainder of the flight (except for landing), as conditions may exist that will eventually damage the battery.



Inspect the battery after landing. Refer to the Nickel Cadmium Battery Condition Check in the manufacturer’s AFM. If the annunciator remains on after the battery switch is off, there is a malfunction in either the battery relay or the current detector. The battery may still be connected to the electrical system. Land as soon as practical.Turn the battery switch on prior to landing to avoid electrical transients caused by power fluctuations.

NTTc: SI-0701-356; Electrical Battery Installation -Installation of an improved battery charge current detector.

Starter/GeneratorsTwo 28V to 28.5 VDC engine-driven starter/generators are rated at 250A. In the start mode, each functions as a starter with power from the aircraft battery via the Main Battery bus. With the start mode deselected and the generator mode selected, the starter/ generator then generates regulated voltage to the aircraft electrical system via the Generator buses. Reverse current protection prevents each generator from absorbing power from the Generator bus if the generators are not operating or if generator voltage is less than bus voltage. (See DC Distribution System, Generator Buses, this chapter.)See Table 5C-1 for generator load limits.

GENERATOR LOAD MINIMUM GAS GENERATOR RPM — N1

WITHOUT AIR CONDITIONING

WITH AIR CONDITIONING1

King Air 200 0 to 0.70 52% 60%0.70 to 0.75 55% 60%0.75 to 0.80 60% 60%0.80 to 0.85 65% 65%King Air B200 0 to 0.75 56% 62%0.75 to 0.80 60% 62%0.80 to 0.85 65% 65%

Table 5C-1: King Air 200 and B200 Generator Load Limits1 Right engine only

NTTc: Some aircraft may have either of two optional Lear Siegler 300A starter/generators. See the Quick Reference chapter of this manual for load limits.

Electrical System



Ghnhrmaor wtaSchsThe GEN 1 and GEN 2 switches on the pilot’s left subpanel control generator operation. The MASTER SWITCH gang bar may be used to switch off the battery and generator switches simultaneously.On S/N BB-088 and subsequent, to turn the generator on, first hold the generator switch upward in the spring-loaded RESET position for a minimum of one second, then release it to the ON position.

Npatonml Ghnhrmaor Nvhrchma Montaortng ysahe If the temperature in either generator exceeds 315°F, the corresponding red L or R GEN OVHT annunciator on the Master Warning panel illuminates and both MASTER WARNING annunciators flash. If either generator is off-line, the corresponding amber L or R DC GEN caution annunciator illuminates and both MASTER CAUTION annunciators flash.

Vola/LomdehahrThe left and right dual-scale DC volt/loadmeters are below the overhead light switch panel. The pointer normally indicates generator load from 0 to 100% on the upper scale. To read DC voltage, press the button to the lower left of the meter and read the indicated voltage, which ranges on the scale from 0 to 30 VDC.

Figure 5C-6: DC Volt/Loadmeters



Generator Voltage Regulation SystemsThe generator voltage regulation systems provide paralleling, reverse-current cutout, 28 VDC to 28.5 VDC voltage regulation, and 32 VDC to 34 VDC overvoltage protection.Two basic types of voltage regulation systems are used in the Super King Air 200/B200. On S/N BB-002 through BB-087, transistorized regulators, over-voltage relays, paralleling relays, and reverse-current relays comprise the system. On S/N BB088 and subsequent, an integrated generator control panel with line contactor relays comprise one unit that performs all functions.

Volamgh Rhgulmaton ysahe(S/N BB-002 through BB-087)During variations in engine speed and electrical load requirements, two transistorized voltage regulators, one for each engine, maintain a constant level voltage output.The voltage regulator system uses voltage developed in each generator’s compensator windings to equalize the generators. A sensor reads the voltage at the compensator winding interpole; this voltage then connects from its respective paralleling rheostat to the paralleling rheostat of the opposite generator through the intervening voltage regulators.The field current of the generator carrying the higher current load reduces, while that of the generator carrying the lower current load increases until the load on each generator is equal.The paralleling circuit also includes paralleling relays. When one generator is on-line and the other is off-line at the same voltage, the paralleling circuit depresses the voltage of the on-line generator and increases that of the off-line generator until both are on-line.

Overvoltage ProtectionIf an overvoltage condition occurs, the overvoltage relay trips and removes the overvoltage generator from the line; this leaves the other unit to supply the entire load.The overvoltage relay is set to trip at 32V to 34V. When the overvoltage relay actuates, it opens to remove the voltage regulator from its source of sensing voltage, interrupting the field current of the generator.Actuation of the overvoltage relay also removes the voltage from the SW terminal of the reverse-current relay, to permit it to open and remove the generator from the bus. If the overvoltage condition is the result of a voltage regulator malfunction, the overvoltage relay stops the overvoltage condition. If some other failure, such as a fault, causes the overvoltage condition, the reverse-current relay isolates the overvoltage generator from the bus to protect the rest of the system from damage.

Electrical System



Ghnhrmaor Conarol Pmnhls(S/N BB-088 and subsequent, BT, BL, and BN Models)Two generator control panels mounted below the center aisle floor control generator operation. The generator control panels provide the voltage regulation, generator paralleling, reverse-current sensing and control, overvoltage/undervoltage protection, and over-excitation protection.Individual voltage regulators control the generators to provide a constant voltage to the buses during variations in engine speed and electrical load requirements. The generator switches on the pilot’s left subpanel connect the generators to the voltage regulating circuits (see Generator Switches, this chapter). The voltage regulating circuit automatically disables or enables a generator when voltage limits are exceeded. The left or right volt/loadmeter in the overhead panel indicates the load on its respective generator.

STARTER-GENERATOR

EXTERNAL POWER CONNECTOR

INVERTER

BATTERY PRINTED CIRCUIT BOARDS

GENERATOR CONTROL PANEL (BB-88 and Subsequent)

STARTER-GENERATOR

INVERTER

Figure 5C-7: Generator Control Panels

Control Panel Overvoltage ProtectionThe generator control panel provides overvoltage protection. If the output voltage reaches 32 VDC, the overvoltage protection portion of the generator control panel opens the coil circuit of the bus contactor and isolates the overvoltage generator from the aircraft bus. The normally regulated generator then comes on line to supply the electrical system.

Control Panel Reverse-Current ProtectionThe control panel provides reverse-current protection and generator control. Bus voltage is sensed at the bus side of the bus contactor. The generator output voltage is sensed at the generator side of the bus contactor.



When the generator is operating and the control switch is in the RESET position, the generator output voltage rises to the regulated voltage. When the generator switch is in the ON position, a voltage output from the control panel closes the bus contactor to connect the generator to the bus.When the generator slows down to the point where it can no longer maintain a positive load, the generator begins to draw current from the aircraft bus. This reverse-current passes through the compensator windings of the generator and the resultant voltage is sensed at the interpole terminal of the generator. The generator control panel then removes the voltage from the coil of the bus contactor, which opens and removes the generator from the aircraft bus.If the bus voltage exceeds 28.25 VDC, reverse-current begins to flow in the normally regulated generator.This generator is then removed temporarily from the bus by the reverse-current relay. The generator with the overvoltage condition attempts to assume the entire electrical load.The resultant bus voltage then depends on the generator speed, the electrical load, and the nature of the fault.When the engines are operating, generator power flows through the reverse-current protection circuitry to the right and left Generator buses, which are interconnected by two 325A current limiters to protect against ground faults.

External Power

CAAUTIO A maximum continuous load in excess of 350A will damage the external power relay and power cables of the airplane.

CAAUTIO Any current in excess of 1,000A may overtorque the drive shaft of the starter generator or produce heat sufficient to shorten the life of the unit.

CAAUTIO Voltage is required to energize the avionics master power relays that remove power from the avionics equipment. Therefore, never apply external power to the aircraft without first applying battery voltage. If the battery is removed from the aircraft or if the battery switch is to be turned off, connect an external battery parallel to the external power unit prior to energizing the power unit.

CAAUTIO The battery may be damaged if exposed to voltages higher than 30V for extended periods.

Electrical System



An external power receptacle under the right wing outboard of the nacelle allows connection of an auxiliary power unit for ground operation. The receptacle is designed for a standard AN plug.A relay and a diode in the external power circuit protect the aircraft electrical system from reverse polarity damage. The relay closes only if the external power source polarity is correct (24V to 28 VDC).

Figure 5C-8: External Power Receptacle

On S/N BB-364 and subsequent and BL models, the battery switch must be on before the external power relay can close and allow external power to enter the aircraft electrical system.External power can be used to operate all the airplane electrical equipment (this includes avionics checkouts) during ground operations without the engines running, and it can be used to start the engines. The external power circuit is capable of accepting 400 Amps continuously, and it can withstand current surges up to 1,100 Amps for short durations (up to 100 milliseconds), which may occur during engine starting. Refer to the BEECHCRAFT Super King Air 200 Series Maintenance Manual for ground checkout information.The EXT PWR caution annunciator illuminates when an external DC power plug is connected to the aircraft. Observe the following precautions when using an external power source. Use only a negatively grounded auxiliary power source. If the

polarity of the power source is unknown, determine the polarity with a voltmeter before connecting the unit to the aircraft.

Before connecting an external power source, turn off all radio equipment and generator switches. Leave the battery on to protect transistorized equipment against transient voltage spikes.

If battery voltage indicates less than 20V, the battery must be recharged or replaced with a battery indicating 20V or greater before using auxiliary power. The battery switch must be on when starting an engine with auxiliary power. Generator switches should remain off until auxiliary power is disconnected.



If the ground power source does not have a standard AN plug, check the polarity of the plug. The positive lead from the ground power source must connect to the center post, the negative lead must connect to the front post, and a positive voltage of 24V to 28 VDC must be applied to the small polarizing pin of the aircraft’s external power receptacle.

NTTc: The generators do not come online with auxiliary power connected.

DC Distribution SystemThe DC distribution system receives power from the battery. and both generators. Each of four Dual-Fed buses receives power from at least two sources. If one generator fails, the Dual-Fed buses and the systems they supply continue to receive power. Eleven buses (12 buses if the optional Avionics bus No. 3 is installed) comprise the DC distribution system: Hot Battery bus Main Battery bus Isolation bus Left and Right Generator buses No. 1 Dual-Fed bus No. 2 Dual-Fed bus No. 3 Dual-Fed bus No. 4 Dual-Fed bus No. 1 and No. 2 Avionics buses Optional No. 3 Avionics bus.

Each DC power source feeds directly to its own bus: the battery to the Hot Battery bus and the generators to their respective Generator buses. Two 325A current limiters interconnect the buses so that any one source can feed all buses.If a major ground fault occurs on the Generator bus, the current limiter opens to isolate the affected bus. The fault also causes the affected generator to trip off-line.When the battery, generators, or external power unit are providing power, the Isolation bus and left and right Generator buses function as one bus if both current limiters are intact.

Hoa Bmaahry BusThe Hot Battery bus, outboard of the battery box, powers the cabin entry lights, fire extinguishers, and firewall shutoff valves. On S/N BB-002 through BB-1097, the standby fuel pumps are also wired to the Hot Battery bus.On BB-1098 and subsequent, the standby fuel pumps are connected to Dual-Fed buses only. Items powered by the Hot Battery bus may be operated even though the battery master switch is off and the battery relay is open. Isolation diodes prevent interaction between the Hot Battery bus and the Main Battery bus.

Electrical System



The Hot Battery bus supplies power to the following: left and right engine fire extinguishers (optional) left and right firewall shutoff valves left and right standby fuel pumps (S/N 002-1097) entry lights clock external power sense battery relay optional RNAV memory optional stereo.

Mmtn Bmaahry BusThe Hot Battery bus supplies DC through the battery relay to the Main Battery bus. When an external power source is used, the external power connector directs power through the external power relay to the Main Battery bus.

Isolmaton BusThe Isolation bus receives battery power directly from the Main Battery bus and generator power via two separate 325A current limiters from the left and right Generator buses. Therefore, the battery and both generators power the entire system. The current limiters provide isolation of the battery from a faulted generator bus in the event of a ground fault.

Ghnhrmaor BushsEach generator directly supplies its respective Generator bus and indirectly supplies the opposite Generator bus through the Isolation bus.After starting, the generators supply power for the DC systems and battery charging through their respective left and right Generator buses.The individual generators also supply power for items not on the Dual-Fed buses. The right Generator bus supplies the following: DC Avionics No.2 bus No.2 Inverter landing gear motor vent blower copilot’s windshield heat right radiant heat aft evaporator blower power air conditioner clutch miscellaneous equipment.

Items supplied by the left Generator bus include the following: DC Avionics No. 1 bus No. 1 Inverter



if installed, the DC Avionics No. 3 bus pilot’s windshield heat condenser blower left radiant heat miscellaneous equipment.

Duml-Fhd BushsBoth Generator buses feed each of four Dual-Fed feeder buses through a 60A current limiter, a 70A diode, and a 50A CB. The Dual-Fed buses supply power to the majority of the aircraft’s utility systems; they continue to receive power even if one Generator bus is lost.The four BUS FEEDER CBs for the No. 1 and No. 2 Dual-Fed buses are on the copilot’s side panel; the No. 3 and No. 4 CBs for the Dual-Fed buses are on the pilot’s fuel panel. Figure 5C-9 shows CB panel for SN 1444 and subsequent aircraft. The diodes at each end of the Dual-Fed buses -prevent reverse current flow across the bus to a faulted generator bus or feeder circuit.

Figure 5C-9: Copilot’s Side Panel

All electrical loads are divided among these buses with equipment arranged so that all items with duplicate (i.e.,left and right) functions are split between two different buses. The items on Dual-Fed buses No. 1 and No. 3 are left functions; those on buses No. 2 and No. 4 are right functions.For example, the left landing light is on the No. 1 Dual-Fed bus, while the right landing light is on the No. 2 Dual-Fed bus.On S/N BB-088 and prior, the DC power from each generator routes from the engine-driven generator to the reverse-current relay, the volt/loadmeter shunt and Generator bus to the Dual-Fed buses on the DC power panel assembly.

Figure 5C-10: Pilot’s Fuel Panel

Electrical System



S/N BB-088 and subsequent, BT, BL, and BN models have generator bus contactors that are controlled by the generator control panels to provide the reverse-current protection. The generator contactors and the volt/loadmeter shunts are mounted under the center aisle floor.

DC AvtontSs BushsThe left Generator bus feeds the DC Avionics bus No. 1 through a 40A current limiter and a 30A CB and the DC Avionics bus No. 3, if installed, through a 30A CB. The right Generator bus feeds the DC Avionics bus No. 2 through a 40A current limiter and a 30A CB. Equipment powered from these buses varies according to the individual aircraft avionics installations.

AvtontSs Mmsahr PowhrThe avionics systems installed on each aircraft vary, but they usually consist of individual nav/com units, each having its own control switch. In addition, a factory-installed two-position AVIONICS MASTER PWR/OFF switch on the pilot’s left sub-panel controls power to receivers and transmitters in the aircraft.

Figure 5C-11: AVIONICS MASTER PWR/OFF Switch



Three avionics master power relays automatically open to isolate avionics equipment when power is applied to the aircraft bus system. This fail-safe design allows operation in the event of master switch failure.Selecting the AVIONICS MASTER PWR switch on overrides the isolation function to allow checkout and maintenance of the avionics subsystems while using DC external or aircraft power. The avionics master CB on the copilot’s circuit breaker panel powers the avionics master power relays.

AVIONICS MASTER POWER SWITCH

AVIONICS NO. 2 BUS

AVIONICS NO. 1 BUS

NO. 2 AVIONICS MASTER POWER RELAY

NO. 1 AVIONICS

MASTER POWER RELAY

AVIONICS MASTER POWER SWITCH

AVIONICS NO. 2 BUS

AVIONICS NO. 1 BUS

ONOFF

ONOFF

NO. 1 DUAL FED BUS 5A

30A

LEFT GEN BUS

RIGHT GEN BUS

NO. 1 DUAL FED BUS 5A

LEFT GEN BUS

RIGHT GEN BUS

NO. 2 AVIONICS MASTER POWER RELAY

NO. 1 AVIONICS

MASTER POWER RELAY

40A

30A 40A

30A 40A

30A 40A

Figure 5C-12: Avionics Master Power

Circuit Breaker System

CAAUTIO If a non-essential CB or 50A feeder CB on either of the two circuit breaker panels trips while in flight, do not reset. Resetting can cause further damage to components or systems, especially to the eight 50A Dual-Fed bus feeders and the three 30A avionics bus feeders. If an essential system CB (less than 20A) trips, wait a minimum of five seconds for the breaker to cool, then reset. If it fails to reset, do not attempt to reset. Take corrective action according to Emergency procedures in your CAE Operating Handbook. This procedure also applies to the switch CBs on pilot’s right sub-panel.

Both AC and DC power are distributed to the various aircraft systems via two separate circuit breaker panels that protect most of the aircraft components. Either switch-type circuit breakers or push-pull circuit breakers protect the individual bus components. The right side panel contains circuit breakers for Dual-Fed buses No. 1 and No. 2 and the Avionics buses No. 1, No. 2, and optional No. 3. This panel contains CBs for the electrical distribution system and the following equipment: major engine-related electrical systems all avionics components

Electrical System



environmental system lights annunciator warning systems.

The fuel control panel contains circuit breakers for the Dual-Fed buses No. 3 and No. 4. This panel protects the bus feeders and other systems including: standby boost pumps propeller de-ice flap motor and control igniter power start control firewall shutoff valves propeller governor test.

Tngtnh amratng mnd Igntaton

CAAUTIO Do not exceed the starter motor operating time limits of: 40 sec. ON/60 sec. OFF 40 sec. ON/60 sec. OFF 40 sec. ON/30 min. OFF.

NTTc: Battery voltage of 23V or higher permits a battery start. 20V or higher permits an external power application. If voltage is below 20V, the battery should be charged or changed before using external power.

The Main Battery bus provides starter power to each individual starter/generator through a starter relay.



The IGNITION AND ENGINE START three-position (ON/OFF/ STARTER ONLY) switches on the pilot’s left subpanel control the start relays.With the switch in the ON position, the starter system is in both the start and ignition mode, and activates the following: the starter relay and starter motor the ignition excitor and corresponding IGNITION ON annunciator the generator control relays overvoltage relay (BB-002 through

BB-087) or field power disconnect (BB-088 and subsequent) the fuel purge valve.

Figure 5C-13: IGNITION AND ENGINE START Switch

At 50% N1, turn the starter and ignition switch off; this closes the fuel purge valve, opens the field disconnect relay (or overvoltage circuit), turns off the ignition, and stops power to the starter windings. At this point, the engine should be self-sustaining.With the switch in the STARTER ONLY position, the above systems are activated, except for the ignition excitor and IGNITION ON annunciator. The STARTER ONLY mode is normally used for compressor wash or for purging fuel after an aborted start.An illuminated L or R IGNITION ON annunciator indicates one of the following conditions: the corresponding starter/ignition switch is in the engine/ignition

mode

Electrical System



the corresponding auto ignition is armed and the associated engine torque is below 400 ft-lbs.

The starter/generator drives the compressor section of the engine through the accessory gearing. The starter/ generator initially draws approximately 1,000A, then drops rapidly (to about 300A for a battery start or 400A for a 28V external power unit start) as the engine reaches 20% N1.

Auao IgntatonThe auto ignition system provides automatic ignition to prevent engine loss due to combustion failure. This system ensures ignition if flame-out occurs during critical phases (e.g., takeoff, landing, turbulence, and ice and rain conditions).Arm the system prior to takeoff and turn it off after landing. To arm the system, move the ENG AUTO IGNITION lever lock switches on the pilot’s left subpanel to ARM. If the engine torque falls below 400 ft-lbs, the igniter automatically energizes and the IGNITION ON annunciator illuminates.For extended ground operation, turn the auto ignition system off to prolong the life of the igniters. (See the Power-plant chapter.)



AC Electrical System

YAW

RIGHT ENGINE TORQUEMETER

LEFT ENGINE TORQUEMETER

LEFT ENGINE FUEL FLOW (PRIOR TO BB225)

RIGHT ENGINE FUEL FLOW (PRIOR TO BB225)

LEFT ENGINE TORQUEMETER

RIGHT ENGINE TORQUEMETER

YAW RATE (BB113 AND AFTER)

380 100

120 130

11

AC VOLTS

390 400 410

420

INVERTER INST INV

TEST JACK

VOLTS/FREQUENCY METER

KING AIR B200 KING AIR 200

MODEL VARIATION ANNUNCIATOR

KING AIR 200

KING AIR B200

2 6 V A C

B U S

2 6 V A C

B U SNO. 2

INV

115V

26V AC

AC AVIONICS

AND RADAR

INV CONTROL

RELAY

INV SELECT SWITCH

INV SELECT RELAY

INV CONTROL

RELAY

115V

26 V AC

NO. 1 INV CONTROL

NO. 2 INV CONTROL

FROM NO. 1 DUAL-FED BUS


50A

5A

5A

5A

50 A

10A

5A

5A

5A

5A

10A

FROM NO. 2

NO. 1 INV

1 BB1097, 1095 AND PRIOR

1

1

RIGHT GEN BUS

LEFT GEN BUS

DUAL-FED BUS

Electrical System



AC Electrical System

AC InvertersThe left and right Generator buses supply 28 VDC power to their respective inverters through the inverter control relay. Two static 400 Hz, single-phase 750VA inverters supply AC electrical power for the AC avionics equipment and AC-powered engine instruments. Each inverter provides two levels of power: 115 VAC, 400 cycle and 26 VAC, 400 cycle.Either inverter can be selected to power, via a 10A CB, the 115 VAC systems, which include the AC avionics and radar, the inverter warning relay, and the AC volts/frequency meter.In addition, either inverter can power the 26 VAC systems. The 26 VAC is tapped from a stepdown transformer within the inverter and routes to the 26 VAC bus via a 5A CB.On the 200 Series, the 26 VAC bus powers the following: yaw rate (BB-113 and subsequent) left and right engine torquemeters left and right engine fuel flow (BB-001 through BB-224).

On the B200 Series, the 26 VAC bus powers the following: yaw rate left and right engine torquemeters.

The inverters are in the wing center section outboard of each nacelle.



INVERTER SwitchThe pilot selects either inverter with a three-position (NO. 1/OFF/NO. 2) switch on the left sub-panel. Selection of either inverter actuates the associated inverter power relay to supply the inverter with DC power. An inverter select relay provides the necessary switching to permit the operating inverter to supply 26 VAC instrument power and 115 VAC power for test jack, avionics, radar, and the inverter warning relay.The inverter warning relay energizes to prevent illumination of the inverter warning annunciator.

Figure 5C-14: INVERTER Switch

AnnunciatorsIn the event of an AC bus or inverter failure, the red INVERTER annunciator on the Master Warning panel illuminates. In addition, the MASTER WARNING annunciator flashes until it is manually reset.

Electrical System



AC VoltmeterAll aircraft have a frequency/voltage meter that monitors the inverter output. The single pointer on the dual scale normally monitors bus frequency, indicated on the top scale from 380 to 420 Hz. To read voltage, press the button in the lower left corner of the meter; the voltmeter reads from 100 to 130 VAC on the lower scale.

NTTc: Nominal AC voltage should be 115 ± 3% and 400 ± 1% Hz.

Figure 5C-15: AC Voltmeter

Inverter Circuit ProtectionThe 5A INV CONTROL NO. 1 and NO. 2 CBs on the right circuit breaker panel protect the inverter circuits. In addition, two inverter power current limiters in the DC power distribution panel and two inverter select relay CBs provide circuit protection.




Electrical System




PreflightDuring the interior cockpit inspection, ensure that all switches on the fuel control panel, pilot and copilot subpanels, overhead panel, and pedestal are either in the OFF or AUTO position. Verify the CBs on the fuel control panel and right side panel are in and the landing gear relay is in. Check that the landing gear control switch on the pilot sub-panel is down.Perform a Hot Battery bus check as detailed in the Preflight chapter. Check the battery; both voltmeters should read normal battery voltage of 24V to 25V to verify the condition of the battery and current limiters. During the exterior inspection, check the battery air inlet on the bottom of the wing between the nacelle and the fuselage. The thermostatically controlled valve should be securely in place, should not bind, and should be in the proper position for battery box temperature (fully open at 80°F and fully closed at 30°F).Ensure that all exterior lights are secure and in good condition; check for cracked lenses or burned out bulbs.Refer to the Expanded Normal Procedures and Preflight chapters for detailed checklists.

Abnormal ProceduresThe following provides a brief explanation and overview of abnormal procedures. Please refer to the CAE King Air 200/B200 Operating Handbook for detailed checklists for electrical system abnormal procedures.

BATTERY CHARGE Annunciator Illuminated on the GroundDuring engine start, the BATTERY CHARGE annunciator normally illuminates approximately six seconds after a generator is on-line to indicate a charge current above normal. The annunciator should extinguish within five minutes following normal engine start; if it does not, the battery is partially discharged.After the loadmeter stabilizes, momentarily turn the battery switch OFF and note the change in meter indication. Failure to obtain a change value of below 2.5% within five minutes indicates a partially discharged battery.Continue to charge the battery, repeating the above check every 90 seconds until the charge current decreases below 2.5%. No decrease in charging current between checks indicates an unsatisfactory condition. Do not take off. Have the battery removed and serviced by a qualified nickel-cadmium battery shop.



BATTERY CHARGE Annunciator Illuminated in FlightA BATTERY CHARGE annunciator illuminated in flight indicates a possible malfunction. Turn the battery switch to OFF momentarily to note the change in the loadmeter reading; if the change value exceeds 2.5%, turn the battery switch off, and proceed to the destination.If the BATTERY CHARGE annunciator does not extinguish when the battery control switch is placed in the OFF position, land as soon as possible.To avoid electrical transients caused by power fluctuations, turn the battery switch to ON for landing gear and flaps extension. Perform the During Engine Shutdown Battery Check according to your CAE Operating Handbook. If the check is unsatisfactory, remove the battery and have it serviced by a qualified nickel-cadmium battery shop.

Emergency ProceduresThe following provides an overview of emergency procedures. Please refer to the CAE King Air 200/ B200 Operating Handbook for detailed check-lists for electrical system emergency procedures.

Electrical System Failure

Ghnhrmaor Inophrmatvh (DC GT AnnunStmaor N)If the left or right DC GEN annunciator illuminates in flight, turn that generator off. Wait one second, then turn the generator switch to RESET and ON. If the generator does not reset, turn it off and rely on the operating generator. Monitor the loadmeter so that the load on the remaining generator does not exceed 100%. Manage essential systems; plan for landing gear deployment, lights, and other high-draw items.If either loadmeter exceeds 100%, turn off the battery switch and monitor the loadmeter. The battery may be taking an excessive charge; watch for a possible loadmeter drop when the battery is switched off.If the loadmeter still indicates 100%, turn off all non-essential electrical equipment. The problem is not a -battery charge but could be a malfunctioning airframe system or avionics system or a direct short.If the reading goes below 100%, turn the battery switch on. Continue to monitor the loadmeters during the remainder of the flight.In the event of an overvoltage condition to the generator buses, the overvoltage protection circuits isolate the affected generator, the appropriate DC GEN annunciator illuminates, and the remaining generator assumes the total electrical load.

Electrical System



Current Limiter Check With One Generator Off LineIf a generator fails, check the voltmeter on the failed generator’s side to ensure the current limiter on the corresponding bus is operative. A 28 VDC reading indicates both current limiters are intact. A 24 VDC reading indicates the current limiter on the operating generator side has opened. A reading of no voltage indicates the current limiter on the failed generator side has opened, or both could have opened.

Bus Feeder Circuit Breaker Tripped in FlightIf a non-essential CB or 50A feeder CB on either of the two circuit breaker panels trips while in flight, do not reset it. Resetting a tripped breaker can cause further damage to the component or system; this is especially true of the eight 50A Dual-Fed bus feeders and the 30A Avionics bus feeders.If an essential system CB (less than 20A value) trips, wait one second, then reset it. If it fails to reset, do not attempt to reset it again.To determine the systems and components affected by the isolated bus, refer to the appropriate power distribution schematic for the aircraft. Take corrective action according to Emergency procedures in your CAE Operating Handbook.

Inverter FailureIn the event of an AC bus or inverter failure, the red INVERTER annunciator on the Master Warning panel illuminates. In addition, the MASTER WARNING annunciator flashes until it is manually reset.If this occurs, select the other inverter with the inverter switch. Check AC voltage and frequency. The normal range is 396 to 404 Hz and 107 to 120 VAC.




Electrical System



Lighting Systems

LightingLighting systems on the Super King Air 200/B200 consist of interior and exterior lighting. This section provides a general overview of typical lighting found on the aircraft. Passenger compartment lighting systems may vary with equipment modifications and customer preference.

Interior LightingThe interior lighting systems consist of the cockpit, cabin, and emergency lighting.

Cockpit LightingThe cockpit lighting includes an overhead floodlight; individual post lights and indirect lighting on the instrument panel; edge lighting of the pedestal, overhead panel, side panels, and sub-panels.The following cockpit lighting controls are on the overhead panel. The MASTER PANEL LIGHTS ON/OFF switch controls the overhead

light control panel lights, fuel control panel lights, engine instrument lights, radio panel lights, subpanel and console lights, instrument lights and gyro instrument lights.

Separate rheostat switches control instrument indirect lights under the glareshield and overhead map lights.A pushbutton FREE AIR TEMP switch adjacent to the outside air temperature gauge on the left sidewall panel turns the gauge lights on and off.

Figure 5C-16: MASTER PANEL LIGHTS

ON/OFF Switch Figure 5C-17: FREE AIR TEMP Switch



Cabin LightingCabin lighting includes the following: two continuous rows of fluorescent lights along the cabin aisle headliner an adjustable eyeball light assembly above each passenger seat

and belted lavatory seat two lighted NO SMOKING/ FASTEN SEAT BELTS signs overhead threshold door light baggage compartment light airstair door steps entry floodlight

A CABIN three-position START/ BRIGHT-DIM-OFF switch on the copilot’s subpanel controls the cabin fluorescent lights. The adjacent NO SMOKE & FSB/OFF/FSB switch controls the NO SMOKING/FASTEN SEAT BELT signs and their accompanying chimes. In the up position, the switch illuminates both signs; in the down position, it illuminates only the FASTEN SEAT BELT sign.When the master switch in the cockpit is on, each passenger can control the individual reading light above each seat with a pushbutton switch adjacent to the light assembly.The threshold light is forward of the airstair door at floor level. An aisle light is at floor level aft of the spar cover. A switch adjacent to the threshold light controls both these lights as well as the air-stair entry light. These lights extinguish when the airstair door is closed, regardless of the switch position. To illuminate the baggage area with the airstair door closed, press the ON/OFF pushbutton on the light assembly.

Figure 5C-18: START/ BRIGHT-DIM-OFF

Switch Figure 5C-19: Spar Cover

Electrical System



A two-position switch just inside the airstair door, aft of the door frame, controls the baggage compartment light. The Hot Battery bus powers the baggage compartment light.

Figure 5C-20: Two-Position Switch

Power SourcesAlthough wiring variations exist between aircraft, typically, the No. 1 Dual-Fed bus powers the overhead and side panel lights, instrument indirect lights, and flight and gyro instrument lights. The No. 2 Dual-Fed bus powers the cabin fluorescent and reading lights, no smoking/fasten seat belt signs, subpanel and console lights, and avionic and engine instrument lights.

Exterior LightingThe exterior lighting system includes the landing, taxi, wing ice, navigation, recognition, rotating beacon, wing-tip, and tail strobe lights. Appropriately labeled switches for these lights are on the pilot’s subpanel.Optional tail floodlights in the horizontal stabilizer illuminate both sides of the vertical stabilizer. If these lights are installed, a TAIL FLOOD/OFF switch is on the pilot’s subpanel.

Figure 5C-21: TAIL FLOOD/OFF Switch



In addition, an optional flush-mounted floodlight in the lower left wing may be installed. This entry light illuminates the area around the airstair door for passenger convenience at night.

Figure 5C-22: Flush-Mounted Floodlight

Power SourcesAlthough wiring variations vary between aircraft, typically, the No. 1 Dual-Fed bus powers the beacon lights, strobe lights, and the left landing light.The No. 2 Dual-Fed bus powers the recognition light, navigation light, ice lights, taxi light, and right landing light.

Electrical System



Data Summaries

Electrical System

DC SystemPower Sources Battery

Two starter/generators 250A (STD) External power unit

Control Battery switch Start switch Generator switch

Two-position on BB-88 and priorThree-position on BB-89 and subsequent

Distribution Hot Battery bus Battery relay Main Battery bus Left and right start relays Isolation bus Left and right Generator buses Nos. 1 through 4 Dual-Fed buses Avionics buses No. 1, No. 2, and optional No. 3

Monitoring L/R DC voltmeter L/R GEN annunciators Battery charge annunciators EXT POWER annunciators Volt/loadmeter

Protection Voltage regulator Generator paralleling Reverse current sensing and control Over-voltage protection Over-excitation protection Under-excitation protection GPU reverse polarity sensing Generator buses 325A isolation limiters CBs and current limitersDual-Fed buses 60A current limiters and/or 50A CBs and diodes (70A)Hot Battery Bus Fuses or CBs



Electrical Systems (Continued)

AC SystemPower Sources No. 1 Inverter

No. 2 Inverter Control INVERTER NO. 1/OFF/NO. 2 switchDistribution Generator buses

Inverters No. 1 and No. 2 26V AC bus 115V AC avionics

Monitoring INVERTER inundator Volt/Frequency meter

Protection DC to inverter 50A or 60A current limiters Inverter output Fuses and CBs

Fire Protection



5DContentsFire Protection

Figure: Fire Detection System ....................................................5D-4Fire Detection System

Description .............................................................................................5D-5Operation ................................................................................................5D-5

Fire Extinguishing SystemEngine Fire Extinguishing System (Optional) ....................................5D-7

Control ...............................................................................................5D-7Portable Fire Extinguishers ..................................................................5D-8

Figure: Fire Extinguishing Systems ............................................5D-9Preflight and Procedures

Preflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5D-11Fire Detection and (Optional) Extinguishing Test . . . . . . . . . . . . . . 5D-12

Detection System Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5D-13Extinguisher System Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5D-13

Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5D-13Engine Fire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5D-13Environmental or Electrical System Smoke or Fumes . . . . . . . . . . . 5D-14

Data SummaryFire Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5D-15




Fire Protection



Fire ProtectionThis chapter describes the Super King Air 200/B200 fire protection system.All models have an electrically powered engine fire detection system. An optional electrically actuated engine fire extinguishing system may be installed.Manually operated portable fire extinguishers are available in the cockpit and passenger cabin.



Fire Detection System

5

FIRE

DET

5 7.5

ICE VANE

CONTROL

FUEL CONTROL

HEAT

LEFT LEFT

RIGHT RIGHT OFF

TEST SWITCH FIRE DET & FIRE EXT

3 2

1

FIRE DETECTORS

CONTROL AMPLIFIERS

FIRE DETECTORS

32

1

L FUEL PRESS

L BL AIR FAIL

L GEN OVHT

INVERTERL ENG FIRE

L OIL PRESS

L CHIP DETECT

AP TRIM FAIL

AP DISC R BL AIR FAIL

R GEN OVHT R OIL PRESS

R CHIP DETECT

R FUEL PRESS

CABIN DOOR ALT WARN R ENG FIRE

APPROACH PLATE

BRT

PRESS TO TEST

PRESS

TO RESET

MASTER CAUTION

PRESS

TO RESET

MASTER WARNING

PRESS

TO RESET

PRESS

TO RESET

MASTER WARNING

MASTER CAUTION

FIRE DETECTION CB2

ANNUNCIATOR PANEL

1

3 TEST SWITCH

5 7.5

Fire Protection



Fire Detection System

DescriptionThe electrically powered fire detection system provides immediate warning in the event of fire in either engine compartment.The detection system consists of the following: three photoconductive cells for each engine control amplifier for each engine red L/R ENG FIRE annunciators test switch on the copilot’s left sub-panel FIRE DET CB on the right side panel.

OperationOn S/Ns BB-2 through BB-1438, BB-1440 through BB-1443, BT-1 through BT-34, BL-1 through BL138 and BN-1 through BN-4, the fire detection system provides warning of an engine compartment fire.Three photoconductive cells, sensitive to infrared rays, are positioned in each engine compartment to receive direct and reflected rays. The cells, which cover the entire engine compartment, emit electrical signals based on infrared intensity and radiation striking the cell. The rate of heat increase and heat level do not activate the cell’s electrical signal.

Figure 5D-1: Photoconductive Cells



A control amplifier mounted on a panel on the forward side of the forward pressure bulkhead contains a relay that closes when the electrical signal strength reaches a preset alarm level. When the relay closes, the appropriate warning annunciator (FIRE L/R ENG or L/R ENG FIRE) illuminates.When the fire has been extinguished, cell output voltage drops below the alarm level. The control amplifier automatically resets. No manual resetting is required to reactuate the system.A test switch on the inboard side of the copilot subpanel and the FIRE DET circuit breaker on the right panel are also part of the system.On S/Ns BB-1439, BB-1444 and subsequent, BT-35 and subsequent, BL-139 and subsequent and BN-5 and subsequent, the main component of the system is a temperature sensing element. This element is routed through the three sections of each nacelle; it terminates in a responder unit.

NTTe: This system is subject to false alarms; moisture in circuitry connections or sun rays (as when the aircraft flies toward a rising or setting sun) can trigger the photocell detectors.Check for secondary indication of fire (temperatures, fuel flow, visual indications of fire, etc.) prior to engine shutdown or prior to activation of the extinguishing system.

The sensing element of corrosion-resistant, stainless steel, is hermetically sealed. The outer tube is filled with an inert gas while the inner core is filled with an active gas. The gases form a pressure barrier to keep the contacts of the responder integrity switch closed for continuity test functions of the fire alarm.The responder unit, attached to the engine mount in each engine accessory section, transmits signals to the L/R annunciator fault detection printed circuit cards and the L/R fire extinguisher control switches (if installed).Simultaneously, the red L or R ENG FIRE warning annunciator in the warning annunciator panel, located on the center glareshield, illuminates the pilot’s and copilot’s red MASTER WARNING lights. Those lights will continue to flash, even if the fire is extinguished.

Figure 5D-2: Temperature Sensing Element

Fire Protection



Fire Extinguishing System

Engine Fire Extinguishing System (Optional)The optional engine fire extinguisher system has a pyrotechnic cartridge on the extinguisher bottle in each wheel well. When fired, the extinguisher bottle discharges its entire contents into the engine nacelle and must be recharged for subsequent use. The system cannot be cross-fed.

Figure 5D-3: Pyrotechnic Cartridge

ContuolThe Hot Battery bus powers the fire extinguisher control switches; therefore, the extinguishers can be discharged in flight with no other electrical buses operating or on the ground with engines off. The detection system, however, is operative only when the No.1 Dual-Fed bus is operating.On S/Ns BB-2 through BB-1485 except BB-1484, BL-1 through BL140, BT-1 through BT-38 and BN-1 through BN-4, each push-to-actuate extinguisher switch has three indicator lights that illuminate and are marked as follows: L or R ENG FIRE-PUSH TO EXT (red), fire in the engine compartment D (amber), system discharged and the cartridge is empty OK (green), provided only for the preflight test function.



On S/Ns BB-1484, BB-1486 and subsequent, BL-141 and subsequent, BT-39 and subsequent and BN-5 and subsequent, each push-to-actuate extinguisher switch has three indicator lights that illuminate and are marked as follows: EXTINGUISHER PUSH (yellow), fire in the engine compartment DISCH (yellow), system discharged and the cartridge is empty OK (green), provided only for the preflight test function.

The system is discharged by raising the safety-wired clear plastic cover and pressing the face of the red L or R ENG FIRE-PUSH TO EXT lens. The D lens then illuminates and remains illuminated, regardless of battery switch position, until the pyrotechnic cartridge is replaced.

Figure 5D-4: Push-To-Actuate Extinguisher Switch

Portable Fire ExtinguishersTwo portable fire extinguishers, one in the cockpit and one in the cabin, are mounted in quick-release brackets. The bottles are normally filled with Halon. Prior to flight, verify that each is fully charged according to the pressure chart on the bottle and that it is securely stowed in its bracket.

Figure 5D-5: Portable Fire Extinguisher

Fire Protection



Fire Extinguishing Systems

1

APPROACH PLATE

BRT

L CHIP DETECT

PRESS TO TEST L BL AIR FAIL

L GEN OVHT

INVERTERL ENG FIRE

L OIL PRESS AP TRIM FAIL

AP DISC R BL AIR FAIL

R OIL PRESS

R CHIP DETECT

CABIN DOOR ALT WARN

L FUEL PRESS R FUEL PRESS

R ENG FIRE

R GEN OVHT

L ENG FIRE PUSH TO EXT

D OK

PRESS

TO RESET

PRESS

TO RESET

MASTER WARNING

MASTER CAUTION

PRESS

TO RESET

MASTER CAUTION

PRESS

TO RESET

MASTER WARNING

R ENG FIRE PUSH TO EXT

D OK

RIGHT MONITOR MODULE

TEST SWITCH

FIRE EXTINGUISHER SUPPLY CYLINDER

ANNUNCIATOR PANEL

2

1

3

OFF TEST SWITCH

FIRE DET & FIRE EXT

3 2

1

EXT

LEFT

RIGHT

2

LEFT MONITOR MODULE

FIRE EXTINGUISHER SUPPLY CYLINDER

3

EXPLOSIVE SQUIB

PRESSURE GAGE




Fire Protection




PreflightThe following is an overview of preflight procedures that apply to the fire protection system. Please see the Preflight and Expanded Normal Procedures chapters for details and specific checklists.If extinguishers are installed, verify during the preflight inspection that the fire extinguisher gauge in each wheel well indicates a proper charge.The fire extinguisher bottle contains 2 1/2 lbs of bromotrifluoromethane (CBrF3) pressurized to 450 PSI at about 70°F (see Fire Extinguisher Cylinder Pressure Limits at right). If a small leak allows the extinguishing agent to slowly leak from the bottle, there is no cockpit indication that contents are depleted. Checking the pressure indication on the gauge during preflight is the only method of determining the unfired bottle’s discharged state.Ensure the cylinder is checked by qualified maintenance personnel and recharged if necessary.Check the portable fire extinguishers in the cockpit and cabin for proper stowage and charge, based on the bottle placard.During the interior preflight inspection, perform the fire protection system check with the TEST SWITCH on the copilot’s subpanel.

Figure 5D-6: Fire Extinguisher Gauge

Fire Extinguisher Cylinder Pressure Limits

Temp. °F Indicated Pressure

-40 . . . . . . . . . . .190 to 240 PSI-20 . . . . . . . . . . .220 to 275 PSI0 . . . . . . . . . . .250 to 315 PSI

20 . . . . . . . . . . .290 to 365 PSI40 . . . . . . . . . . .340 to 420 PSI60 . . . . . . . . . . .390 to 480 PSI80 . . . . . . . . . . .455 to 550 PSI

100 . . . . . . . . . . .525 to 635 PSI120 . . . . . . . . . . .605 to 730 PSI140 . . . . . . . . . . .700 to 840 PSI



Fire Detection and (Optional) Extinguishing TestOn S/Ns BB-2 through BB-1438, BB-1440 through BB-1443, BT-1 through BT-34, BL-1 through BL138 and BN-1 through BN-4, a rotary TEST SWITCH on the copilot’s left subpanel (FIRE DET & EXT) has six positions, if the aircraft has the optional fire extinguishing system installed.The positions are: left of the switch – L and R EXT right of the switch – 3, 2, 1 (test positions) OFF – down.

On aircraft without engine extinguishers, the switch does not have the two L and R EXT positions.On S/Ns BB-1439, BB-1444 through BB-1462, if the optional fire extinguishing system is installed, the switch is placarded TEST SWITCH FIRE DET & EXT, OFF, EXT L-R and DET L-R. If the fire extinguishing system is not installed, the switch is placarded TEST SWITCH FIRE DETECT, OFF, LH and RH.On S/Ns BB-1463 and subsequent, BT-35 and subsequent, BL-139 and subsequent and BN-5 and subsequent, if the optional fire extinguish ing system is installed, the switch is placarded TEST SWITCH ENG FIRE SYS, OFF, EXT L-R and DET L-R. If the fire extinguishing system is not installed, the switch is placarded TEST SWITCH ENG FIRE SYS, OFF and DET L-R.

NTTe: The fire detection system does not test the photocell detectors in the engine.

Figure 5D-7: Rotary Test Switch

Fire Protection



DrtrctFon Systrm TrstOn the right side of the test switch are three test positions: 3, 2, and 1. A normal test begins with the rotary switch pointer in the six o’clock position.When the switch is rotated counterclockwise to each of these three positions, the output voltage of the corresponding detector in each engine compartment increases to a level sufficient to signal the amplifier that a fire is present.With the switch at one of the three positions, the following annunciators illuminate: left and right red MASTER WARNING flashers red L and R ENG FIRE annunciators red L and R ENG FIRE – PUSH TO EXT lenses on the fire extinguisher

activation switches (if extinguishers installed).The system may be tested on the ground or during flight. Place the test switch in all three positions to verify that the circuitry for all six fire detectors is functional. If any one or more of the fire detection system annunciators fail to illuminate during the test at each test position, there is a malfunction in one or both of the two detector circuits (one in each engine) being tested by that particular position of the test switch.

TxtFnigFshru Systrm TrstIf extinguishers are installed, the fire extinguisher system test functions in the rotary TEST SWITCH test the circuitry of the fire extinguisher pyrotechnic cartridges.During preflight, rotate the switch counterclockwise to each of the two positions to the left of the switch (R and L EXT). Verify the illumination of the amber D lens and the green OK lens on each fire extinguisher activation switch on the glareshield. Illumination of the green OK light indicates that the detector circuitry and squib-firing circuitry are operational and that the squib is in place; it does not indicate bottle pressure. Illumination of the amber D light only tests the bulb.

Emergency ProceduresThis section presents a brief overview of emergency procedures that apply to the fire protection system. Please see the CAE King Air Operating Handbook for detailed checklists on emergency procedures.

TniFnr FurHistorically, engine fire on the Super King Air 200/B200 is rare. Illumination of the L or R ENG FIRE warning annunciator indicates a fire in the corresponding engine. On aircraft with engine fire extinguishers, the red L or R ENG FIRE-PUSH TO EXT lens illuminates as well.



Check for secondary indication of fire (temperatures, fuel flow, visual indications of fire, etc.) prior to engine shutdown or prior to activation of the extinguishing system. The check for secondary indications is necessary because other sources (e.g., sunlight or moisture in circuitry) can cause false alarms.On aircraft without engine fire extinguishers, select CUT OFF on the affected engine’s condition lever to shut down the engine, feather the propeller, and close the fuel firewall valve to stop combustible liquid flow. The fuel firewall valve is aft of the engine’s stainless steel firewall.On aircraft with fire extinguishers, follow the above procedure, then discharge the fire extinguisher cartridge. Raise the safety-wired clear plastic cover and press the face of the illuminated ENG FIRE-PUSH TO EXT lens. After the cartridge discharges, the D lens illuminates and remains illuminated, regardless of battery switch position, until the pyrotechnic cartridge is replaced.Follow the CAE Operating Handbook or AFM checklist procedures and land as soon as practical.

TnvFuonmrntal ou TlrctuFcal Systrm Smokr ou gmrsIn either situation, don crew masks (100% position) first. Give primary attention to flying the aircraft. Identify the source of the smoke or fumes. Gray or tan smoke and irritation of the nose and eyes identifies an electrical failure. An environmental system failure produces white smoke that is much less irritating to the nose and eyes.If necessary, descend and depressurize. Follow procedures in the CAE Operating Handbook or the AFM. After the aircraft is depressurized, open a storm window to facilitate smoke and fume removal.

Fire Protection



Data Summary

Fire Protection

Power Source No. 1 Dual-Fed bus – fire detection Hot Battery bus – optional fire extinguishing Portable fire extinguishers

Distribution Extinguisher bottle to corresponding engine (no crossfiring) Control TEST SWITCH FIRE DET (& EXT, if installed)

ENG FIRE-PUSH TO EXT (L/R) lens/switch (if installed) Monitor ENG FIRE (L/R) annunciators

Red ENG FIRE-PUSH TO EXT (L/R) lens (if installed)Amber D lens to confirm electrical wiring continuity (if installed)Green OK lens to test system (if installed)MASTER WARNING (L/R) flashers

Protection FIRE DET CB (5A) FIRE extinguisher fuse (prior to BB-1096) (if installed)FIRE EXTINGUISHER CB (5A) (BB-1096 and subsequent) (if installed)




Flight Controls 5E

King Air 200 5E-1December 2011


ContentsFlight Controls

Figure: Primary Flight Controls ................................................. 5E-5ScremaFSe: Rudder Boost System ............................................ 5E-6Figure: Trim Systems ................................................................ 5E-7

Primary Flight ControlsDescription ........................................................................................... 5E-9Ailerons ................................................................................................ 5E-9

Roll Trim ........................................................................................... 5E-10Elevators .............................................................................................. 5E-10

Pitch Trim ......................................................................................... 5E-11Rudder .................................................................................................. 5E-13

Rudder Trim (Servo) ........................................................................ 5E-13Rudder Boost ....................................................................................... 5E-14Pedal Adjustment Levers .................................................................... 5E-15Yaw Damper ......................................................................................... 5E-15Controls Gust Lock ............................................................................. 5E-16

ScremaFSe: Primary Flight Controls .......................................... 5E-17Secondary Flight Controls

Flaps ..................................................................................................... 5E-19Flap Control Lever ........................................................................... 5E-19Electrical Motor ................................................................................ 5E-19

Stall Warning ........................................................................................ 5E-21Preflight and Procedures

Preflight ................................................................................................ 5E-23Abnormal Procedures ......................................................................... 5E-23

Unscheduled Electric Elevator Trim ................................................. 5E-23Unscheduled Rudder Boost Activation ............................................ 5E-24Asymmetrical Flaps ......................................................................... 5E-24

Data SummaryFlight Controls ..................................................................................... 5E-25

Flap System ..................................................................................... 5E-25

King Air 2005E-2December 2011


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Flight Controls



Flight ControlsThe primary flight controls are manually operated by the pilot and copilot; they consist of the following: ailerons elevators rudder.

The secondary flight controls are the flaps. The inboard and outboard flap panels operate electrically.Trim tabs on the left aileron, the rudder, and one on each elevator provide for roll, yaw, and pitch trim. The pilot manually controls the tabs through drum-cable trim systems with jackscrew actuators. Electric trim is also provided for pitch.In addition to the primary and secondary flight controls, the following complete the flight controls system: stall warning autopilot (see Avionics chapter) rudder boost yaw damper.

NTTe: A fixed trim tab is located on right aileron.





5E-5For Training Purposes Only

Flight Controls

Primary Flight Controls

Aileron Control System

Elevator Control System

Rudder Control System


5E-6 For Training Purposes Only

Rudder Boost System

P SWITCH (60 PSI DIFFERENTIAL)

18 PSI PNEUMATIC PRESSURE

REGULATOR CHECK VALVE

CHECK VALVE

P3 AIR

N.O. LEFT INSTRUMENT AIR VALVE


15 PSI PRESSURE

REGULATOR FILTER

RIGHT RUDDER SERVO

LEFT RUDDER SERVO

N.C. N.C.

REGULATED BLEED AIR

UNREGULATED P3 AIR

N.O. RIGHT INSTRUMENT AIR VALVE

P3 AIR

TO LEFT RUDDER CABLE

TO RIGHT RUDDER CABLE

RUDDER BOOST


OFF

ENVIR OFF

INSTR & ENVIR OFF

BLEED AIR VALVES OPEN

LEFT RIGHT FROM NO. 1

DUAL-FED BUS

FROM NO. 2 DUAL-FED BUS 5A

5A

5A

RUDDER BOOST



Flight Controls

Trim Systems

Roll Trim (Aileron)

Pitch Trim (Elevator)

Rudder Trim




Flight Controls




DescriptionPrimary flight controls permit command of the aircraft through the roll, pitch, and yaw axes. The ailerons and elevators receive input from conventional push-pull control wheels interconnected by a T-bar. The rudder pedals control rudder displacement through direct linkup to the rudder.

AileronsAilerons on the outboard trailing edge of each wing provide roll control mechanically through the control wheel or electrically through the autopilot servo. Positive stops on the flight control surfaces limit travel.The control wheel provides conventional aileron control through a range of 15° ± 1½° up and 15° ± 1½° down.

Figure 5E-1: Ailerons



Roll TuFeA trim tab on the left aileron’s inboard trailing edge provides lateral trim capability. An AILERON TAB knob in the center pedestal allows manual adjustment of the roll trim. Around the knob, a dial indicator displays the position of the tab. A drum-cable system incorporating a jackscrew actuator moves the tabs to the selected setting. Double-ended, clevis rod assemblies remove joint free play.

Figure 5E-2: AILERON TAB Knob

ElevatorsThe elevators on the trailing edge of the T-tail horizontal stabilizer provide pitch control mechanically through fore and aft movement of the control column or electrically through the autopilot servo. The travel for the elevator is 20° +1° -0° up and 14° +1° -0° down.

Figure 5E-3: Elevators

Flight Controls



PFaSc TuFe

WARNRN Do not allow full travel of the pitch trim either mechanicaly, electricaly, or through action of the autopilot. This can result in damage or mis-rigging of the pitch trim system.

Each elevator contains a trim tab that is manually or electrically trimmed. Move the handwheel on the left side of the pedestal to manually control the tab. The handwheel is a conventional trim wheel that rolls forward for nose-down trim and aft for nose-up trim. The trim indicator displays trim position.The following control the electric elevator-trim system: an ELEV TRIM/ON/OFF switch on the pedestal a dual element THUMB switch on each control wheel a trim-disconnect switch on each control wheel a 5A PITCH TRIM CB in the flight group on the right side panel. The

No. 1 Dual-Fed bus provides primary electrical power.

Figure 5E-4: Trim Tab

With the ELEV TRIM switch ON, moving either the thumb switch forward (nose-down) or aft (nose-up) achieves pitch trim. The switch returns to the center (OFF) -position when released. The pilot’s thumb switch overrides the copilot thumb switch. Before takeoff, perform a check of both dual-element thumb switches by moving each of the four switch elements individually; no one switch element should activate the -system. Activating both elements on either control wheel results in trim movement.



Figure 5E-5: ELEV TRIM Switch

A bi-level, push-button, momentary-on, trim-disconnect switch inboard on the dual-element thumb switch is on the outboard grip of each control wheel.With each installation of an autopilot, pressing the switch to the first level disconnects the autopilot and the yaw damper system. Pressing the switch to the second level additionally disconnects the electric elevator-trim system.Without an autopilot, pressing the switch to the first level has no effect since the yaw damper switch on the pedestal controls the damper. The second level disconnects the elevator-trim system.Pressing the trim-disconnect switch with the ELEV TRIM switch ON illuminates a green ELEC TRIM OFF light on the center pedestal. This light notifies the crew of a disabled electric trim system. To reset the system, cycle the ELEV TRIM switch on the pedestal to OFF then ON.

NTTe: The manual-trim wheel allows trimming of the aircraft anytime (with or without electrical trim active).

NTTe: As a part of each pre-flight check, set pitch trim to zero, and check that the trim tab aligns with the trailing edge of the elevator. There is NO tolerance.

Figure 5E-6: Thumb Switch

Flight Controls



RudderA direct connect cable system from both sets of rudder pedals to the tail section drives the aircraft rudder. The rudder, at the trailing edge of the vertical stabilizer, provides directional control of the aircraft about the vertical axis. Full range of motion is 25° +1° -0° left and right of center. Rudder pedals or the yaw damper servos control the rudder. On auto-pilot equipped aircraft, the yaw damper utilizes the autopilot servos. On non-autopilot equipped aircraft, the yaw damper system utilizes the rudder boost servos.

Figure 5E-7: Rudder

Rgddru TuFe (ruvo)The rudder trim system reduces pilot input forces to aid the pilot in maintaining yaw control of the aircraft. The RUDDER TAB LEFT/RIGHT rotary-type knob on the center pedestal allows manual control of the rudder trim tab. A dial indicator around the knob displays the position of the tab.

Figure 5E-8: RUDDER TAB LEFT/RIGHT Rotary-Type Knob



Rudder BoostA rudder boost system aids the pilot in maintaining directional control if an engine fails and/or if there is a large variation of power between the engines. The rudder boost system consists of: a differential pressure switch an in-line air filter a 15 PSI pressure regulator two vented solenoid valves two pneumatic rudder servos.

The differential pressure switch, ΔP, receives bleed air pressure from both engines. The switch senses pressure changes and moves toward the low pressure side. Upon reaching a differential pressure of 60 ±5 PSI, the respective switch on the low pressure side closes. The switch provides power to activate the applicable solenoid valve; this allows regulated pneumatic bleed air pressure to actuate the rudder servo.An in-line air filter prevents contaminants from entering the pressure regulator and servos. The 15 PSI pressure regulator prevents high pressure damage to the servo’s diaphragms.To compensate for asymmetrical thrust, the pneumatic servos provide additional rudder control pressure.Electrical power for the system passes through a two-position toggle switch on the pedestal labeled RUDDER BOOST/OFF.The RUDDER BOOST/OFF toggle switch below the RUDDER TAB wheel on the center pedestal controls the rudder boost system by interrupting the 28 VDC to the differential pressure switch. The 5A RUDDER BOOST CB on the copilot’s CB panel protects the system. Select the switch to RUDDER BOOST prior to flight.Additionally, the selection of either or both BLEED AIR VALVES switches to INSTR & ENVIR OFF disengages the rudder boost system by opening the rudder boost relay.

Figure 5E-9: Rudder Servo Figure 5E-10: RUDDER BOOST/OFF

Toggle Switch

Flight Controls



Pedal Adjustment LeversEach set of rudder pedals individually adjusts through a lever on the side of the pedals. The crewmember adjusts the pedals to the desired position by unlocking the pedals with the foot. Return the lever to the locked position.

Yaw DamperThe yaw damper provides automatic stabilization about the yaw axis by controlling transient yaw motion. Use of the yaw damper system is required at altitudes above 17,000 ft. Deactivate the yaw damper for takeoffs and landings.The independent yaw damper system (no autopilot) consists of the following: a yaw sensor an amplifier a control valve.

A YAW DAMP switch adjacent to the RUDDER BOOST switch on the pedestal controls the yaw damper system. Regulated air pressure from the control valve utilizes the same pneumatic servos used for the rudder boost system when no autopilot is installed. During takeoff or landing with the YAW DAMP switch ON, a safety switch on the left landing gear interrupts power to the system to prevent operation of the yaw damper. The yaw damper system drives the right and left rudder boost servos to deflect the rudder and stabilize the yaw axis of the aircraft.On autopilot-equipped aircraft, the yaw damper system is a part of the autopilot operation. The YD ENGAGE/DISENGAGE controls the yaw damper system. The autopilot compensates for yaw characteristics through the pilot’s turn and bank instrument or a rate gyro. and signals the autopilot servos adjusting the rudder. For specific details, refer to the appropriate Air-plane Flight Manual Supplement for autopilot installation.

Figure 5E-11: YD ENGAGE/DISENGAGE

Controls Figure 5E-12: Rate Gyro



Controls Gust Lock

CWAUNIR Remove the rudder lock before towing.

The controls gust lock secures movable control surfaces. When installed, the lock holds rudder and ailerons in neutral and down elevator. Check the controls for proper movement and freedom from binding prior to flight.

Figure 5E-13: Controls Gust Lock



Flight Controls


FLAPS

20A

FLIGHT 2LH HOUR LIMIT SWITCHES GEN METER

BUS UPSPLIT FLAP FLAP

SWITCHES DYNAMICAPPROACH BRAKE RELAY

5A

1 4 DOWN

3POSITION

TRANSMITTER

TO LANDING GEAR WARNING SYSTEM

POSITION INDICATOR

LIFT COMPUTER

2 3 41




Flight Controls



Secondary Flight Controls

FlapsTwo flaps installed on each wing are operated by an electric motor-driven gearbox mounted on the forward side of the rear spar at the centerline of the airplane. The gearbox drives four flexible driveshafts, each connected to a jackscrew at each flap. The landing gear warning system is affected by the position of the flaps. Any time the flaps are set beyond the approach position and the landing gear is not extended, the landing gear warning horn will sound until the flaps are retracted, or the gear extended. Otherwise, in the Up or Approach positions the gear warning horn may be silenced.

lmp Conauol LrvruAny of the three flap positions, UP, APPROACH or DOWN may be selected by moving the flap selector lever up or down to the selected detent position indicated on the pedestal. The lever position detent is provided to select flaps UP (0°), 40% flap extension (14°) for APPROACH position, or 100% flap extension (35°) for DOWN.

Figure 5E-14: Flap Control Lever

TlrSauFSml MoaouThe motor incorporates a dynamic braking system, through the use of two sets of motor windings, which helps to prevent over travel of the flaps.

GearboxThe gearbox drives four flexible driveshafts connected to jackscrews at each flap.



Split Flap ProtectionA safety mechanism is provided which will disconnect power to the electric motor in the event of any type of failure, which causes any flap to be 3 to 5 degrees out of phase with the adjacent flap panels.

Figure 5E-15: Split Flap Protection

Flap Position IndicationWing flap position is marked in percent of flap deflection and is located on the forward throttle quadrant assembly.The indicator is actuated by a potentiometer driven by the right inboard flap.

Figure 5E-16: Flap Position Indicator

Flight Controls



Stall Warning

WARNRN The formation of ice at the stall transducer vane results in erroneous indications.

The stall warning system provides an audible warning to notify the crew of an impending stall. The system consists of a lift transducer, a lift computer, a warning horn, and a test switch.Aerodynamic pressure strikes the leading edge of the left wing’s lift transducer vane, which measures the Aircraft’s Angle-of-Attack (AOA). The lift transducer signals the lift computer if a stall is imminent. The lift computer processes signals from the lift transducer and the flap position sensors prior to triggering the warning horn in the pilot’s headliner.During flight, the lift transducer provides signals to the lift computer, which adjusts the stall warning limits to the following: 5 to 13 Kts above stall with flaps retracted 5 to 12 Kts above stall with flaps at 40% (APPROACH) 8 to 14 Kts above stall with flaps fully extended.

The electrically heated mounting plate and vane of the lift transducer prevents ice formation down to -22°F. In flight, a supply of 28 VDC warms the vane and plate. For ground operation, electrical power activated by the LH squat switch drops to 12 VDC for heating the plate and vane.The STALL WARNING TEST switch on the copilot subpanel tests the operation of the stall warning system with the aircraft on the ground. An electromagnetic circuit within the transducer enables the vane to simulate a stall condition and sounds the horn for testing purpose. The stall warning system may be un-reliable during operations in icing conditions with accumulations of ice on airframe surfaces.

Figure 5E-17: Left Wing’s Lift Transducer

Vane Figure 5E-18: STALL WARNING TEST

Switch




Flight Controls




PreflightDuring the external preflight inspection, check all control surfaces for freedom of movement and general security.During the cockpit inspection, ensure the flap lever agrees with flap position and the elevator trim is within the takeoff range; remove the control lock (see Preflight chapter for details).During engine run-up, perform a preflight check of the rudder boost system by retarding the throttle of one engine to idle and advancing the power lever on the opposite engine until obtaining a bleed air split between the engines of 60 ±5 PSI. The differential pressure switch then closes and activates the rudder boost system. The respective rudder pedal should move (i.e., left power lever forward, left rudder pedal moves forward) to indicate the switch is closed. Pedal movement verifies proper system operation. Repeat the check with opposite power settings to check for movement of the opposite rudder pedal.Prior to takeoff check all controls, rudder, aileron and elevators for freedom of movement and surface agreement.

Abnormal ProceduresThis section provides a brief discussion of flight controls abnormal procedures.For a list of specific procedural steps, please refer to your CAE Operating Handbook.

UnsScrdglrd TlrSauFS Tlrvmaou TuFe

WARNRN Determine and correct the electric trim system malfunction prior to reactivating the system.

If an unscheduled electric elevator trim situation occurs, the nose of the aircraft ascends or descends depending on direction of trim runaway. Maintain the attitude of the aircraft with the elevator control and press the control wheel disconnect switch down to the second level. Manually retrim the aircraft and select the ELEV TRIM control switch on the pedestal off.

NTTe: Pressing the control wheel disconnect switch to the second level disengages both the autopilot and yaw damper.



UnsScrdglrd Rgddru Boosa ASaFvmaFonAn unscheduled rudder boost activation requires the RUDDER BOOST switch be moved to OFF. If the condition persists, adjust the rudder trim manually, select either BLEED AIR VALVE to INSTR & ENVIR OFF.An unscheduled rudder boost activation accompanied by an increase in ITT and a decrease in torque on the engine opposite to the direction of the rudder deflection indicates a bleed air line rupture on the side of the abnormal engine instrument readings. Placing the bleed air valve switch on that side to INST & ENVIR OFF should correct the rudder boost problem.

NTTe: During single engine operations, the rudder boost switch may need to be placed in OFF position to prevent unnecessary (or intermittent) rudder boost - activation at certain power settings.

AsyeerauFSml lmpsThe asymmetrical flap protection functions when any segment’s position varies 3° to 6° from the adjacent side segment. The safety mechanism disconnects power to the electric flap motor in the event of a malfunction.

Flight Controls



Data Summary

Flight Controls

lmp ysarePower Source Electric motor and control No. 3 Dual-Fed busControl Flap handle Monitor Flap position indicator Protection Power and Control CBs

Split flap protection Limit switches




Fuel Systems 5F

King Air 200 5F-1December 2011


ContentsFuel System

SchematSc: Fuel System (King Air B200) .....................................5F-5SchematSc: Fuel System (King Air 200) ........................................5F-5

Fuel StorageDescription . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5F-9Wing Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5F-9Nacelle Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5F-9Auxiliary (Center Section)Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-10Tip Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-11Vent System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-11Fuel Quantity Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-12

Fuel Probes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-12Fuel Quantity Indicators . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-13

Fuel Distribution/ControlWing Tanks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-15Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-15

Standby Boost Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-15Standby Boost Pump Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-17Jet Transfer Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-17Defueling Adapter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-17Firewall Shutoff Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-18Fuel Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-18Crossfeed Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-18Crossfeed Valve Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-19Motive Flow Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-19Fuel Drains . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-19Fuel Drain Collector System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-20Fuel Drain Purge System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-21

Servicing and ProceduresRefueling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-23Draining . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-23Defueling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-23Additive Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-24

King Air 2005F-2December 2011


Preflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-24Abnormal Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-25

FUEL CROSSFEED Annunciator Illuminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-25Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-25Crossfeed (One Engine Inoperative Operation) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-25FUEL PRESS Annunciator Illuminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-26

Data SummaryFuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-27

Main Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-27Auxiliary Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5F-27

Fuel Systems



Fuel SystemThe King Air 200/B200 has separate fuel systems connected by a valve-controlled crossfeed line. The fuel systems consist of the following major subsystems: main fuel storage system includes the wing and nacelle tanks, and

center section includes the auxiliary tanks. The fuel vent and fuel quantity indicating systems are part of the storage system.

fuel distribution system that includes the pumps, valves, and plumbing required to move fuel through the aircraft to the engines. Fuel filtration is part of the fuel distribution system.





5F-5For Training Purposes Only

Fuel Systems

Fuel SystemKing Air B200

0

2

4 6 8

12

14

F UE L

F UE L C ONT R OL UNIT

P B LE E D AIR LINE

E NG INE F UE L P UMP

F UE L HE AT E R

AIR F I LTE R

F UE L F LOW T R ANS MIT T E R AND INDIC ATOR

LE F T F UE L P R E S S UR E S WIT C H

F UE L C ONT R O L P UR G E VA LVE

G R AV IT Y F LOW C HE C K VALVE

T R ANS F E R C ONT R OL MOT IV E F LOW VA LV E

P R E S S UR E S WIT C H F OR LE F T NO F UE L T R ANS F E R LIG HT ON F UE L PANE L

E NG INE F UE L MANIF OLD

P R E S S UR E TANK

F IR E WALL F UE L F ILT E R

DR AIN VA LV E

E NG INE DR IV E N B OOS T P UMP

S TANDB Y B OOS T P UMP

S T R AINE R AND DE F UE LING DR AIN VA LV E

V E NT F LOAT VALV E

NAC E LLE TANK

C R OS S F E E D VA LV E

LE F T F UE L QUANT IT Y T R ANS MIT T E R

F IR E WALL S HUT OF F VA LV E

F 13 G A L

35 G A L

40 G A L

25 G A L 23 G A L

AIR INLE T

V E NT F L OAT VALVE

DR AIN

DR A IN

R E C E S S E D V E NT

HE AT E D R AM V E NT

F LAME AR R E S TOR

DR AIN

F

79 G A L

T R ANS F E R J E T P UMP

S T R AINE R , DR AIN & F UE L S WIT C H

F UE L UNDE R P R E S S UR E

STAT IC P R E S S UR E

J E T P UMP P R E S S UR E

F UE L IN S T OR AG E

F UE L F ILLE R

F UE L P R OB E S

F

3

F IR E WA L L S HUT OF F VA LVE

S TANDB Y P UMP ON

OF F

P R IOR TO B B 1098

L F UE L P R E S S

5A 5A

10A

5A

5A

5A

INT E G R A L (WE T C E L L ) A UXIL IA R Y

5A

136 G A L TO TA L 57 G A L

F R OM R IG HT WING

NO. 3 DUA L -F E D B US

HOT B AT T B US


5F-6 For Training Purposes Only



5F-7For Training Purposes Only

Fuel Systems

Fuel SystemKing Air 200

F R OM LE F T

WING

F IR E WA L L S HUT OF F VA LV E

S TANDB Y P UMP ON

OF F

P R IOR TO B B 1098

F R OM NOZZLE MANIF OLD

F LOA T S WIT C H

F UE L DR AIN C OLLE C T OR TANK

F UE L DR AIN R E T UR N P UMP

R IG HT F UE L QUANT IT Y T R ANS MIT T E R

57 G A L


S TAT IC P R E S S UR E

J E T P UMP P R E S S UR E

F UE L IN S T OR AG E

F UE L F ILLE R

F UE L P R OB E S

HOT B AT T B US

5A 5A

10A

5A

5A

5A NO 4 DUA L -F E D B US

136 G A L TO TA L

5A

F

F

F

4 6 8

F UE L

0

2

14

DR A IN

DR A IN

40 G A L 13 G A L

35 G A L 25 G A L 23 G A L

79 G A L

R F UE L P R E S S

12


5F-8 For Training Purposes Only


Fuel Systems



Fuel Storage

DescriptionFive wing tanks and the nacelle tank comprise the main fuel system: the wing center section tank is referred to as the auxiliary fuel system.

Wing TanksThe main fuel system in each wing consists of two wing leading edge bladder tanks, two box section bladder tanks, one integral tank, and a nacelle tank. All of the wing tanks interconnect by gravity feed lines and flow into the nacelle tank. The total usable fuel capacity of each wing and nacelle main system is 193 U.S. gallons. Fuel for each engine normally is pumped directly from its respective nacelle tank.The filler cap for the main fuel system is on the leading edge of the wing near the wing tip. An anti-siphon valve is installed at each filler port.Primary venting for all tanks is the air inlet on the underside of each wing tip. In addition, a heated vent and a NACA vent provide pressure venting for the wing tanks. The NACA vent also provides overboard porting of fuel if overexpansion occurs.

Figure 5F-1: Filler Cap

Nacelle TanksA flapper-type check valve between the nacelle tank and the wing tank allows fuel to gravity flow freely from the wing tank to the nacelle tank, but not back into the wing tank.A standby electrically-powered boost pump in each nacelle tank is available to move fuel directly to its respective engine if the engine-driven boost pump fails.



For single-engine operation, each nacelle tank connects to the fuel system on the opposite side through a crossfeed line. Fuel in each main fuel system is available to either engine. The system is not designed to transfer fuel from tanks on one side to tanks on the other side.

Auxiliary (Center Section) TankThe auxiliary fuel tank consists of a center section tank in each wing root. Because a portion of the center section tank is lower than the wing tanks and nacelle tank, a jet transfer pump in the sump of the center section tank moves fuel to the nacelle tank. Each auxiliary tank has its own filler opening, and each tank has a usable capacity of 79 U.S. gallons.If the float switch in the sump of the auxiliary tank senses any quantity of fuel in the tank and there is adequate fuel pressure (motive flow) to operate the jet pump, that fuel automatically transfers to the nacelle tank. When the float switch indicates empty, transfer automatically is discontinued (the motive flow valve closes).If the tank has fuel but fails to transfer because of a system discrepancy, an amber NO TRANSFER light on the fuel panel light illuminates. Switching on the fuel management panel offers a manual override which, when selected, bypasses normal sensing and applies power to the motive flow valve to allow the jet pump to resume transfer. The override position should be manually de-selected when the auxiliary tank has been transferred.

Figure 5F-2: Auxiliary Tank

Fueling LimitationDo not fuel auxiliary tanks unless the main tanks are full .

Fuel Systems



Tip TanksAn optional wing-tip installation includes a 53 U.S. gallons tank on each wing tip. The wing-tip installation increases total fuel system capacity to 650 U.S. gallons (4,355 lbs (1,975.3 kg)).Vent and fuel lines connect the tip tanks to the outboard wing tank on their respective side; fuel transfer is accomplished by gravity flow. An anti-siphon fuel filler cap is integral to the tank, and a float-type fuel quantity transmitter is in each tank. An additional set of fuel quantity gauges, labeled TIP TANK FUEL RH and LH, is on the fuel panel next to the standard gauges.

DR A IN V E NT F L OAT VALVE

A IR INL E T

F UE L QUA NT IT Y T R A NS MIT T E R

F IL L E R C A P

C HE C K VALVE

Figure 5F-3: Tip Tanks

Vent SystemThe fuel system vents through an air inlet on the underside of each wing tip and through a recessed NACA vent, coupled to a heated external vent, on the underside of the wing adjacent to the nacelle. One vent is recessed to prevent icing; the external vent is heated to prevent icing. Each vent serves as a backup if one becomes plugged. In each fuel system, the tanks are cross-vented with one another.

Figure 5F-4: Vent System



If the main tanks are full and the auxiliary tank is empty, fuel vents into the auxiliary tank when it expands from heating. If all the tanks are full, fuel expands overboard through the heated vent or NACA port.

Fuel Quantity IndicatingThe fuel quantity system is a capacitance gaging system with one quantity indicator per wing. Gauges normally show main system quantity; a momentary toggle switch allows selection of auxiliary fuel quantity. The main fuel system has a total capacity of 390 U.S. gallons and a maximum usable fuel quantity of 386 U.S. gallons. The auxiliary fuel system has a total capacity of 160 U.S. gallons with a maximum usable quantity of 158 U.S. gallons. The fuel quantity gauges and the engine fuel flow indicators read in pounds times 100 (e.g., at 6.7 lbs (3.03 kg) per U.S. gallon, 2,586.2 lbs (1,173.08 kg) of usable fuel are available in the main system; 1,293.1 lbs (586.5 kg) per side).The fuel quantity indication system automatically compensates for fuel temperature-density variations.

Figure 5F-5: Fuel Quantity Indicator

Fuhl ProbhsEach side of the aircraft has an independent fuel quantity system. A series of eight fuel probes supplies information to each fuel quantity indicator through the FUEL QUANTITY selector switch. These capacitance type tubular probe systems read the total capacitance (the sum of the eight tubular probes), which varies proportionally to quantity of fuel present.Openings in the probe allow fuel to flow between the electrodes. Capacitance is a function of plate (electrode) area, plate separation, and fluid dielectric constant. The fluid constant will either be air or fuel, and the total capacitance, or ability to store an electrostatic charge, can, therefore, be calculated precisely. Since the fluid density is measured rather than fluid quantity, these probes provide protection against contamination from fluids other than petroleum products. Water, for example, with a dielectric constant of 80, will have a noticeable affect on fuel gauge performance.

Fuel Systems



Fuhl Qumnatay IndtSmaorsThe left fuel quantity indicator, on the fuel management panel, indicates the amount of fuel remaining in the left side fuel system tanks when the FUEL QUANTITY select switch is in the MAIN (upper) position, and the amount of fuel remaining in the left side auxiliary tank when the FUEL QUANTITY select switch is held in the AUXILIARY (lower) position. The right fuel quantity has duplicate functions. The gauges are marked in pounds times 100.An electronic circuit in the system processes the signals from fuel quantity (capacitance) probes in the fuel cells for an accurate readout (±3%) by the fuel quantity indicators.




Fuel Systems



Fuel Distribution/Control

Wing TanksFuel flow from each wing tank system is automatic without pilot action. The wing tanks gravity feed into the nacelle tank through a line extending from the aft inboard wing tank to the forward side of the nacelle tank. A flapper-type check valve in the end of the gravity feed line prevents any backflow of fuel into the wing tanks.

Components

amndby Boosa PuepEach fuel system has a submerged standby boost pump in the bottom of the nacelle tank. This pump can supply a pressure head of about 30 PSI to the engine-driven main fuel pump if the engine-driven boost pump fails. These electrically-driven boost pumps are submerged rotary vane-type impeller pumps that provide low pressure fuel for the following functions: to serve as a backup, in case the engine-driven fuel boost pump fails to supply pressure to engine-driven fuel boost pump when using

aviation gasoline at altitudes above 20,000 ft to provide fuel to the opposite engine in crossfeed operations.

Standby pumps are powered by either the Hot Battery bus or the No. 3 Dual-Fed bus (left) or the No.4 Dual-Fed bus (right); two 5A fuses on the Hot Battery bus and a single 10A CB on the fuel panel provide protection. Some later S/Ns replace the fuses with a 10A CB. On S/N 1096, 1098 and subsequent, the standby pumps are no longer powered by the Hot Battery bus.Check the standby boost pumps prior to flight. If an engine-driven fuel pump fails, its respective FUEL PRESS annunciator illuminates as fuel pressure decreases below 10 (+1) PSI. The light should extinguish when the standby boost pump for that side is switched ON.



Auxiliary Fuel Transfer System

PR

E

S S T O T ES

T

N O

3

O R

4

D U A L

F E D

B U S

REFSNART XUAS WIT C H MOT IV E F LOW

AUX P R E S S UR E NO T R ANS

KNAT XUAS WIT C H

5A

AUX R E T UR N

OV E R F LOW)

TO E NG INE

NO. 3 OR 4 E NG INE -DR IV E N

LINE (NAC E LLE

F UE L P R E S S WA R N

OV E R R IDE

A UT O

F LOAT S WIT C H

E MP T Y

NOT E MP T Y

T R ANS F E R LIG HT

N.C .

MOT IV E F LOW VALVE F LOAT S WIT C H

A UXIL IA R Y TA NK


F UE L P R E S S UR E 5A 5A

A UXIL IA R Y

S WIT C H

5A P R E S S WAR N

S T B Y 10A

P UMP S W

S T R A INE R F R OM C R OS S F E E D

S WIT C H F R O M WING TANK

NA C E L L E TA NK

T R A NS F E R CONT R OL A ND

TI ME R (P.C .B)

HOT B AT B US

1

DUAL-F E D B US B OOS T P UMP


STAT IC P R E S S UR E

N O

3

O R

4

D U A L

F E D

B U S

AUX T R ANS F E R S WIT C H MOT IV E F LOW

F UE L IN S TOR AG E

1 ON B B 1097, 1095 AND P R IOR

F UE L P R E S S WA R N

OV E R R IDE

A UT O

AUX TANK F LOAT S WIT C H

E MP T Y

NOT E MP T Y

P R E S S UR E S WIT C H

NO T R ANS F E R

LIG HT

N.C .

MOT IV E F LOW VALVE F LOAT S WIT C H

A UXIL IA R Y TA NK


F UE L P R E S S UR E S WIT C H

5A P R E S S WAR N

S T B Y P UMP

S W

S T R A INE R

E NG INE -DR IV E N B OOS T P UMP

NO. 3 OR 4 DUAL-F E D B US

TO E NG INE

F R OM C R OS S F E E D

S WIT C H 10A

PR

E

S S T O T ES

T

TO WING TA NK

NA C E L L E TA NK

A UXIL IA R Y T R A NS F E R

CONT R OL A ND TI ME R (P.C .B )

AUX T R ANS

5A

5A 5A

HOT B AT B US

1

Fuel Systems



amndby Boosa Puep wtaSc

CCAUTIO If a standby pump fails, crossfeed can only be accomplished from the side of the operative pump.

Two separate sources of power operate the standby boost pumps. The lever-lock toggle switches on the fuel management panel, labeled STANDBY PUMP ON-OFF, provide power through the No.3 bus (LH pump) and the No.4 Dual-Fed bus (RH pump).On aircraft prior to S/N BB-1096, the Hot Battery bus also provides power to the standby pumps. For this reason, confirm that both standby boost pump switches are OFF during shutdown to prevent battery discharge.

Jha Trmnsfhr PuepA jet pump (venturi) transfers fuel from the sump of the auxiliary tank to the nacelle tank for each wing system. Fuel pressure from the engine-driven boost pump or the standby boost pump provides the motive flow by which the jet transfer pump operates.The pressured fuel line that supplies the motive flow to the jet transfer pump routes along the outboard side of the nacelle through the jet pump motive flow valve aft of the firewall to the jet transfer pump in the sump of the auxiliary tank.The lever-lock switches on the fuel management panel labeled AUX TRANSFER OVERRIDE-AUTO control the motive flow for the jet transfer pump.With the switch in AUTO, the automatic fuel transfer module under the center aisle floor applies power to the motive flow valve. If fuel is present in the auxiliary tank and no fuel transfer is taking place from the auxiliary tank to the nacelle tank (no PSI is sensed by the motive flow pressure switch), the NO TRANSFER light illuminates.If the motive flow valve fails to open due to an auxiliary tank float switch failed in the EMPTY position, and/ or a malfunction in the auxiliary transfer control module, select AUX TRANSFER OVERRIDE to transfer fuel.With the switch in AUX TRANSFER OVERRIDE, the fuel transfer module is bypassed, and power is applied directly to the normally-closed motive flow valve. When the valve opens, auxiliary fuel should transfer.

Dhfuhltng AdmpahrA defueling adapter is an integral part of the nacelle tank. The adapter, aft of the standby pump, contains a check valve that prevents fuel drainage when the access cover and plug is removed. Attaching a fuel hose to the adapter allows fuel to flow from the aircraft to the defueling vehicle. (See Servicing and Procedures section, this chapter.)

Crossfeed Limitations

Crossfeeding of fuel is permitted only when one engine is inoperative .

Maximum allowable fuel imbalance between wing fuel systems is 1,000 lbs .

Engine-Driven Pump Limitation

Operation with the FUEL PRESSURE annunciator on is limited to 10 hours between main engine-driven fuel pump overhaul or replacement .



Ftrhwmll cuaoff VmlvhsGuarded, two-position switches on the fuel management panel control the motored firewall shutoff valves, which are installed immediately aft of the engine firewall.The OPEN position allows uninterrupted fuel flow to the engine. The CLOSED position cuts off all fuel to the engine. When the red guard closes, it forces the switch into the OPEN position and protects it in that position.The No.3 and No.4 Dual-Fed buses, respectively supply power to the left and right firewall shutoff valves. The Hot Battery bus also powers them. A 5A fuse (5A CBs on S/N 1096 and subsequent) provides protection at the Hot Battery bus, while 5A CBs on the lower fuel panel protect the Dual-fed buses. Total failure of both power sources results in the valves remaining in the last selected position.

Figure 5F-6: Firewall Shutoff Valves

Fuhl FtlahrFrom the firewall shutoff valve, fuel is routed to the fuel strainer filter. The 20-micron filter incorporates a bypass valve to permit fuel flow in case of plugging. A pressure switch mounted directly above the filter senses boost pump fuel pressure at the filter. At approximately 10 PSI (±1 PSI) of decreasing pressure, the switch closes and actuates the red FUEL PRESSURE annunciator.

Crossfhhd VmlvhA crossfeed valve is in the left inboard nacelle, aft of the firewall. With one engine inoperative, the crossfeed switch on the fuel management panel opens the crossfeed valve and turns on the standby pump on the supply side to furnish fuel to the operating engine from the inoperative side.

Fuel Systems



Crossfhhd Vmlvh wtaScA manually-operated switch on the fuel management panel, labeled CROSSFEED FLOW, controls the crossfeed system. When the crossfeed switch is actuated, the fuel management panel provides power to the solenoid to open the crossfeed valve. Additionally, the switch activates the standby boost pump on the side from which crossfeed is desired, regardless of the standby pump switch position, and closes the motive flow valve on the side being fed.The green FUEL CROSSFEED annunciator is illuminated. Illumination of the annunciator indicates switch position. It is not a positive indication that the crossfeed valve has moved or that the standby boost pump is operating. See Hot Battery Bus preflight check. A 5A CB on the fuel management panel from the No.4 Dual-Fed bus powers the crossfeed.

Moatvh Flow VmlvhThe motive flow valves are installed in their respective engine nacelles just aft of the firewall. The AUX TRANSFER switches electrically control the normally closed rotary actuator-type valves.

Fuhl DrmtnsThe main and auxiliary fuel systems have five sump drains, a drain manifold, and a firewall filter drain. These push-to-open drain valves are semi-flush and externally-mounted to allow draining of sediment, moisture, and/or fuel from the system.When draining the flush-mounted drains, do not turn the draining tool. Turning or twisting of the draining tool unseats the O-ring seal and causes a leak.Since jet fuel and water are of similar densities, water does not settle out of jet fuel as easily as it does from aviation gasoline. For this reason, prior to draining the sumps to maximize water removal, do not move the aircraft after refueling for approximately three hours. Although turbine engines are not as critical as reciprocating engines regarding water ingestion, water should be drained periodically to reduce fungus growth and contamination-induced inaccuracies in the fuel gauging system.

Figure 5F-7: Fuel Drains



Fuhl Drmtn CollhSaor ysaheThe fuel system on S/N BB-2 to BB-665 and BL-1 to BL-9 incorporates a fuel drain collector system.After engine shutdown, a small amount of fuel in the fuel nozzle manifold drains into a small collector tank in the aft engine compartment. A normally-open float switch senses fuel level in the collector tank and closes when the tank fills.The switch activates an electric pump that transfers the fuel back to the nacelle tank at the next activation of the fuel line heater.A check valve in the line prevents backflow of fuel during purging of the fuel system at engine start. The drain line branches into the fuel return line aft of the firewall. A vent line runs from the top of the collector tank downward through a flame arrestor to the fuel drain manifold on the underside of the nacelle.When the collector tank is emptied, the float switch turns the pump off. The automatic operation requires no input from the pilot.

E NG INE

C OL L E C TOR TA NK

F L OAT S WIT C H

F UE L DR A IN C OL L E C TOR P UMP

F L A ME A R R E S TOR

F UE L MA NIF OL D

TO P UR G E VA LVE

TO DR A IN MA NIF OL D

F IR E WA L L

Figure 5F-8: Fuel Drain Collector System

Fuel Systems



Fuhl Drmtn Purgh ysaheThe fuel purge system on S/N BB666 and subsequent and BL-10 and subsequent ensures that any residual fuel in the fuel manifolds is consumed during engine shutdown. During engine starting, fuel manifold pressure closes the fuel manifold poppet valve to allow P3 air to pressurize the purge tank. During engine operation, engine compressor discharge air (P3 air) routes through a filter and check valve to maintain pressurization of the small purge tank.Upon engine shutdown, fuel manifold pressure subsides to allow the engine fuel manifold poppett valve to open. The pressure differential between the purge tank and fuel manifold causes air to be discharged from the purge tank; this forces residual fuel out of the engine fuel manifold lines through the nozzles and into the combustion chamber. As the fuel burns, a momentary surge in (N1) gas generator RPM can be observed. The entire operation is automatic and requires no input from the pilot.

E NG INE

A IR

P 3

P 3

P 3

F L OW DIV IDE R P UR G E TA NK P 3

F ILT E R

C HE C K VA LVE E NG INE F IR E WA L L

Figure 5F-9: Fuel Purge System




Fuel Systems



Servicing and Procedures

RefuelingDetermine the amount of fuel required. Ensure the fuel supply unit is grounded and then grounded to the aircraft. Also, ensure the fuel nozzle is grounded to the aircraft. Remove the filler caps and add required fuel. Once fueling is complete, replace filler caps and remove ground wires.Refuel the aircraft through the filler ports on each wing and in the auxiliary tank. Approved grounding procedures for the aircraft and fuel truck must be followed during refueling.

DrainingOpen each fuel drain daily to drain any water or other contamination collected in the bottom of the tanks. There are five sump drains, a manifold drain, and a firewall fuel filter drain for each side. The drain for the firewall fuel filter is found on the underside of the nacelle; the fuel pump and tank drains are accessible from the underside of each wing.

Defueling

CAOTON During defueling operations, place the battery and generator swithes in the OFF position; disconnect external power.

The aircraft can be defueled with the aid of a fuel truck. Ensure that there is no smoking within 50 ft of the aircraft while conducting defueling operations.Remove aircraft filler caps and properly ground all equipment used in defueling. Remove the cover on the bottom of the nacelle to gain access to the adapter plug. Attach the hose from the defueling truck to an AN 832-12 union.Remove the plug from the defueling adapter aft of the standby boost pump and screw the union into the adapter. Fuel can then flow from the aircraft to the defueling truck. Attach the clamp to the hose to prevent fuel leakage around the hose-union-adapter connection. Start the defueling pump.When defueling is complete, turn the defueling pump off. Disconnect the hose and the union from the defueling adapter. Re-install the adapter plug, remove the union from the hose, and re-install the fuel filler caps. Remove the grounding cables to complete the procedure.



Additive Procedures

CAOTON If fuel contacts the eyes, rinse with cool, fresh water and seek medical attention immediately. Avoid allowing fuel to contact the skin; when contact cannot be avoided, wash and mild soap and water.

CCAUTIO Anti-ice additive is toxic. It is dangerous when breathed and/or absorbed into the skin. When in contact with anti-ice additive, utilize appropriate protective equipment (eye goggles/shield, respirator with organic vapor cartridges, non-absorbing gloves) . If anti-ice enters the eyes, flush with water and contact a physician immediately.

Insert fuel nozzle with fuel additive nozzle attached into fuel filler. Ensure additive is directed into the flowing fuel stream and additive begins to flow after fuel. Stop the additive before the fuel flow.Additive concentration range should be maintained in accordance with instructions in the AFM.Biobor JF is approved for use as a biocide additive when premixed with fuel in the fuel supply facility. Use a metering device to inject Biobor JF. If no device available, blend the additive either through batch blending or over-the-wing blending when filling the tanks. Biobor JF may be used as the only fuel additive, or it may be used with an anti-icing additive conforming to MIL-I27686. Lack of anti-icing additive may cause fuel filter icing and subsequent engine flameout. Military JP-4 type fuels refined in the United States have anti-icing additive conforming to MIL-I-27686 blended at the refinery; no additional treatment is necessary. Before refueling, always check with fuel supplier to determine if the fuel contains anti-ice additive meeting MIL-I-27686.Prolonged aircraft storage may result in water buildup in the fuel that “leaches out” the anti-icing additive. Excessive water accumulation in the fuel tank sumps is an indication of this.

Preflight

CCAUTIO Do not allow concentrated additive to contact coated interior of fuel tank or aircraft painted surfaces .

During preflight, the following components of the fuel system are checked (see Preflight chapter for details): fuel gauges for any servicing requirements excessive fuel leakage (some seepage is expected) fuel sample at each sump drain

Fuel Systems



fuel caps closed and locked fuel tank vents unobstructed Hot Battery bus firewall shutoff valves crossfeed valve standby boost pumps.

Abnormal ProcedureThe following provides a brief discussion of what happens to the fuel system during abnormal conditions. For a list of specific procedural steps, refer to your CAE Operating Handbook.

FUEL CROFEED AnnunStmaor IlluetnmahsIllumination of the FUEL CROSSFEED annunciator is an indication that the CROSSFEED switch is in the crossfeed position. Monitor fuel gauges to verify that fuel is crossfeeding; when crossfeed is discontinued, verify that crossfeeding has stopped.

EehrghnSy ProShdurhsThe following provides a brief discussion of what happens to the fuel system during emergency conditions.For a list of specific procedural steps, refer to your CAE Operating Handbook.

Crossfhhd (Onh Engtnh Inophrmatvh Ophrmaton)Turn the standby boost pumps to OFF and position the crossfeed switch either LEFT or RIGHT (as required). Check that the FUEL CROSSFEED annunciator is ON. The auxiliary transfer switch must be in AUTO on the side receiving fuel. With a firewall shutoff valve closed, the auxiliary fuel supply is not available, and the FUEL PRESS light remains illuminated on the side supplying the fuel.To discontinue crossfeeding, move the crossfeed switch to the center position (OFF).

Engine-Driven Fuel Pump Limitations

Operation with the FUEL PRESS annunciator on is limited to 10 hours between main engine-driven fuel pump overhaul or replace-ment .



FUEL PRE AnnunStmaor Illuetnmahs

CCAUTIO AUX TRANSFER switch must be in AUTO position on the side being cross-fed. If the firewall valve is closed on the side of the inoperative engine, the auxiliary fuel supply is not available (usable) and the FUEL PRESS annunciator remains illuminated on the side supplying fuel.

If the FUEL PRESS annunciator illuminates, check for evidence of a fuel leak. If fuel is leaking, move the condition lever to CUT OFF. Feather the propeller and close the fuel fire-wall valve to isolate fuel from the engine. Finally, actuate the fire extinguisher if required.If there is no evidence of a fuel leak, it is likely that an engine-driven boost pump has failed. Turn on the standby pump on the failed side; verify that fuel pressure restores by checking the FUEL PRESS annunciator extinguishes. If the light remains illuminated, turn the pump off (to remove power from the fuel tank pump). Begin recording time.If engine operation is abnormal, perform Emergency Engine Shutdown checklist in the Operating Handbook.

Fuel Systems



Data Summary

Fuel System

Mmtn Fuhl ysahePower Source Hot Battery bus (BB-1097, 1095 and prior) or

No. 3 Dual-Fed bus Left standby pump Hot Battery bus (BB-1097, 1095 and prior) orNo. 4 Dual-Fed bus Right standby pump No. 4 Dual-Fed bus Crossfeed valve Hot Battery bus and/or No. 3 and No. 4 Dual-Fed buses Firewall shutoff valves (L/R)

Distribution Wing tanks (gravity feed) to nacelle tank Nacelle tank to engine

Control Switches STANDBY PUMP CROSSFEED FIREWALL SHUTOFF VALVES

Monitor Main fuel gauges Fuel flow indicator Annunciators FUEL CROSSFEED FUEL PRESS CROSSFEED (closes motive flow valve on receiving side,

opens crossfeed valve, turns on standby boost pump on feeding side, and illuminates crossfeed annunciator)

Protection Circuit breakers Check valves Fuses Fuel drain system Fuel filters (pressure switches) Vent systems Oil/Fuel heat exchanger

Auxtltmry Fuhl ysahePower Source Motive Flow Distribution Auxiliary (center) tank (automatic transfer to nacelle tankwith

AUX TRANSFER switch in AUTO) Control Switches

AUX TRANSFER OVERRIDE-AUTO (opens motive flow valve)

Monitor Aux fuel gauges NO TRANSFER lights

Protection Circuit breakers Fuses




Ice and Rain Protection 5G

King Air 200 5G-1December 2011


ContentsIce and Rain Protection

SchematSc: Deice System .......................................................... 5G-5Surface Protection

Surface Deice ......................................................................................... 5G-7Wing and Stabilizer Deice Boots .......................................................... 5G-7

Wing Ice Lights .................................................................................. 5G-9Powerplant Protection

Engine Anti-Ice .......................................................................................5G-11Engine Air Inlet Lip Heat ....................................................................5G-11

SchematSc: Propeller Deice System ...........................................5G-12Engine Inertial Separators .................................................................5G-13Engine Auto-Ignition System .............................................................5G-16

Propeller Electric Deice System ...........................................................5G-16Dual Heating Element Deicer Boots ....................................................5G-17Single Heating Element Deicer Boots .................................................5G-18

SchematSc: Windshield Anti-Ice System .....................................5G-19Additional Anti-Ice Systems

Fuel System Anti-Ice .............................................................................5G-21Pitot Heat Anti-Ice ..................................................................................5G-22Stall Warning Vane ................................................................................5G-23Windshield Protection ...........................................................................5G-23

Windshield Wipers .............................................................................5G-24Windshield Anti-Ice ............................................................................5G-25

Brake Deice System ..............................................................................5G-27Servicing and Procedures

Servicing ................................................................................................5G-29Preflight ..............................................................................................5G-29

Abnormal Procedures ...........................................................................5G-29ICE VANE Annunciator Illuminates ....................................................5G-29Reduction in Airspeed at Altitude .......................................................5G-30Windshield Electrical Fault .................................................................5G-30

Data SummariesIce and Rain Systems ............................................................................5G-31

King Air 2005G-2December 2011


Surface Deice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-31Prop Heat System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-31

Anti-Ice Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-31Brake Deice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-31Pitot Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-32Stall Warning Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-32Fuel Vent Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-32Windshield Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-32Ice Vanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5G-32

Ice and Rain Protection



Ice and Rain ProtectionA variety of ice and rain protection systems permit operation in inclement weather conditions. The systems use electrical power, bleed air and exhaust to protect the aircraft surfaces, power-plant, and additional systems.The surface system protects the wing and stabilizer, and the powerplant systems protect the engine, its components, and the propellers. Additional systems protect the fuel, pitot/static, and stall warning systems, the windshield, and optionally, the brakes.





5G-5For Training Purposes Only


Deice System

DEICE DISTRIBUTOR

VALVE

TIMER



RUDDER BOOST

OFF

OFF

MANUAL

SINGLE CYCLETO TAIL BOOTS

TO WING BOOTS

18 PSI PRESSURE

REG

ENVIR OFF

INSTR & ENVIR OFF


LEFT RIGHT

RIGHT P INSTRUMENT AIR

3

N. O. INSTRUMENT AIR VALVE

RIGHT BRAKE DEICE VALVE

N. O.

INSTRUMENT AIR VALVE

LEFT BRAKE DEICE VALVE

LEFT P INSTRUMENT AIR

3

LEFT BRAKE DEICE MANIFOLD

RIGHT BRAKE DEICE MANIFOLD

15 PSI PRESSURE

REG

TIMER

R BL AIR FAIL

23

25 30

35

2119

17

15

3

1

2 4

5

6 7 8 9

10 11

13

0

CABIN

ALT

1000 FT

RATE CABIN ALT

ACFT ALT 100 FT

15

M

N

M A X

PNEUMATIC PRESSURE

GAUGE

COPILOT'S ATTITUDE

INDICATOR

COPILOT'S TURN & SLIP INDICATOR

SUCTION GAUGE

PRESSURIZATION CONTROLLER

VACUUM AIR FILTER

CABIN AIR

FILTER

4 PSI REG

N. O.

TO DOOR SEAL

N.O. PRESET SOLENOID

N.C. DUMP SOLENOID

TO DUMP VALVE

FLT HOUR METER

N. C.

N. C.

N. C.

N. C.

L BL AIR FAIL

FILTER

UP

DNGEAR

UPLOCK

P SWITCH LEFT RUDDER BOOST SERVO

RIGHT RUDDER BOOST SERVO

UNREGULATED

REGULATED

VACUUM

RUDDER




5A

5A

5A

5A

5A

5A

5A




VACUUM REGULATOR

20 20

1010

L R

AIR

FLIGHT

0 4 5 6 2 HOURS 1/10

GYRO SUCTION

INCHES OF MERCURY

3 4 6

PNEUMATIC PRESSURE

0

10

20

TO OUT FLOW

VALVE

BRAKE DEICE ON

ON

OFF

BRAKE DEICE


5G-6 For Training Purposes Only





Surface Protection

Surface DeiceThe surface deice system removes ice accumulations from the leading edges of the wings and horizontal stabilizers. Alternately inflating and deflating the pneumatic deice boots accomplishes ice removal.

Wing and Stabilizer Deice BootsPressure-regulated bleed air from the engines supplies pressure to inflate the wing and stabilizer boots. A venturi ejector, operated by bleed air, creates a vacuum that deflates the boots and holds them down while not in use.

Figure 5G-1: Wing and Stabilizer Boots

A check valve incorporated in the bleed air line from each engine assures operation of the system if one engine fails, and prevents loss of pressure through the compressor of the inoperative engine. A distributor valve controls the inflation and deflation phases of the boots.



A three-position switch on the pilot’s subpanel, placarded DEICE CYCLE SINGLE-OFF-MANUAL, controls the deicing operation. The switch is spring-loaded, and returns to the OFF position from SINGLE or MANUAL. Selecting the SINGLE position opens the distributor valve and inflates the wing boots. After an inflation period of approximately six seconds, a timer switches the distributor to deflate the wing boots, and a four-second inflation begins in the horizontal stabilizer boots. When the stabilizer boots have inflated and deflated, the cycle is complete.Holding the switch in the MANUAL position inflates all the boots simultaneously; they remain inflated until the switch is released. The switch returns to the OFF position when released. After the cycle, the boots remain in the vacuum hold-down condition until actuated again by the switch. Power for the system is from the No.1 Dual-Fed bus.Allow at least 1/2 inch of ice to form before attempting ice removal to achieve the most effective deicing operation. Very thin ice can crack and cling to the boots instead of shedding. Subsequent cyclings of the boots then have a tendency to build up a shell of ice outside the contour of the leading edge, making ice removal efforts ineffective.

Figure 5G-2: DEICE CYCLE SINGLE-OFF-MANUAL, Controls




Wtng ISh LtgcasWing ice lights illuminate the wing leading edges for determination of ice buildup in icing conditions. The wing lights are on the outboard side of each nacelle. The control switch for the lights is on the pilot’s right subpanel in the LIGHTS group. Power for the lights is from the No.2 Dual-Fed bus.

Figure 5G-3: Wing Ice Lights Figure 5G-4: Control Switch







Powerplant Protection

Engine Anti-IceAn inertial separation system provides engine ice protection via electrical actuation of the ice vanes. Engine exhaust air continuously anti-ices the leading edge lip of the engine air inlet. Electrothermal boots on each propeller blade automatically cycle to prevent the formation of ice on the propellers.

Engtnh Atr Inlha Ltp HhmaHot exhaust gases heat the lip around each air inlet to prevent the formation of ice during inclement weather. A scupper in each engine exhaust stack deflects the hot exhaust gases downward into the hollow lip (overboard at the 6 o’clock position) tube that encircles the engine air inlet. No shutoff or temperature indicator is necessary for this system; heat flows through the inlet whenever the engine is running.

EXHAUST PIPE

EXHAUST SEALPLATE

INLET LIP ANTI-ICE

CLAMP

FLEXIBLE HOSE AIR INTAKE INLET ANTI-ICE LIP

EXHAUST STACK

WITHOUT KIT NO. 101-9048 WITH KIT NO. 101-9048

Figure 5G-5: Engine Air Inlet Lip Heat

On S/Ns BB-2 through BB-1265, BL-1 through BL-28, BN-1 through BN-4 and BT-1 through BT-31, the standard engine exhaust is depicted on the left side of Figure 5G-5.On S/Ns BB-2 through BB-1265, BL-1 through BL-28, BN-1 through BN-4 and BT-1 through BT-31 with Kit No. 101-9048 installed and including S/Ns BB-1266 and subsequent, BL-129 and subsequent, BT-32 and subsequent, and BN-5 and subsequent, the engine exhaust routed around the air intake inlet lip is depicted on the right side.

NTEc: With Kit No. 101-9048 installed, engine exhaust is routed around the air intake inlet lip from one stack to the other and then exhausted overboard.



Propeller Deice System

RH PROP LH PROP

PROP AMMETER

SHUNT

INNER

OUTER OUTER MANUAL OVERRIDE RELAY

INNER MANUAL OVERRIDE RELAY

NO 1

D U A L

F E D

B U S

PROP TIMER (34 SEC)

AUTO PROP DEICE

LH MANUAL

PROP DEICE

MANUAL PROP DEICE

CONTROL

RH MANUAL PROP DEICE

5A

20A

20A

20A

NO 3

D U A L

F E D

B U S

NO 4

D U A L

F E D

B U S

MODEL 200 BB-2, BB-6 thru BB-815, BB-817 thru BB-824, BL-1 thru BL29

BB-816, BB825 and subsequent, BL-30 and subsequent

PROP TIMER (90 SEC)

AUTO PROP DEICE

MANUAL PROPMANUAL

PROP DEICE CONTROL SWITCH

MANUAL PROP DEICE

CONTROL

LH MANUAL

PROP DEICE

RIGHT MANUAL OVERRIDE RELAY

LEFT MANUAL OVERRIDE RELAY

RH PROP LH PROP

PROP AMMETER

SHUNT

NO 4

D U A L

F E D

B U S

NO 1

D U A L

F E D

B U S

NO 3

D U A L

F E D

B U S

5A

20A

20A

20A




Engtnh Inhratml hpmrmaors

AAUIIO Once the manual override system is engaged (i.e., anytime the manual ice vane T-handle is pulled), do not attempt to retract or extend the ice vanes electrically, even if the T-handle is pushed back in, until the override linkage in the engine compartment is properly reset on the ground. The override linkage cannot be reset in flight. The manual system may be reused until the electric system has been reset.

An electrically actuated inertial vane system on each engine prevents ice or other foreign objects, such as dust or sand, from entering the engine inlet plenum and prevents ice from accumulating on the engine inlet screen.At temperatures above 15°C, close (RETRACT) the motor-driven ice vane and bypass door to assure adequate cooling ram airflow through the oil cooler.

Figure 5G-6: Motor-Driven Ice Vane (RETRACT)

OIL COOLER

BYPASS DOOR ACCESS PLATE

INERTIAL VANE

INDUCTION AIR

INLET LIP ANTI-ICE

OIL COOLER

INERTIAL VANE

INDUCTION AIR

INLET LIP ANTI-ICE

BYPASS DOOR

ACCESS PLATE

Figure 5G-7: Cooling Ram Airflow Through The Oil Cooler



When in icing conditions (5°C or colder and visible moisture) EXTEND the ice vane to deflect the airstream slightly downward. This creates a venturi effect and introduces a sudden turn of the air into the engine; the bypass door in the lower cowling at the aft end of the airduct opens with the vane.

Figure 5G-8: Motor-Driven Ice Vane (EXTEND)

As ice particles or water droplets enter the air inlet, they accelerate by the venturi effect of the extended vane. Due to their greater mass and greater momentum, the moisture particles or ice accelerate past the screen area and discharge overboard through the bypass door as the airstream makes the sudden turn before entering the engine through the inlet screen.

Ice Vane ControlsThe ice vane and bypass doors extend or retract simultaneously through a linkage system connected to electric actuators. Switches in the ICE VANE group on the pilot’s left subpanel energize the actuators. The ICE VANE switches EXTEND the separators in the switch up position and RETRACT them in the switch down position. Use the retract position for all normal flight operations at temperatures above 15°C.

Figure 5G-9: Ice Vane Controls Figure 5G-10: Ice Vane Controls




Extend the ice vanes whenever there is visible moisture at 5°C or colder. When the ice vanes are extended, the two green advisory annunciators illuminate. Observe a drop in torque and a slight increase in ITT because the airflow into the engine is restricted. When the ice vanes and bypass doors retract, the annunciators extinguish, torque is restored, and ITT decreases.A manual override mechanical backup system operates the system by flexible cable. If the vanes and doors do not move within 15 seconds after electrical actuation, an amber annunciator on the advisory panel provides information to the pilot. Pull the ICE VANE CBs on the copilot’s CB panel to disable electric power and pull the manual T-handle for the appropriate engine to activate the system. Once the manual system is engaged, do not attempt to retract or extend electrically until linkage is properly reset.When the vane is successfully positioned with the manual system, the yellow annunciator lights extinguish and green annunciators illuminate to confirm ice vane extension. Verify ice vane movement by observing movement of the bypass doors.The vane may also be retracted with the manual system. During manual system use, the electric motor switch position must match the manual handle position for a correct annunciator readout. Maximum airspeed for manual extension of the ice vanes is 160 Kts.On S/Ns BB-1439, BB-1444 and subsequent, BN-5 and subsequent, BT-35 and subsequent and BL-139 and subsequent, the inertial ice vane and bypass door are extended or retracted simultaneously through a linkage system connected to an electric dual-motor actuator. Two switches control both the left and right engine systems. The ACTUATOR switch is in the MAIN position except when the ACTUATOR STANDBY position is used to actuate the backup motor because the main motor is inoperable.

Figure 5G-11: T-Handle



Engtnh Auao-Igntaton ysaheThe engine auto-ignition system provides automatic ignition to prevent engine power loss due to combustion failure. Once armed, the system ensures ignition during takeoff, landing, turbulence, and penetration of icing or precipitation conditions. If ice or rain cause an engine flameout, or if engine torque falls below 400 ft-lbs once the system is armed, electrical power automatically energizes the engine igniter; the green IGNITION ON annunciator on the Caution/Advisory panel illuminates. During ground operation, turn the system off to prolong the life of the igniters.The switches that arm the auto-ignition system are on the pilot’s left sub-panel, above the ice vane switches and left of the control column. Moving the switches into the up or ARM position activates the system. Lift each switch over a detent to move it into or out of the ARM position. This lever-lock feature prevents inadvertent movement to the OFF position. Power for the left engine auto ignition system is from the No.3 Dual-Fed bus and from the No. 4 Dual-Fed bus for the right engine.

Figure 5G-12: Engine Auto-Ignition Switches

Propeller Electric Deice SystemAn electrically heated boot for each propeller blade provides automatic and manually controlled anti-ice protection for the propellers; the system includes the following additional components: brush assemblies slip rings an ammeter a timer select relays on-off control switches.




When the ON-OFF switch is ON, the ammeter registers the amount of current (l4 to l8A for three-bladed propellers; 18 to 24A for optional four-bladed, single-boot propellers) passing through the system. If the current rises beyond the switch limitations, a CB shuts off power to the deicer timer. Current flows from the timer through the brush assemblies to the slip rings, where it is distributed to each of the individual propeller deicer boot elements.Heat produced by the elements in the deicer boots reduces ice adhesion; the centrifugal effect of the propeller and the blast of the airstream removes the ice from the propeller.

Figure 5G-13: Ammeter

Dual Heating Element Deicer BootsPower to the deicer boot heating elements cycles in a continuous programmed sequence. S/Ns BB-2 to BB-815, BB-817 to BB-824 BB-991, and BL-1 to BL-29 have dual heating element deicer boots. One element deices the inner portion of the propeller boot and the other element deices the outer portion of the deicer boot.The deicer timer cycles power to these heating elements in the following sequence: RH outboard, RH inboard, LH outboard, and LH inboard. Each sequence has a duration of approxi mately 30 seconds; note a flicker of the deice ammeter as the timer switch es elements. Power for the automatic systems is from the No. 1 Dual-Fed bus, through a 20A CB switch labeled PROP AUTO-OFF on the pilot’s right hand sub panel.A manual propeller deicer system provides a backup to the automatic system. A control switch on the pilot’s right subpanel controls the manual relays. The switch is labeled PROP INNER-OUTER, and receives power through the No. 4 Dual-Fed bus. It is spring-loaded to the center (OFF) position. When the switch is in the OUTER position, the outer heating elements of both propellers receive power.



When the switch is moved to the INNER position, the inner heating elements of both propellers receive power. Hold the switch in the desired position approximately 45 seconds, or until the ice dissipates. Power for the outer elements is from the No. 4 Dual-Fed bus and power for the inner elements is from the No. 3 Dual-Fed bus, each through a 20A CB. The propeller ammeter does not indicate any load in the manual mode of operation.

Single Heating Element Deicer BootsS/Ns BB-816, BB-825 to BB-990, BB-992 and subsequent, and BL-30 and subsequent incorporate improved single heating element deicer boots. The boots are controlled by a 20A CB-type control switch on the pilot’s right subpanel labeled PROP AUTO-OFF, and receives power from the No.1 Dual-Fed bus.Power to these deicer boots cycles in 90-second phases. The first 90-second phase heats all the deicer boots on the RH propeller. The second phase heats all the deicer boots on the LH propeller. The deicer timer completes one full cycle every three minutes. Note a momentary deflection of the propeller ammeter needle as the deicer timer moves from one phase to the next.The manual override switch is labeled PROP MAN-OFF. When the switch is in the MAN position, the entire deice surface of both props receive power. The manual override switch is momentary; hold it in place until the ice has been dislodged from the propeller surface (approximately 45 seconds). The load meters in the overhead panel indicate an increase of load when the manual propeller deicer system is in operation. The propeller ammeter does not indicate any load in the manual mode of operation.

Figure 5G-14: Manual Override Switch

NTEc: The heating sequences for the deicer boots that follow are normal operating sequences. Because the timer does not return to any given point when power is removed, the sequence may restart at any point.




Windshield Anti-Ice System

WINDSHIELD HEAT SWITCH

NORMAL

OFF

HIGH

50A

5A

NORMAL

OFF

HIGH

50A

5A

WINDSHIELD HEAT SWITCH

265 IN2 AT 4.5 WATTS/IN2 HIGH SWITCH POSITION

360 IN2 AT 2.4 WATTS/IN2 NORMAL SWITCH POSITION

G E N

B U S

G E N

B U S

TEMP CONTROLLER (100-105° F)

TEMP CONTROLLER (100-105° F)







Additional Anti-Ice Systems

Fuel System Anti-IceTwo anti-ice systems protect fuel flow through the fuel lines to the engine; one system protects the fuel vents and the other protects the fuel itself.Two types of fuel vents on the underside of each wing outboard of the nacelle prevent ice blockage of the fuel vent system. The primary fuel vent is a NACA vent, recessed into the wing and designed so ice cannot obstruct it. The second vent, which is heated to prevent icing, serves as a backup to the primary vent. Left and right fuel vent heat switches in the ICE PROTECTION group on the pilot’s right subpanel operate the fuel vent heat. Power for the left vent heat is from the No. 1 Dual-Fed bus, and from the No. 2 Dual-Fed bus for the right vent heat. Turn these switches ON for all flights.

Figure 5G-15: Fuel Vents Figure 5G-16: Fuel Vent Heat Switches



An oil-to-fuel heat exchanger on the engine’s accessory section provides ice protection to the fuel control unit. An engine oil line within the heat exchanger is next to the fuel line; heat transfer to the fuel occurs through conduction, and melts any ice particles in the fuel. This operation is automatic whenever the engines are running. Refer to the Limitations section for temperature limitations.

FUELOUT

OILIN

OILOUT

FUELIN

THERMAL

HEAT EXCHANGER CORE

SLIDE VALVE

Figure 5G-17: Fuel System Anti-Ice

Pitot Heat Anti-IceTwo pitot tubes on the nose of the aircraft contain heating elements that protect against ice accumulation. The pitot tubes are electrically heated to ensure that proper airspeed is indicated during icing conditions. Two PITOT switches on the pilot’s right subpanel control left and right pitot heat. Power for the left pitot heat is from the No. 1 Dual-Fed bus, and from the No. 2 Dual-Fed bus for the right pitot heat.

Figure 5G-18: Pitot Tube Figure 5G-19: PITOT Switches




Separate pitot-static systems are available for each pilot’s instruments. An alternate source of static air, aft of the rear pressure bulkhead, is for the pilot’s instruments only. The static ports in the aft fuselage are not heated, since they are not susceptible to icing.

NTEc: Limit operation of pitot heat on the ground to two minutes to prevent damage to the pitot heat elements.

Stall Warning Vane

AROIOG Stall speed increases when ice accumulates on any airplane.

The stall warning vane is heated to ensure against freeze-up during icing conditions. A two-position ON-OFF switch to the right of the surface deicer cycle switch (pilot’s right subpanel) activates the stall warning anti-icer. Power for the stall warning vane heat is from the No. 2 Dual-Fed bus. Turn the stall warning vane heat on for all flights.A safety switch on the left landing gear limits the current flow to approximately 12V to prevent overheating while the airplane is on the ground. In flight, after the left strut extends, the stall warning vane and mounting plate receive 28V current. The heating elements protect the lift transducer vane and face plate from ice. A buildup of ice on the wing may change or disrupt the airflow and prevent the system from accurately indicating an imminent stall.

Figure 5G-20: Stall Warning Vane

Windshield ProtectionElectric heating elements in the windshield provide protection against the formation of ice, while air from the cabin heating system prevents fogging. Heavy duty windshield wipers provide improved visibility during rainy flight conditions.



Wtndscthld WtphrsSeparate windshield wipers are on the pilot’s and copilot’s windshield. The dual wipers are driven by a single electric motor, installed forward of the instrument panel.

Figure 5G-21: Windshield Wipers

The windshield wiper control is on the overhead light control panel. It provides the wiper mechanism with SLOW, FAST and PARK positions. The wipers may be used either on the ground or in flight, as required; however, they must not be operated on a dry windshield. The windshield wiper CB is on the copilot’s right side CB panel in the WEATHER group.

Figure 5G-22: Windshield Wiper Control




Wtndscthld Anat-ISh

AAUIIO In the event of windshield icing during sustained icing conditions, it may be necessary to reduce the airspeed to 226 Kts or below in order to keep the windshield ice-free.

The pilot’s and copilot’s windshields each have independent controls and heating circuits. The control switch allows the pilot to select a HI or a NORMAL intensity heat level (Windshield Anti-Ice System, page 5G-19).The windshields are composed of three physical layers. The inner layer is a thick panel of glass that is the structural member. The middle layer is a polyvinyl sheet that carries fine wire heating grids. The outer layer is a protective layer of glass bonded to the first two layers. The outside of the windshield is treated with a static discharge film, called a NESA coating.Electrical heating elements protect the windshields against icing. The heating elements connect at terminal blocks in the corner of the glass to wiring leading to the control switches mounted in the pilot’s right subpanel.The laminations of each windshield incorporate a transparent material (usually stannic oxide) that has high electrical resistance to improve the deicing capability of the windshield. Each windshield also has electrical connections for the resistive material and for temperature sensing elements. The resistive material is arranged to provide primary and secondary heated surfaces.Switches in the ICE PROTECTION group on the pilot’s right subpanel, labeled WSHLD ANTI-ICE NORMAL-OFF-HI (PILOT and COPILOT) control windshield heat. Selecting the PILOT and COPILOT switches to the NORMAL (up) position, heats the inboard and outboard areas of the windshields. Selecting the switches to the HI (down) position, heats only the outboard areas. The smaller areas heat faster, but to the same temperatures as in the NORMAL position. A lever-lock feature prevents inadvertent selection of the HI position when moving the switches from NORMAL to the OFF (center) position.

Figure 5G-23: Windshield Anti-Ice System



A temperature sensing element embedded in each windshield and a temperature controller in each windshield circuit automatically control windshield temperature. The temperature sensors signal the controller to remove power from the windshield heat elements when the window temperature reaches between 32.2°C (90°F) and 37.7°C (100°F).When the NORMAL level of heating is selected, an automatic temperature controller senses the windshield temperature, and attempts to maintain it at approximately 32.2°C (90°F) to 37.7°C (100°F) by energizing the “low” heat relay as necessary. In this mode, a larger area of the windshield heats.When the HI level of heating is selected, the same temperature controller senses the windshield temperature and attempts to maintain it at 32.2°C (90°F) to 37.7°C (100°F). In this mode, however, the controller energizes the “high” heat relay switch, which applies the electrical heat to a more concentrated, but more essential viewing area of the windshield. In HI, approximately two thirds of the windshield heats at the outboard portion.The power circuit of each system routes through 50A current limiters in the power distribution panel under the center aisle floor. Windshield heater control circuits route through 5A CBs on a panel mounted on the forward pressure bulkhead (forward of the pilot’s left subpanel).

NTEc: Windshield heat may be used at any time and in any combination. Use of windshield heat, however, may cause erratic operation of the magnetic compass because of the electrical field created by the heating elements. Windshield heat can cause slight visual distortion.




Brake Deice System

AAUIIO Use of brake deice during engine-out procedures substantially reduces the effectiveness of rudder boost assistance. Turn brake deice off for takeoff.

Optional brake deicers may be installed on the 200 or B200 King Air. Bleed air from each engine compressor flows to a distributor manifold that directs the hot air to the brakes to prevent ice and slush from building up between the wheels and freezing the brakes.The brake deicer is plumbed into the bleed air system that provides air for surface deice and instrument vacuum operation. Bleed air flows through a line on the left side of each nacelle to a solenoid-operated shutoff valve on the left side of each main gear wheel well.

FROM ENGINE BLEED AIR

BRAKE DEICE DISTRIBUTOR MANIFOLD

Figure 5G-24: Brake Deice System

From the shutoff valve, bleed air flows through a hose secured to the aft side of the landing gear strut to a distributor manifold attached to the piston and axle assembly. The bleed air routes to the brake for each wheel through orifices around the circumference of each ring of the distributor manifold.An ON-OFF toggle switch on the pilot’s right subpanel controls the brake deice system. When the switch is ON, a control module under the center aisle floor receives power from the aircraft electrical system through a 5A CB in the copilot’s side panel. This module supplies current that opens the solenoid shutoff valves in each wheel well, and allows the hot bleed air to enter the distributor manifold for diffusion through the orifices.



The control module simultaneously provides a signal that illuminates the green BRAKE DEICE ON annunciator on the pedestal. Performing a takeoff without switching the system OFF completes a circuit through the uplock switch to a timing circuit in the control module when the main landing gears reach the up and locked position. The timing circuit turns the deice system off after 10 minutes of operation, closing the solenoid valve in the wheel well and stopping the flow of bleed air to the brakes. This is done to prevent overheat damage to adjacent components in the wheel well. Do not activate the system again until the landing gear completes one cycle from the up and locked position.





Servicing

PreflightDuring preflight inspection, check the wing inspection lights, static ports, and pitot tubes to ensure they are clear and to verify heat operation. Check the wing, stabilizer, and propeller deice boots for security, cuts, cracks, or tears.The windshield wiper blades should be free of cracks, nicks, or breaks. Drain the static system and the fuel sumps.

Abnormal ProceduresThe following section provides a brief discussion of what happens to the ice and rain protection systems during abnormal conditions. For specific procedural steps, please refer to the CAE Operating Handbook.

ICE VAE AnnunStmaor Illuetnmahs(Prior to BB-1439 and 1444)

AAUIIO Once the manual override system is engaged (i.e., anytime the manual ice vane T-handle has been pulled out), do not attempt to retract or extend the ice vanes electrically, even if the T-handle is pushed back in, until the override linkage in the engine compartment is properly reset on the ground. The override linkage cannot be reset in flight.

If either vane does not attain the selected position in 15 seconds, a yellow caution annunciator illuminates. In this event, two T-handles below and between the pilot’s left and right sub-panels provide a mechanical backup. Pull the ICE VANE CBs to remove electrical source, then pull the T-handle to override the electrical malfunction.To manually extend the ice vanes, first pull the affected ice vane CB on the right CB panel. Decrease the airspeed to 160 Kts or less, then pull the affected handle fully aft. Advance power to resume cruise airspeed, but not to exceed ITT limits.

BB-1439, 1445 and SubsequentIf failure is confirmed, place engine Anti-Ice Actuator in STANDBY. If green anti-ice annunciator illuminates, resume airspeed. If green anti-ice annunciator fails to illuminate, exit icing conditions.



RhduSaton tn Atrsphhd ma AlataudhIf during flight at altitude, the pilot observes a gradual reduction in airspeed indication, suspect pitot icing. Turn the pitot heat ON to restore airspeed; if airspeed recovers, leave the pitot heat ON, as icing conditions exist. It is standard practice to keep the pitot heat on during all flights. Do not operate the pitot heat system on the ground, except for testing or for short intervals to remove snow or ice from the tube (maximum, two minutes). Turn on pitot heat before takeoff and leave it on in flight during icing conditions, or whenever icing conditions are expected.Perform the following Anti-ice/Deice check of each heated item: Turn all heated items on. Turn one generator off. Watch loadmeter decrease as each heated item is turned off. Once tests are completed, turn generator back on.

Wtndscthld ElhSartSml FmulaObservation of smoke and/or fire at inboard corner of either windshield may indicate overheat condition in the electrical power terminal for normal heat mode of windshield heat. If this occurs, turn WINDSHIELD ANTI-ICE switches – OFF. If smoke and/or fire persists, conduct ELECTRICAL SMOKE OR

FIRE Checklist. If smoke and/or fire ceases, continue flight with Windshield Anti-Ice

switches OFF. If windshield anti-ice required, isolate if possible and turn unaffected

side – NORMAL/HI.




Data Summaries

Ice and Rain Systems

urfmSh DhtSh ysahePower Source Bleed Air

No. 1 Dual-Fed bus Distribution Wing leading edge boots

Horizontal stabilizer leading edge boots Control DEICE CYCLE switch

SINGLE – inflation/deflation of wing boots then horizontal stabilizer boots MANUAL – inflation of all boots simultaneously

Monitor Visual monitoring for wing Pneumatic gauges Protection Circuit breakers

Prop Hhma ysahePower Source No. 1 Dual-Fed bus (auto)

No. 3 and 4 Dual-Fed bus (manual) Distribution Heated boot for each propeller blade Control Switches

PROP AUTO PROP MANUAL INNER/OUTER (200 only)

Monitor Prop ammeter Loadmeters

Protection Circuit breakers/Circuit breaker switch (AUTO)

Anti-Ice Systems

Brmkh DhtSh ysaheSource Engine P3 bleed air Control BRAKE DEICE switch Monitor BRAKE DEICE ON annunciatorProtection 10-minute timer



Anti-Ice System (continued)

Ptaoa HhmaPower Source Dual-Fed buses Nos. 1/2 Control PITOT circuit breaker switches Protection Circuit breaker switch

amll Wmrntng HhmaPower Source Dual-Fed bus No. 2 Control STALL WARN circuit breaker switches

Landing gear safety switch Protection Circuit breaker switch

Fuhl Vhna HhmaPower Source Dual-Fed buses Nos.1/2 Control FUEL VENT circuit breaker switches

Wtndscthld HhmaPower Source L/R GEN bus Control WSHLD ANTI-ICE switches Protection Circuit breaker (5A)

Temperature sensing element Temperature controller 50A current limiters

ISh VmnhsPower Source Dual-Fed buses Nos. 1/2 Control ICE VANE switches

VANE MANUAL PULL handle Monitor ICE VANE amber and green annunciators Protection Circuit breakers

Manual override system

Landing Gear Systems 5H

King Air 200 5H-1December 2011


ContentsLanding Gear Systems

SchematSc: Electro-Mechanical Landing Gear System ...............5H-4Mechanical Landing Gear

Main Gear ................................................................................................5H-5Main Gear Strut/ Torque Knees ..........................................................5H-6Main Gear Wheels/Tires .....................................................................5H-6Main Gear Doors ................................................................................5H-6Motor and Gearbox .............................................................................5H-7

Nose Gear ................................................................................................5H-7Nose Gear Wheel/Tire ........................................................................5H-8Duplex Chains/Uplock/Downlock ........................................................5H-8Nose Wheel Steering ..........................................................................5H-8

Landing Gear Operation ........................................................................5H-9Retraction/Extension ...........................................................................5H-9Emergency Gear Operation ...............................................................5H-11

SchematSc: Hydraulic Landing Gear ...........................................5H-13SchematSc: Hydraulic Landing Gear System ..............................5H-15tiguhc: Hydraulically Operated Landing Gear ............................5H-16

Hydraulic Landing GearMain Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-17

Main Gear Strut/ Torque Knees . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-17Main Gear Wheels/Tires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-17Main Gear Doors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-18Main Gear Hydraulic Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-18

Nose Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-18Nose Gear Wheel/Tire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-19Nose Gear Uplock/Downlock . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-19Nose Wheel Steering . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-19

Hydraulic Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-20Fill Reservoir . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-20Primary Reservoir . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-20Landing Gear Powerpack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-21HYD FLUID LOW Annunciator . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5H-21

King Air 2005H-2December 2011


System Filters ................................................................................... 5H-22Thermal Expansion Valve ................................................................. 5H-22Hydraulic Accumulator ...................................................................... 5H-22Service Valve .................................................................................... 5H-22

Landing Gear Operation ...................................................................... 5H-23Retraction/Extension ......................................................................... 5H-23Emergency Gear Operation .............................................................. 5H-25

SchematSc: Brake System ......................................................... 5H-27Brake System

Brakes .................................................................................................... 5H-29Wheel Brakes ........................................................................................ 5H-29Shuttle Valve ......................................................................................... 5H-29Brake System Operation ...................................................................... 5H-29Parking Brake ....................................................................................... 5H-30Brake Deicer System ............................................................................ 5H-30

Servicing and ProceduresPreflight ................................................................................................. 5H-31Servicing ............................................................................................... 5H-31

Tire Inflation ...................................................................................... 5H-31Struts ................................................................................................ 5H-32Hydraulic Service Valve .................................................................... 5H-32

Abnormal/ Emergency Procedures ..................................................... 5H-33Landing Gear Manual Extension (Mechanical Gear) ........................ 5H-33Landing Gear Manual Extension (Hydraulic Gear) ........................... 5H-33

Data SummariesLanding Gear Systems ......................................................................... 5H-35




Landing Gear SystemsThe landing gear system on the Beechcraft King Air is a conventional tricycle configuration with air/oil shock strut-type nose and main gear. The gear is operated either electrically by a 28V motor and gearbox, or hydraulically by an electrically-driven hydraulic powerpack.The main gear consists of a dual wheel and brake assembly, and the nose gear utilizes a single wheel. High flotation main gear tires and wheels are available as an option.Rudder pedals operate the multiple disc brakes on the main gear wheels. The system is mechanically controlled and hydraulically actuated. A dual-valve parking brake provides main wheel braking when the aircraft is parked. There is no emergency brake system.



Electro-Mechanical Landing Gear System

N O.

2

D U LANDINGA GEAR DOWNL RELAY LIMIT DOWN F DOWN SWITCH E D 5A

EMERGENCY LIMITEXTENSION GROUND

U B

SWITCH S UP AIRBORNE

LANDINGENGAGED GROUND GEARGEAR CONTROLLERDOWNLOCK AIRBORNE HOOK

GEAR SAFETY SWITCH LIMIT

UP LIMITR

I SWITCH G UP H T

60A G TORQUE E TUBESN 1 B U S

RIGHT MAIN TORQUE GEAR TUBES

SPRING-LOAD IDLER SPROCKETS

CROSS SHAFT

DUPLEX CHAIN

DOWNLOCK INDICATOR SWITCH

DOWNLOCK HOOK

1 LEFT MAIN GEAR

NOSE GEAR DRAG BRACE

ASSEMBLY




Mechanical Landing Gear

Main GearThe mechanical main gear assembly, which is on all S/Ns to BB1192, consists of dual wheels and brakes attached to a conventional air/oil shock strut. Other components include: torque knees gear doors electro-mechanical retract system – motor, gearbox, torque tubes,

mechanical actuators position indicators multiple disc brakes (see Brakes section, this chapter).

Two types of main gear are available:standard gear and high-flotation gear, which provide larger main gear wheels and a shorter main gear shock strut. On the high-flotation gear, the tires will protrude through landing gear cutouts when the landing gear is fully retracted.

ACTUATOR (ATTACHED TO STRUCTURE)

DRAG BRACE

SHOCK STRUT

BRAKE ASSEMBLY

TORQUE KNEE

PIVOT POINT

Figure 5H-1: Main Gear Assembly



Mmtn Ghmu auga/ Touqgh KnhhsEach strut has an isolation piston with hydraulic fluid on one side and a nitrogen pressure charge on the other; the piston floats within a cylinder for shock absorption during taxi, takeoff, and landing.Torque knees connect the strut cylinder to the piston and axle assembly; they hold the wheel in alignment and limit the maximum extension of the strut.

Mmtn Ghmu Wchhls/TtuhsEach main landing gear wheel consists of two wheel halves bolted together. Each wheel on standard-equipped aircraft is equipped with a 18 x 5.5 inch, 10- or 8-ply, tubeless rim-inflation tire. On aircraft equipped with high flotation landing gear, a 22 x 6.75 inch, 10 or 8-ply, tubeless rim-inflation tire is used for both the main and nose gear wheels.

Mmtn Ghmu DoousTwo doors enclose each main gear after retraction. The doors are mechanically actuated by gear movement during extension and retraction.Rollers on the shock strut contact cams in the wheel well during retraction. Linkages transmit the cam movement to close the doors. During extension, roller action opens the doors. When the rollers leave the cams, springs drive the linkage over-center to hold the doors open.The main gear doors on aircraft equipped with high-flotation landing gear incorporate cutouts in the doors that allow the oversized wheels to protrude out of the gear bay. These doors are mechanically linked to the shock strut and open and close as the gear extends or retracts.

Figure 5H-2: Main Gear Doors




Moaou mnd GhmuboxA 28 VDC split-field motor and gearbox, on the forward side of the center-section main spar, extends and retracts the landing gear. The motor incorporates a dynamic braking system controlled by UP and DOWN limit switches. In conjunction with the landing gear locking mechanism, the dynamic braking system prevents overtravel of the landing gear.A spring-loaded friction overload clutch in the gearbox prevents damage to the aircraft structure and torque tubes in the event of malfunction. On S/N 002 to 185, a 200A, remote CB on the landing gear panel under the center floor protects the system from electrical overload. On S/N 186 and subsequent, a 150A current limiter replaces the CB; Service Bulletin No. 2035 replaces the current limiter with a 60A CB.

Main Gear Torque TubesThe landing gear motor drives the torque tubes. The torque tubes transfer the energy of the motor to main gear actuators that drive the main gear up or down. Internal friction of the jackscrew in each actuator holds the main gear in the retracted position.

Main Gear DownlocksNotched hook and plate attachments fitted to each main gear drag brace provide positive mechanical downlocks.

Nose GearThe nose gear assembly consists of a single wheel assembly attached to a conventional air/oil shock strut. Other components include: torque knees gear doors motor and gearbox (see description under Main Gear).

ACTUATOR (ATTACHED TO STRUCTURE)

SHOCK STRUT

SHIMMY DAMPER

TORQUE KNEE

ROLLER (NOSE WHEEL DOOR)

PIVOT POINT

DRAG BRACE

Figure 5H-3: Nose Gear Assembly



The nose gear incorporates an air/oil shock strut that operates in the same manner as the main gear strut. Two doors enclose the nose gear after retraction. The doors are mechanically actuated in the same manner described for standard main gear doors.The nose gear assembly adds a shimmy damper that bleeds hydraulic fluid through an orifice to dampen nosewheel shimmy.

Figure 5H-4: Nose Gear after Retraction

Nosh Ghmu Wchhl/TtuhThe nose landing gear wheel consists of two wheel halves bolted together, similar to those in the main gear. Each nose gear wheel on standard-equipped aircraft is equipped with a 22 x 6.75 inch, 8- or 10-ply, tubeless rim-inflation tire.

Dgplhx Ccmtns/UploSk/ DownloSkThe landing gear motor drives duplex chains. These chains transfer the energy of the motor to the nose gear actuators, which in turn drive the nose gear up or down. Internal friction in the jackscrew in each actuator holds the gear in the retracted position, just as in the main gear. The over-center action of the drag brace on the nose gear assembly provides positive mechanical downlock.

Nosh Wchhl ahhutniDirect linkage of the nose gear to the rudder pedals permits nose wheel steering when the nose gear is down. A spring-loaded link in the system absorbs some of the force applied to any of the interconnected rudder pedals until the nose wheel is rolling. Once the nose wheel is rolling, the resisting force reduces; more pedal motion results in more nose wheel deflection. The nose gear automatically centers for retraction.




Since motion of the pedals is transmitted via cables and linkage to the rudder, rudder deflection occurs whenever force is applied to any of the rudder pedals. With the nose gear retracted, some of the force applied to the rudder pedals is absorbed by the spring-loaded link in the steering system. Normal rudder pedal steering of the nose wheel allows up to 14° left and right of center. Augmenting normal deflection with differential power and/or main wheel braking allows up to 48° left and right of center.

Landing Gear OperationA manually-actuated, wheel-shaped switch controls landing gear system operation. The switch, labeled LDG GEAR CONTROL UP and DN, is on the pilot inboard panel. The LANDING GEAR CONTROL CB alongside the landing gear control provides protection for the landing gear switch and associated relay circuits.A squat switch on the right main gear torque knee opens the control circuit when the strut is compressed (aircraft on the ground). The squat switch also actuates a solenoid-operated downlock hook on the landing gear control switch. This mechanism prevents the landing gear handle from being raised when the aircraft is on the ground. The hook automatically unlocks when the aircraft leaves the ground. If the down-lock solenoid fails, the down-lock releases by pressing the red DOWNLOCK REL button alongside the landing gear control handle.

Figure 5H-5: Landing Gear Operation (Switch Controls)

RhaumSaton/ExahnstonA 28 VDC split-field, 11/2 HP motor drives the mechanical landing gear via a gearbox, torque tubes, and duplex chains. The landing gear motor has two series of windings; one field drives the motor to retract the gear and one field drives the motor in the opposite direction to extend the gear.



To prevent overtravel of the gear, a dynamic brake relay simultaneously breaks the power circuit to the motor and makes a complete circuit through the armature and the unused field winding. The motor then acts as a generator and the resultant electrical load on the armature stops the gear.Torque tubes from the gearbox carry torque to drive the main gear actuators. Duplex chains that attach to a sprocket on the gearbox torque tube drive the nose gear actuators. A spring-loaded friction clutch between the gearbox and the torque tube protects the motor in the event of mechanical malfunction.

Position Indication

CAUTIO Stop pumping the handle when the three green GEAR DOWN lights illuminate. Further movement of the handle could damage the drive mechanism and prevent subsequent electrical gear retraction.

An assembly of three individual GEAR DOWN lights on the landing gear control panel provide position information to the crew. One light in each segment, when illuminated, makes the segment appear green and indicates the particular gear is DN and locked. Absence of illumination indicates gear UP. The green position indicator lights test by pressing.A landing gear warning horn provides a warning to the pilot that the landing gear is not in the down and locked position during specific flight conditions. Various warning modes result, depending on the position of the flaps.A landing gear warning horn silence button is next to the landing gear switch handle or is on the left power lever. Two red parallel-wired indicator lights in the landing gear control handle indicate that the gear is in transit or unlocked.The red lights extinguish in a GEAR UP condition. The red lights also illuminate when the landing gear warning horn is actuated. These lights test by pressing the HDL LT TEST button next to the handle itself.The indicator lights receive information from the normally-closed, up position switches, one of which is in the upper portion of each wheel well. When the gear is in the fully retracted position, each strut actuates its respective switch and opens the circuit from the intransit light to ground. As the gear moves from the fully retracted position, the switches close and illuminate the respective intransit light. The intransit light extinguishes when the drag brace in each landing gear passes over-center and actuates its respective downlock switch.Illumination of the landing gear intransit lights indicates one or more of the following conditions: landing gear handle is in the UP position and the aircraft is on the ground

with weight on the gear one or both power levers are retarded below a preset power level and

at least one landing gear is not down and locked any one or more landing gear is not fully retracted or in the down and

locked position




warning horn has been silenced and will not operate; the light remains illuminated when the horn is silenced.

one or more of the landing gear is not down and locked and the flaps are selected past approach. In this condition, the warning horn can only be silenced by retracting the flaps or extending the landing gear.

EehuihnSy Ghmu OphumatonA separate, manually powered chain-drive system provides manual (emergency) gear extension. To operate the gear manually, pull the LANDING GEAR RELAY CB on the pilot right panel. Verify that the landing gear handle is in the DN position; pull up on the EMERGENCY ENGAGE handle (on the floor) and turn clockwise about 60° to lock it into position.When the emergency engage handle is up, the motor electrically disconnects from the system; the emergency drive system is locked into the gearbox. Pumping the handle adjacent to the emergency engage handle ratchets the gearbox and lowers the landing gear. Approximately 70 to 90 pumping actions move the gear from full up to full down.After a practice manual extension, the landing gear may be retracted electrically. Rotate the EMERGENCY ENGAGE handle counterclockwise and push down. Stow the extension lever, push in the LANDING GEAR RELAY CB on the pilot panel, and retract the gear in the normal manner with the landing gear switch handle. Be sure to return clutch disengage handle to normal position (J-hook down).

RIGHT MAIN

NOSE GEAR

DOWN LIMIT SWITCH

UP LIMIT SWITCH

LEFT MAIN

MANUAL EXTENSION CABLE TENSIONER

MANUAL EXTENSION LEVER

LANDING GEAR

ALTERNATE EXTENSION

1. LIFT AND TURN HANDLE TO

ENGAGE.

CLUTCH DISENGAGE HANDLE

1

1

2. PUMP HANDLE UP AND DOWN UNITL THREE

GREEN GEAR LIGHT ILLUM.

Figure 5H-6: Emergency Gear Operation



Tcts pmih tnahnatonmlly lhfa blmnk.


5H-13Developed for Training Purposes Only


Hydraulic Landing GearNormal Extension

FILL RESERVOIR

RH MAIN DOWN LOCK SWITCH

NOSE DOWN LOCK SWITCH

LH MAIN DOWN LOCK SWITCH

CONTROL SWITCH

DOWN

UP

SERVICE VALVE SWITCH

RH SAFETY SWITCH

TIME DELAY

14 SEC

5A 60A

4.87 ΩRESISTOR

RELAY

HAND PUMP

VENT PORT AUXILIARY RETURN PORT (PLUGGED) POWER PACK ASSEMBLY

PRIMARY RESERVOIR

SYSTEM RELIEF VALVE

PUMP CHECK VALVE

PUMP MOTOR

FILTER RELIEF VALVE

SOLENOID SOLENOID

DOWN UP

PRESSURE CHECK VALVE

HAND PUMP DUMP VALVE

FILTER THERMAL RELIEF VALVE

ACCUMULATOR

VENT VALVE

AUXILIARY PRESSURE PORT (PLUGGED)

FILTER

GEAR DOWN PORT

GEAR UP PORT

RH MAIN ACTUATOR

NOSE ACTUATOR

LH MAIN ACTUATOR

SERVICE VALVE

HAND PUMP PRESS PORT

PRESSURE SWITCH

HAND PUMP SUCTION PORT

FILL PORT

SECONDARY RESERVOIR

ORIFICE

OVERBOARD VENT

2A

RETURN FILTER

PUMP

PRESSURE

RETURN

STATIC

BLEED AIR

REGULATED ENGINE BLEED AIR (18 TO 20 PSI)

LH SAFETY SWITCH

RH GEN BUS

1 ACTUATOR (DETAIL)

DOWN

UP

1

N O.

2

D U A L

F E D

B U S


5H-14 Developed for Training Purposes Only




Developed for Training Purposes Only

Hydraulic Landing Gear SystemEmergency Extension

ORIFICE FILL RESERVOIR

HAND PUMP

VENT PORT AUXILIARY RETURN PORT (PLUGGED) POWER PACK ASSEMBLY

PRIMARY RESERVOIR

SYSTEM RELIEF VALVE

PUMP CHECK VALVE

PUMP MOTOR

FILTER RELIEF VALVE

SOLENOID SOLENOID

DOWN UP

PRESSURE CHECK VALVE

HAND PUMP DUMP VALVE

FILTER THERMAL RELIEF VALVE

ACCUMULATOR

VENT VALVE

AUXILIARY PRESSURE PORT (PLUGGED)

FILTER

GEAR DOWN PORT

GEAR UP PORT

RH MAIN ACTUATOR

NOSE ACTUATOR

LH MAIN ACTUATOR

SERVICE VALVE

HAND PUMP PRESS PORT

PRESSURE SWITCH

HAND PUMP SUCTION

PORT

FILL PORT

SECONDARY RESERVOIR

OVER-BOARD VENT

RETURN FILTER

PUMP

1

ACTUATOR (DETAIL)

DOWN

UP

1 PRESSURE

RETURN

STATIC

BLEED AIR

SUCTION



Hydraulically Operated Landing Gear

RIGHT MAIN

DOWNLOCK INDICATOR SWITCH

DRAG BRACE ASSEMBLY

DOWNLOCK HOOK

1

POWER PACK SERVICE (RESERVOIR PUMP VALVE & MOTOR) LEFT

GEAR ACTUATOR

NOSE GEAR ACTUATOR

1

HAND PUMP

LEFT MAIN GEAR ACTUATOR

PLUMBING NETWORK TO POWER PACK

FULL CAN LEFT WING

WING

ACCUMULATOR LEFT




Hydraulic Landing Gear

Main GearThe hydraulic main gear assembly, which is on S/N BB 1193 and subsequent, consists of dual wheels and brakes attached to a conventional air/oil shock strut. Other components include: torque knees gear doors hydraulic retract system – hydraulic actuators, electrically-driven

power-pack position indicators multiple disc brakes (see Brakes section, this chapter).

Two types of main gear are available: standard gear and high-flotation gear, which provide larger main gear wheels and a shorter main gear shock strut. The high-flotation gear does not retract completely into the wheel well and protrudes through cutouts in the gear doors approximately five inches into the airstream.

Mmtn Ghmu auga/Touqgh KnhhsEach strut has an isolation piston with hydraulic fluid on one side and an air pressure charge on the other; the piston floats within a cylinder for shock absorption during taxi, takeoff, and landing.Torque knees connect the strut cylinder to the piston and axle assembly; they hold the wheel in alignment and limit the maximum extension of the strut.

Mmtn Ghmu Wchhls/TtuhsEach main landing gear wheel consists of two wheel halves bolted together. Each wheel on standard-equipped aircraft is equipped with a 18 x 5.5 inch, 10- or 8-ply, tubeless rim-inflation tire. On aircraft equipped with high flotation landing gear, a 22 x 6.75 inch, 10 or 8-ply, tubeless rim-inflation tire is used for both the main and nose gear wheels.



Mmtn Ghmu DoousTwo doors enclose each main gear after retraction. The doors are mechanically actuated by gear movement during extension and retraction.The main gear doors on aircraft equipped with high-flotation landing gear incorporate cutouts in the doors that allow the oversized wheels to protrude out of the gear bay. These doors are mechanically linked to the shock strut and open and close as the gear extends or retracts.

Figure 5H-7: Main Gear Doors

Mmtn Ghmu HydumgltS ASagmaouA hydraulic powerpack in the left wing root pumps hydraulic pressure to the main gear actuators. The main gear actuators have two hydraulic lines that provide fluid for extension (normal extend and hand pump extend) and a single line for retraction. Hydraulic pressure holds the main gear in the retracted position.

Main Gear DownlocksNotched hook and plate attachments fitted to each main gear drag brace provide positive mechanical downlocks.

Nose GearThe hydraulic nose gear assembly is similar to that of the main gear, but consists of a single wheel assembly attached to a conventional air/oil shock strut. Other components include: torque knees gear doors hydraulic actuator.

Hydraulic LimitationUse only MIL-H-5606 red hydrau-lic fluid.




The nose gear incorporates an air/oil shock strut that operates in the same manner as the main gear strut. Two doors enclose the nose gear after retraction. The doors are mechanically actuated in the same manner described for standard main gear doors. Hydraulic pressure actuates the nose gear in the same manner as the main gear.A shimmy damper that bleeds hydraulic fluid through an orifice to dampen nosewheel shimmy is added to the nose gear.

Figure 5H-8: Nose Gear

Nosh Ghmu Wchhl/TtuhThe nose landing gear wheel consists of two wheel halves bolted together, similar to those in the main gear. Each nose gear wheel on standard-equipped aircraft is equipped with a 22 x 6.75 inch, 8- or 10-ply, tubeless rim-inflation tire.

Nosh Ghmu UploSk/DownloSkThe nose gear is held in the retracted position with hydraulic pressure, and in the extended position by an internal lock in the hydraulic actuator.

Nosh Wchhl ahhutniDirect linkage of the nose gear to the rudder pedals permits nose wheel steering when the nose gear is down. A spring-loaded link in the system absorbs some of the force applied to any of the interconnected rudder pedals until the nose wheel is rolling. Once the nose wheel is rolling, the resisting force reduces; more pedal motion results in more nose wheel deflection. The nose gear automatically centers for retraction.Since motion of the pedals is transmitted via cables and linkage to the rudder, rudder deflection occurs whenever force is applied to any of the rudder pedals. Normal rudder pedal steering of the nose wheel allows up to 14° left and right of center. Augmenting normal deflection with differential power and/or main wheel braking allows up to 48° left and right of center.



Hydraulic ComponentsOn S/N 1158, 1167, 1193 and subsequent, an electro-hydraulic powerpack system produces sufficient pressure to extend and retract the landing gear. The system operates on approximately seven to eight U.S. quarts of hydraulic fluid. The hydraulic system consists of the following components: a fill reservoir a primary reservoir a landing gear powerpack filters and check valves an accumulator an emergency hand pump.

tll RhshuvotuThe fill reservoir inboard of the left nacelle and forward of the front spar contains a cap and dipstick assembly to simplify system maintenance. Hydraulic fluid from the fill reservoir supplies the primary reservoir. The dip stick marked HOT/FILL, COLD/ FILL allows for convenient fluid level checking and servicing.Regulated bleed air enters the fill reservoir through a check valve at approximately 18 PSI. This pressure provides a positive fluid flow to the primary reservoir and powerpack to prevent pump cavitation and fluid foaming. A line plumbed to the upper portion of the fill reservoir vents excessive bleed air overboard. The check valve prevents return air flow, maintaining a head pressure on the reservoir.

Putemuy RhshuvotuThe three quart primary reservoir inboard of the left engine nacelle and forward of the main spar employs a standpipe to maintain reserve fluid for emergency use. The primary reservoir supplies fluid to the landing gear powerpack. Fluid returns to the reservoir during the cycling of the landing gear.

Powerpack LimitationThe powerpack requires a resting rate of one minute per cycle and five minutes after the fifth cycle to prevent damage to the electric motor.




Lmndtni Ghmu PowhupmSkThe landing gear powerpack consists of a combined electric motor and hydraulic pump. It draws fluid from the primary hydraulic reservoir and delivers this fluid under pressure through three separate lines to the individual landing gear actuators at 2,775 ±55 PSI.

TO NORMAL EXTEND SIDE OF SYSTEM

FROM EMERGENCY EXTEND SIDE OF SYSTEM

SYSTEM FILTER

GEAR SELECT SOLENOID

FLUID LEVEL SENSOR

TO NORMAL RETRACT SIDE OF SYSTEM

TO HAND PUMP

FROM HAND PUMP

TO TOP OF FILL CAN (VENT)

MOTOR

TO BOTTOM OF FILL CAN

RED KNOB

TO PLUMBING CONNECTED TO BOTTOM OF FILL CAN AND TO POWER PACK RESERVOIR

FILTER BODY

UPLOCK PRESSURE SWITCH

FROM EMERGENCY EXTEND SIDE OF SYSTEM

TO RETRACT SIDE OF SYSTEM

HAND PUMP SUCTION PORT

SERVICE VALVE

Figure 5H-9: Landing Gear Powerpack

Current passes through the landing gear downlock switches and a 2A CB from the No. 2 Dual-Fed bus to the motor relay. The relay closes, and provides a path for 28 VDC from the right GEN bus to the powerpack motor by passing through the 60A relay CB under the center aisle floor. A protection circuit interrupts power to the powerpack by opening the CB if operation continues for more than 14 seconds.The pump housing supports a pressure switch, low fluid level sensor, four-way solenoid selector valve, and the return fluid filter. A check valve downstream of the powerpack prevents fluid from entering the power-pack during emergency hand pump operation.The pressure switch in the hydraulic return fluid line interrupts power to the power relay if the switch senses 2,775 ±55 PSI. A system drop of 300 PSI closes the switch allowing electrical power to flow to the powerpack.

HYD LUID LOW AnngnStmaouThe amber HYD FLUID LOW annunciator illuminates anytime the fluid level sensor in the powerpack senses a low fluid condition. A sensing unit on the end of the motor of the powerpack provides the switching circuitry that illuminates the HYD FLUID LOW annunciator. The No.2 Dual-Fed bus supplies electrical power.



A four-second delay in the system minimizes false indications caused by sloshing of fluid. To test the annunciator, press the HYD FLUID SENSOR TEST button on the pilot’s panel.

Figure 5H-10: HYD FLUID LOW Annunciator

ysahe tlahusA hydraulic filter in the pressure line along with two filters in the return lines prevent foreign matter from circulating in the system and damaging components. Only one filter in the return line incorporates a bypass valve, which allows fluid to return to the reservoir if a filter becomes blocked.

Tchueml Expmnston VmlvhWith the landing gear down and locked, a thermal expansion valve set at 80 PSI in the retract side relieves pressure of hydraulic fluid expansion. The expanding fluid returns to the primary reservoir.

HydumgltS ASSgeglmaouThe cylindrical hydraulic accumulator absorbs and dampens sudden hydraulic surges and helps maintain system pressure. An air charging valve and direct-reading pressure gauge on the air side of the accumulator allow servicing of the unit. The accumulator contains an 800 ±50 PSI charge of dry nitrogen or air.

huvtSh VmlvhA service valve in the left wing root allows ground maintenance personnel to retract the gear manually. The valve is not accessible in flight.




Landing Gear OperationA manually-actuated, wheel-shaped switch controls landing gear system operation on aircraft with either mechanical or hydraulic gear. The switch, labeled LDG GEAR CONTROL UP and DN, is on the pilot inboard panel. The LANDING GEAR CONTROL CB alongside the landing gear control provides protection for the landing gear switch and associated relay circuits.A squat switch on the right main gear torque knee opens the control circuit when the strut is compressed (aircraft on the ground). The squat switch also actuates a solenoid-operated down-lock hook on the landing gear control switch. This mechanism prevents the landing gear handle from being raised when the aircraft is on the ground. The hook automatically unlocks when the aircraft leaves the ground. If the down-lock solenoid fails, the down-lock releases by pressing the red DOWNLOCK REL button alongside the landing gear control handle.

RhaumSaton/ExahnstonThe nose and main landing gear assemblies extend and retract by the action of the hydraulic powerpack in conjunction with the accumulator.Three separate lines from the power-pack to each of the gear actuators supply hydraulic pressure for each of the three landing gear modes (retract, extend, and emergency extend).The landing gear control handle actuates landing gear travel. Moving the handle to the DN position actuates the powerpack down solenoid; fluid flows to the extend side of the actuators. As the actuator piston moves to extend the landing gear, the fluid on the other side of actuators exits through the retract port and flows back to the power-pack through the retract plumbing. Fluid from the pump flows through the selector valve, opens a pressure check valve, and allows the return fluid flow into the primary reservoir.When the actuator pistons position to fully extend the gear, an internal mechanical lock in the nose gear actuator locks the actuator piston and holds the gear in the down position; the main gear are held in the extend position by the mechanical locking system. In this position, the downlock switches interrupt current to the power relay.Moving the landing gear control handle to the UP position provides hydraulic fluid under pressure to the retract side of the gear actuators. As the actuator pistons move to retract the landing gear, the fluid in the other side of the actuators exits through the extend port and flows back to the powerpack through the extend plumbing. The fluid flows to the powerpack through the selector valve and returns to the primary reservoir.When the landing gear reaches the fully retracted position, hydraulic system pressure holds the gear in the UP position. When hydraulic pressure reaches approximately 2,775 PSI, the uplock pressure switch opens the landing gear relay and interrupts current to the pump motor. The same pressure switch causes the pump to actuate and increase hydraulic pressure if it drops below 2,475 PSI.



Position IndicatorsAn assembly of three individual GEAR DOWN lights on the landing gear control panel provide position information to the crew. One light in each segment, when illuminated, makes the segment appear green and indicates the particular gear is DN and locked. Absence of illumination indicates gear UP. The green position indicator lights test by pressing.A landing gear warning horn provides a warning to the pilot that the landing gear is not in the down and locked position during specific flight conditions. Various warning modes result, depending on the position of the flaps.A landing gear warning horn silence button is next to the landing gear switch handle or is on the left power lever.

Figure 5H-11: Position Indicators

Two red parallel-wired indicator lights in the landing gear control handle indicate that the gear is in transit or unlocked. The red lights extinguish in a GEAR UP condition. The red lights also illuminate when the landing gear warning horn is actuated. These lights test by pressing the HDL LT TEST button next to the handle itself.The indicator lights receive information from the normally-closed, up position switches, one of which is in the upper portion of each wheel well.When the gear is in the fully retracted position, each strut actuates its respective switch and opens the circuit from the intransit light to ground. As the gear moves from the fully retracted position, the switches close and illuminate the respective intransit light. The intransit light extinguishes when the drag brace in each landing gear passes over-center and actuates its respective downlock switch.The landing gear intransit lights indicate one or more of the following conditions: landing gear handle is in the UP position and the aircraft is on the

ground with weight on the gear one or both power levers are retarded below a preset power level

and at least one landing gear is not down and locked any one or more landing gear is not fully retracted or in the down and

locked position




warning horn has been silenced and will not operate; the light remains illuminated when the horn is silenced.

one or more of the landing gear is not down and locked and the flaps are selected past approach. In this condition, the warning horn can only be silenced by retracting the flaps or extending the landing gear.

NOTEc: The landing gear may be extended with the hand pump only in flight.

EehuihnSy Ghmu OphumatonThe emergency hand pump supplies fluid to the extend side of the gear actuators in the event of a malfunction or absence of electrical power to the landing gear system. The hand pump, labeled LANDING GEAR ALTERNATE EXTENSION, is between the pilot and copilot seats. It produces hydraulic pressure through the manual actuation of the handle.During manual gear extension, the hand pump manually forces hydraulic fluid to the landing gear actuators.Pull the LANDING GEAR RELAY CB on the pilot inboard subpanel to interrupt electrical power to the power-pack. Select the DN position on the landing gear control handle and remove the hand pump handle from the securing clip. Pump the handle up and down to extend the gear until three green GEAR DOWN indicator lights illuminate and increased resistance is felt. Place the handle in the full down position and secure it in the retaining clip. After a practice manual extension, the landing gear may be retracted hydraulically. Push the LANDING GEAR RELAY CB in and move the LDG GEAR CONT switch handle to the UP position.

Figure 5H-12: Emergency Gear Operation





5H-27Developed for Training Purposes Only


Brake System

RESERVOIR DRAIN

BB-2 TO BB-452

PARKING BRAKE

PILOT'S MASTER CYLINDERS

SHUTTLE VALVE

COPILOT'S MASTER CYLINDERS

BB-453 TO BB-665 BB-666 AND SUBSEQUENT

PARKING BRAKE

PARKING BRAKE



SHUTTLE VALVE



RESERVOIR DRAIN RESERVOIR DRAIN

LEFT BRAKE

RIGHT BRAKE

LEFT BRAKE

RIGHT BRAKE

LEFT BRAKE

RIGHT BRAKE

SUPPLY

STATIC

BRAKE PRESSURE


5H-28 Developed for Training Purposes Only





Brake System

BrakesThe main landing gear wheels on the King Air 200/B200 incorporate multi-disc, non-assisted hydraulic brakes that use MIL-H-5606 hydraulic fluid. Master cylinders attached to both sets of rudder pedals actuate the brakes. The master cylinders receive fluid from a reservoir in the nose avionics compartment.No emergency braking system is provided; use reverse or beta range propeller settings to taxi and to slow the aircraft after landing.

Wheel BrakesThe wheel brake is a multiple disc design, and consists of two rotating disks and a stationary disc. There is one brake assembly for each main gear wheel. Each brake assembly also includes a torque plate and brake housing.Radial tangs that engage slots in the wheel drive each multiple disc brake assembly. The tangs rotate on either side of a stationary disc keyed to the torque tube. The stationary disc and torque plate provide a friction surface for the rotating discs. The torque plate is bolted to the brake housing.The brake housing contains brake pistons, return springs, and an inlet and bleeder port. Cavities interconnect each piston to provide simultaneous brake actuation with equalized pressure.

Shuttle ValveS/Ns prior to BB-666 incorporate a shuttle valve adjacent to each set of pedals. The shuttle valve permits changing the braking action from one set of pedals to the other. The dual brakes on aircraft BB-666 and subsequent are plumbed in series to allow either set of pedals to perform braking action.

Brake System OperationDepressing either set of pedals compresses the piston rod in the master cylinder attached to each pedal. The hydraulic pressure from the movement of the pistons in the master cylinders is transmitted through flexible hydraulic lines and fixed aluminum tubing to the disc brake assemblies on the main landing gear.This pressure forces the brake pistons to press against the discs of the brake assembly.



Parking BrakeDual parking brake valves retain pressure built up by normal brake application in the brake lines. Depress the pilot’s brake pedals to build up pressure in the brake lines. Close both valves simultaneously by pulling out the parking brake handle on the left subpanel.This closes the valves and retains the pressure previously pumped into the brake lines. Depress the brake pedals and push the parking brake control in to release the parking brake.The parking brake system does not compensate for thermal expansion/ contraction and should only be used for short periods of time until the wheels can be properly chocked.

Figure 5H-13: Parking Brake

Brake Deicer SystemAn optional brake deicer system is available for the King Air 200/B200 that routes bleed air heat tubing into a distributor manifold attached to each brake. A BRAKE DEICE switch on the pilot inboard panel activates system.The switch opens shutoff valves to permit hot bleed air to enter the distributor manifolds and deice the brakes. As the shutoff valves open, electrical signals illuminate the green BRAKE DEICE-ON annunciator on the caution/advisory panel.

NOTEc: If the automatic timer terminates brake deice operation after the last retraction of the gear, the gear must be extended again to reset the deice system and allow further operation of the system.

Brake Deice Limitations

If the BRAKE DEICE switch is not turned off manually, a 10-minute time delay safety switch shuts off the brake deice system after the gear retracts into the wheel well. If the safety switch fails, manually select the system off after 10 minutes.

Do not operate the brake deice system at ambient temperatures above 15°C.

Maintain 85% N1 or higher during periods of simultaneous brake deice and wing boot operation. If inadequate pneumatic pressure is developed for proper wing boot inflation, select brake deice system off.

Both sources of instrument bleed air must be in operation when using brake deice.

Select brake deice system off during single engine operation.

The rudder boost system may not operate when the brake deice system is in use.





PreflightDuring the exterior preflight inspection, accomplish the following checks of the landing gear and brake system (see Preflight Inspection chapter for details). Visually inspect the nose gear and doors assembly for fluid leaks,

door condition and security, tire wear, and tire pressure. Check the turn block holes for symmetry to ensure the nose gear has not been turned too far during towing.

Check the main gear and doors for general security, fluid leaks, strut extension, brake condition, and main gear tire pressure.

During cockpit preflight inspection, test the landing gear annunciators by pressing on the light. Test the gear handle lights by pressing and holding the HDL LT TEST button on the right side of the handle.Check that the landing gear control handle is down; check for three green GEAR DOWN lights.

Servicing

Tire InflationManufacturer-recommended tire inflation specifications are: nose wheel – 55 to 60 PSI main gear wheels – standard gear, 94 to 98 PSI; high-flotation gear,

60 to 64 PSI.



augasThe landing gear struts are filled with MIL-H-5606 hydraulic fluid and inflated with dry air or nitrogen. Nose gear strut static deflection limits (depending on takeoff weight) is 3.0 to 3.5 inches. Main strut extension is as follows: standard gear – tire inflation, 96 ±2 PSI; strut extension, 3.95 to 4.19

inches high flotation gear – tire inflation, 62 ±2 PSI; strut extension,

5.56 to 5.93 inches.Due to static friction, the landing gear struts must not be serviced with wheels on the ground; this could cause incorrect readings leading to low inflation and the strut bottoming out when the aircraft is moved.

Figure 5H-14: Struts

HydumgltS huvtSh VmlvhThe service valve forward of the powerpack assembly along with the emergency hand pump allow retraction of the landing gear for maintenance purposes. To operate the valve with the aircraft on jacks and an external power source connected, unlatch the hinged retainer and pull up on the red knob on top of the service valve. Pump the hand pump to raise the gear to the desired position. Upon completion of maintenance push the red knob down and use the hand pump to lower the gear.




Abnormal/ Emergency ProceduresThe following section provides a brief discussion of what happens to the landing gear system during abnormal conditions. For a list of specific procedural steps, please refer to your CAE Operating Handbook.

Lmndtni Ghmu Mmngml Exahnston (MhScmntSml Ghmu)

CAUTIO During manual extension of landing gear, stop pumping the handle when the three green GEAR DOWN lights illuminate. Further movement of the handle could damage the drive mechanism and prevent subsequent electrical gear retraction.

When manual gear extension is required, establish an airspeed of 130 Kts. Pull the LANDING GEAR RELAY CB on the pilot right panel. Verify the landing gear switch is in the DN position and pull up on the EMERGENCY ENGAGE handle (on the floor) and turn clockwise to lock it into position.Pump the handle up and down until the three green GEAR DOWN lights illuminate. If manual landing gear extension is unsuccessful, refer to the appropriate Gear Unsafe Landing checklist.If the three green GEAR DOWN lights do not illuminate, as in the case of electrical failure, continue pumping the handle (even though this may damage the drive mechanism), until sufficient resistance is felt to ensure that the gear is down and locked.After making an emergency landing gear extension, do not stow the pump handle, move any landing gear controls, or reset any switches or CBs until the aircraft is on jacks. The failure may have been in the gear-up circuit and the gear could retract on the ground. The gear cannot be retracted manually.

Lmndtni Ghmu Mmngml Exahnston (HydumgltS Ghmu)When manual gear extension is required, establish an airspeed of 130 Kts. Pull the LANDING GEAR RELAY CB on the pilot’s inboard subpanel to interrupt electrical power to the power-pack. Select the DN position on the landing gear control switch and remove the hand pump handle from the securing clamp. Pump the handle up and down to extend the gear until three green GEAR DOWN indicator lights illuminate and increased resistance is felt. Place the handle in the full down position and secure it in the retaining clip.If the three green GEAR DOWN lights do not illuminate, as in the case of electrical failure, continue pumping the handle until sufficient resistance is felt to ensure that the gear is down and locked.After making an emergency landing gear extension, do not move any landing gear controls, or reset any switches or CBs until the aircraft is on jacks. Since the failure may have been in the gear-up circuit, and the gear could retract on the ground.







Data Summaries

Landing Gear SystemsElectro-Mechanical Landing Gear System

Power Source No. 2 Dual-Fed bus Landing gear control relay Right Generator bus 28 VDC split-field 11/2 HP motor

Control LDG GEAR CONTROL handle EMERGENCY ENGAGE handle

Monitor Gear handle light Gear warning horn Gear DOWN position lights

Protection Landing gear relay (5A) Circuit breaker (60A) Right main gear squat switch Emergency engage handle Limit switches Dynamic brake relay Solenoid-operated down-lock hook (landing gear handle)

Hydraulic Landing Gear SystemPower Source Right Generator bus

No. 2 Dual-Fed bus Landing gear control power relay Electric Motor Driven Hydraulic Pump (Power Pack)

Distribution Landing gear Control LDG GEAR CONTROL handle

Pressure switch Down-lock switches (3) Time delay module

Monitor HYD FLUID LOW annunciator Accumulator precharge direct reading gauge

Protection Circuit breakers LANDING GEAR RELAY (5A) Landing gear powerpack (60A) Pressure switch Thermal relief valve Down-lock switches (3) Internal nose gear mechanical lock Squat switches (L/R) Low fluid level sensor Time delay module Solenoid-operated down-lock hook (landing gear handle)



Landing Gear Systems (continued)Brake SystemPower Source Hydraulic pressure Distribution Master cylinders

Parking brake valves Control Brake pedals

PARKING BRAKE handle Emergency braking: reverse propeller for taxiing or slowing (-3° blade angle zero thrust) top of red and white strips on throttle quadrant Shuttle valves S/N prior to BB-666: valve adjacent to each set of pedals permit changing braking action from one to the

otherS/N BB-666 and subsequent: dual brakes plumbed in

series to allow either set of pedals to perform

Miscellaneous Systems 5I

King Air 200 5I-1December 2011


ContentsMiscellaneous Systems

SchematSc: Oxygen System (King Air 200) .................................. 5I-5SchematSc: Oxygen System (King Air B200) ................................ 5I-5

Oxygen SystemOxygen Bottle .......................................................................................... 5I-7

Shutoff Valve and Regulator ................................................................ 5I-8Crew Oxygen System ..............................................................................5I-11First Aid Oxygen ......................................................................................5I-12Passenger Oxygen ..................................................................................5I-12Oxygen Duration ......................................................................................5I-15Preflight ....................................................................................................5I-16Servicing ..................................................................................................5I-16

Emergency EquipmentEmergency Locator Transmitter ...........................................................5I-17Safety Equipment ....................................................................................5I-18Fire Extinguishers ...................................................................................5I-18Life Preservers .........................................................................................5I-18

Cargo DoorDoor ..........................................................................................................5I-19Operation ..................................................................................................5I-19

Warning SystemsWarnings ..................................................................................................5I-21Landing Gear Warning ............................................................................5I-21

Gear-in-Transit .....................................................................................5I-22Warning Switches ................................................................................5I-22Operation .............................................................................................5I-22

Stall Warning System ..............................................................................5I-24Annunciator System ................................................................................5I-25

Warning Annunciator Panel ................................................................5I-25Caution/Advisory Annunciator Panel ...................................................5I-26

tiguhc: Annunciator Panels ..........................................................5I-27Annunciator Cross Reference

King Air 2005I-2December 2011



Miscellaneous Systems



Miscellaneous SystemsThe Miscellaneous Chapter includes items not normally addressed in other chapters or items that span multiple chapters. This includes the oxygen system, emergency equipment, cargo door, and the warning and caution/advisory annunciator panels.The oxygen system provides supplementary oxygen for the crew and passengers.Emergency equipment includes safety equipment and over-water equipment including the emergency locator transmitter, portable fire extinguishers, first aid kit, and crash axe.The warning and caution/advisory annunciator panels provide visual warnings to the crew of items requiring immediate attention, items of concern, and indications of system operation.





5I-5Developed for Training Purposes Only


Oxygen SystemKing Air 200

PULL ON

OXYGEN OUTLETS

OUTLETS

CONTROL CABLE

OXYGEN CYLINDER

RELIEF

AND SHUT OFF VALVE

PASSENGER BAROMETRIC MANUAL OVERRIDE SHUTOFF VALVE

SWITCH PRESSURE

FORWARD PRESSURE BULKHEAD

ANNUNCIATOR "PASS OXYGEN"

SYSTEM READY

FIRST AID MASK STOWED IN MANUALY OPERATED BOX


HIGH PRESSURE OVERBOARD

PRESSURE REGULATOR FILL GAUGE

OXYGEN PRESSURE SENSE SWITCH

PASS OXYGEN ON

FILL VALVE

COCKPIT OXYGEN GAUGE

PASSENGER MANUAL OVERRIDE

HIGH PRESSURE

OXYGEN CONTROL5A

1

2

NO. 1 DUAL-FED BUS

IN

OUT

1

OXYGENRECHARGE

2

70 PSI

Oxygen SystemKing Air B200

IN

OUT

PASSENGER MANUAL OVERRIDE SHUTOFF VALVE

BAROMETRIC PRESSURE SWITCH

FORWARD PRESSURE BULKHEAD

OUTLETS


HIGH PRESSURE

LOW PRESSURE

HIGH PRESSURE OVERBOARD RELIEF

OXYGEN CYLINDER

PRESSURE REGULATOR FILL GAUGE

FILL VALVE

OXYGEN MASK

PULL ON SYSTEM READY

COCKPIT OXYGEN GAUGE

PASSENGER MANUAL OVERRIDE

OXYGEN OUTLET

CONTROL CABLE

OXYGEN PRESSURE SENSE SWITCH

PASS OXYGEN ON

FIRST AID MASK STOWED IN MANUALLY OPERATED BOX

OXYGENRECHARGE

NO. 1 DUAL-FED BUS

OXYGEN CONTROL5A

1

1

2

2

AND SHUTOFF VALVE


5I-6 Developed for Training Purposes Only





Oxygen SystemThe oxygen system supplies oxygen to two quick-donning masks in the cockpit, drop-down masks in the passenger cabin, and a single therapeutic (first aid) oxygen mask in the entry way. The crew oxygen system is diluter demand and the passenger oxygen system is constant-flow.The oxygen system consists of: single high-pressure oxygen bottle shutoff valve and regulator fill valve and gauge cockpit pressure gauge crew oxygen system first aid oxygen mask passenger oxygen system.

Oxygen BottleA single oxygen bottle is behind the aft pressure bulkhead. The bottle stores 100% aviator’s breathing oxygen at a nominal pressure of 1,800 to 1,850 PSI at 70°F. Depending on the aircraft serial number, owner’s preferences, and modifications, bottle capacity varies from 22 cubic-ft to 115 cubic-ft.

SUPPLY LINE

SHUTOFF VALVE

PRESSURE REGULATOR

OXYGEN CYLINDER

TO COCKPIT PRESSURE GAUGE

TO FILL VALVE

TO PRESSURE GAUGE

OVERBOARD VENT


PULL ON SYSTEM READY

Figure 5I-1: Oxygen Bottle



cgaoff Vmlvh mnd RhiglmaouThe regulator reduces unregulated high pressure oxygen from the bottle to a regulated pressure of approximately 70 PSI. The regulator then provides oxygen to the crew and passenger oxygen systems at a maximum flow rate of approximately 40 Liters Per Minute (LPM). An integral or separate shutoff valve controls the supply of oxygen from the regulator. Low and high pressure ports on the regulator connect to direct reading pressure gauges, a fill valve, overboard vent line, and the crew and passenger oxygen system supply line.If the oxygen bottle overpressurizes due to excessive temperature or overfilling, a disc in the regulator ruptures to release bottle contents through a vent line to the atmosphere.There is a fill valve and direct reading pressure gauge behind an access door on the right side of the aft fuselage near the static ports. A second direct reading pressure gauge is on the copilot subpanel.

Figure 5I-2: Fill Valve And Pressure Gauge Figure 5I-3: Pressure Gauge




The mechanically-operated shutoff valve connects through a cable to the oxygen supply control knob (PULL ON SYS READY) aft of the overhead lights control panel. Pulling the handle out opens the oxygen bottle shutoff valve; pushing it in closes the valve. With the handle out, oxygen flows from the bottle through the primary oxygen supply line to the first aid oxygen mask outlet, the crew mask outlets and the normally closed passenger oxygen system shutoff valve. With the shutoff valve open, oxygen is always available to the crew oxygen masks. The first aid oxygen mask outlet has a shutoff valve that must be opened before oxygen flows.

Figure 5I-4: Oxygen Supply Control Knob



Figure 5I-5: Masks

Figure 5I-8: Headband and by Holding

the Paddles

Figure 5I-11: Mask Microphone

Selector Switch(es)

Figure 5I-6: Pocket Type Overhead Storage Figure 5I-7: Two Red Paddles

Figure 5I-9: Headband once the Mask is in place

Figure 5I-10: Jack

Figure 5I-12: Three-Position(EMER/100%/NORM) Lever




Crew Oxygen System

WARRIRN Do not smoke while oxygen is in use.

The crew oxygen system consists of two quick-donning, diluter-demand oxygen masks. The masks deliver oxygen only on inhalation; there is no loss of oxygen with the system on and the masks plugged-in. The masks are located on the aft partition behind the crew seats and are readily available. The masks connect to individual outlets on the cockpit side just behind the side windows. For aircraft S/N BB-1439, BB-1444 and subsequent, BN-5 and subsequent, BT35 and subsequent, and BL-139 and subsequent, the oxygen ARM and Passenger Manual Override T-handles are located on the left and right side of the center console. The masks have been moved to a pocket type overhead storage. When oxygen is required, the crewmember reaches overhead and presses the two red paddles on the portion of the mask that extends from the pocket holder. This inflates the headband and by holding the paddles, the crew-member dons the mask. Releasing the paddles deflates the headband once the mask is in place.Each mask also has an integral microphone that connects with the audio system through a jack near the mask outlet. The microphone selector switch(es) must be in OXYGEN MASK to enable the mask microphone(s).A three-position (EMER/100%/NORM) lever on the mask selects the mode of operation. NORM dilutes oxygen from the mask with ambient air. As cabin altitude increases, the mask’s regulator increases the ratio of oxygen to ambient air. The use of diluted oxygen reduces the oxygen consumption rate and is more comfortable to breathe than pure oxygen.The 100% position supplies pure, undiluted oxygen to the mask. During all phases of flight select 100% to provide 100% oxygen in the event of excessive cabin altitudes and/or contaminated air. As a safety precaution, always store the mask with the lever in 100%. The EMER position supplies 100% oxygen under positive pressure.



First Aid OxygenWith the oxygen supply on, oxygen flows to an outlet and mask in the lavatory. The mask is behind an access door, (marked FIRST AID OXYGEN – PULL) on the entry way ceiling. The door can be manually opened at any time. Before using the mask turn the supply valve on.

Figure 5I-13: Mask Marked as FIRST AID OXYGEN – PULL

Passenger OxygenThe passenger oxygen system consists of: passenger oxygen system shutoff valve passenger oxygen masks and containers manual override system system annunciator barometric pressure switch.

The passenger oxygen system is a constant flow type system that can be manually or automatically deployed.On S/Ns BB-2 through BB-54 without BEECHCRAFT Kits No.1015006 and 101-5007 installed and BB38 through BB-54 with the Optional Autodeployment Oxygen System installed, an overhead push-pull control turns on the passenger oxygen system. This provides oxygen pressure to the cockpit outlets and the cabin outlet manifolds. To use the oxygen, the oxygen masks are removed from their place of storage under the seat and plugged into the outlet.




Stated Cylinder Size (cu ft)

Number of People Using Oxygen1

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 161 171

Duration in Minutes22 144 72 48 36 28 24 20 18 16 14 13 12 11 10 * * * 49 or 50 317 158 105 79 63 52 45 39 35 31 28 26 24 22 21 19 18 66 422 211 140 105 84 70 60 52 47 42 38 35 32 30 28 26 24 76 or 77 488 244 162 122 97 81 69 61 54 48 44 40 37 34 32 30 28 115 732 366 244 183 146 122 104 91 81 73 66 61 56 52 48 45 43

Table 5I-1: Oxygen Duration

*Will not meet oxygen requirements.1 For oxygen duration computations, count each diluter-demand crew mask in use as two (e.g., with four passengers and two crew members, enter the table at eight people using oxygen.)

On S/Ns BB-55 and subsequent, BT-1 and subsequent, BL-1 and subsequent, BN-1 and subsequent and BB-2 through BB-54 with BEECHCRAFT Kits No.101-5006 and 1015007 installed and BB-38 through BB-54 with the Optional Auto-deployment Oxygen System installed, the autodeployment passenger oxygen system is operated by two push-pull control cables and a barometric pressure switch.On S/Ns BB-1439, BB-1444 and subsequent, BT-35 and subsequent, BL139 and subsequent and BN-5 and subsequent, the push-pull control cables are located on the sides of the control pedestal.The left control cable operates the oxygen system shutoff valve and places the system in the ready mode when the knob is pulled; the right cable is the passenger manual-override control to the shutoff valve that manually turns the passenger oxygen on or off.

Figure 5I-14: Push-Pull Control Cables



With the oxygen bottle shutoff valve open, oxygen flows to the passenger oxygen system shutoff valve. The valve remains closed until the cabin altitude reaches 12,500 ft. At 12,500 ft, a barometric pressure switch opens the shutoff valve and pressurizes the passenger oxygen supply line. The system also can be manually activated by pulling out the PASSENGER MANUAL O’RIDE knob next to the PULL ON SYS READY oxygen supply knob.

Figure 5I-15: Oxygen Supply Control Knob

Oxygen flowing into the passenger supply line activates a pressure switch that illuminates the green PASS OXYGEN ON annunciator and illuminates the cabin lights at full brightness.Oxygen pressure at each mask container activates a plunger that opens the container doors.The masks deploy and hang from a lanyard approximately nine inches from the ceiling. Pulling on the lanyard releases a pin to allow the flow of oxygen to the mask. Reinserting the pin stops the flow of oxygen to the mask.After oxygen is no longer required by the passengers, pulling the OXYGEN CONTROL circuit breaker with the PASSENGER OXYGEN O’RIDE knob in removes power to the barometric switch; oxygen does not flow to the passenger supply lines.

Figure 5I-16: Mask Container Door




Oxygen DurationTypical oxygen duration guidelines for the King Air are shown in Table 5I-1, previous page. Always refer to the applicable operating manual for the oxygen duration table for your particular aircraft. Refer to the Oxygen Available with Partially Full Bottle graph to determine the percentage of a full bottle.Oxygen duration is computed for an auto-deployed type mask, rated at 3.9 liters per minute (LPM-NTPD) color coded orange and white, and approved for pressure altitudes up to 35,000 ft. The Oxygen Duration table is also used for the quick-donning diluter-demand crew oxygen masks. Each crew on-demand mask is counted as two masks (i.e., 8 passengers plus 2 crew equals 12 when figuring duration).

2500

2000

1500

1000

500

0 0 25

GA

UG

E PR

ESSU

RE

– PS

IG

60°C (140 °F)

0°C (32 °F)

-60 °C (-76 °F)

BOTTLE TEMPERATURE

50 75 100

PERCENT OF USABLE CAPACITY

Table 5I-2: Oxygen Available with Partially Full Bottle



PreflightCheck the oxygen pressure gauge to ensure there is sufficient pressure for the flight. Use Table 5I-2 to determine the percentage of a full bottle available. To find the oxygen bottle duration, multiply the percentage available by the full bottle Oxygen Duration chart to obtain time in minutes.As part of the Before Starting Engines checklist, check the bottle pressure gauge. Pull the oxygen supply control handle to PULL ON SYS READY to open the oxygen bottle shutoff valve. Check the crew oxygen masks in 100% to verify proper operation.

ServicingExercise extreme care when working with pure oxygen. When servicing the oxygen bottle, avoid making sparks and eliminate all potential ignition sources in the area. Make sure the oxygen supply control handle is in the closed position (handle in). Ensure that the area around the filler valve and tools, clothes, and hands are clean and free from grease and oils. Pure oxygen can cause oil and grease to ignite spontaneously.Access the filler valve by opening the access door on the right aft fuselage. Remove the protective cap from the valve and connect the oxygen servicing cart hose. Use only MIL-O-27210 aviator’s breathing oxygen. Do not use medical or industrial oxygen because they contain moisture that may freeze and clog system valves and lines at altitude.Fill the oxygen bottle slowly to prevent overheating. Fill the 22 cubic -ft bottle to 1,800 PSI at 70°F and the 49, 64, and 76 cubic-ft bottles to 1,850 PSI at 70°F. Disconnect the hose and replace the valve cap. Secure the access door.




Emergency EquipmentEmergency equipment includes: emergency locator transmitter safety equipment portable fire extinguishers life preservers.

Emergency Locator Transmitter The Emergency Locator Transmitter (ELT) transmits a downward sweeping tone on 121.5 MHz and 243.0 MHz as an aid in locating a downed aircraft. The ELT consists of a battery-powered transmitter behind the aft pressure bulkhead on the right fuselage, an antenna, and a remote arming/ power/testing switch near the transmitter. A spring-loaded door provides access to the transmitter switch. Some aircraft have a remote ARM/XMIT switch on the right instrument panel. An impact switch activates the system if the switch experiences a longitudinal force of approximately 5 Gs.

ELT TRANSMITTER

ON

OFF

AR

RES

ET

ARM - OFF - ON


ARM OFF ON


ANTENNA (TYPICAL)

ELT ACCESS DOOR (TYPICAL)

ELT ACCESS DOOR

Figure 5I-17: Emergency Locator Transmitter (ELT)



Safety EquipmentSafety equipment includes: first aid kit crash axe (some aircraft) two pairs of smoke goggles in the cockpit (some aircraft).

Fire ExtinguishersFederal Aviation Regulations (FARs) require a minimum of two fire extinguishers in the aircraft: one in the cockpit readily accessible to the crew, and one in the passenger compartment. Usually, there is one fire extinguisher under the copilot seat and one in the passenger cabin forward of the entrance door.The cockpit and cabin fire extinguishers use Halon 1301 as a fire extinguishing agent. Halon 1301 is a relatively non-toxic and non-corrosive agent that extinguishes a fire by chemically interfering with the combustion process.

Life PreserversFAR 91.509 requires a life preserver or flotation device for each occupant of an aircraft that travels overwater for more than 50 Nautical Miles (nm) from a shoreline. If the aircraft travels over-water more than 30 minutes flying time or 100 nm from the nearest shoreline, FAR 91.509 requires life preservers with a locating light for each occupant of the airplane and a life raft of suitable buoyancy and capacity to accommodate the occupants of the aircraft.




Cargo Door

DoorThe King Air B200C has a large, top-hinged cargo door with a smaller bottom-hinged airstair door. The cargo door swings upward for loading of large items (i.e., cargo or stretchers). There are no latch handles outside the cargo door; it can be opened only from inside the aircraft. Two covered, inboard latch handles, one upper aft and one lower forward handle, close or release the cargo door latches.

Operation

CAAUIOR After unlatching the bottom latch pins, close the lower forward latch handle access cover. If left open, the cover rotates on its hinge until part of it extends below the bottom of the open cargo door; the access cover suffers damage after closing the cargo door.

To operate the upper aft handle, open the access cover, press the black release button in the handle and rotate the yellow handle upward as far as it goes. This movement via cables releases four latches: two on the forward side and two on the aft side of the cargo door.To operate the lower forward latch handle, open the access cover, lift the orange lock hook from the stud on the yellow latch handle and rotate the handle aft as far as it goes. This movement via linkage transmits to four latch pins on the bottom of the cargo door. The pins move aft to disengage latch lugs mounted at the bottom of the cargo door frame.To open the cargo door, push out on the bottom of the door. After the door manually opens a few feet, gas springs take over and raise the door to the fully open position.To close the cargo door, pull it down and inboard. The gas springs resist the closing effort until the door is only open a few feet; then as the springs move over center, they begin applying a closing force.The door closes against an inflatable rubber seal around the cargo door frame. As the cabin pressurizes, air seeps into the rubber seal through small holes in the pressure vessel side of the seal. The higher the cabin differential pressure, the more the seal inflates. This is a passive seal system and has no mechanical connection to a bleed air source.



To latch the cargo door after it closes, rotate the lower forward latch handle forward until the orange lock hook engages the stud on the handle. Check the security of this handle by attempting to move it aft without raising the lock hook; it should not move. Close the access cover. Next, check the observation window at the lower aft corner of the cargo door. Ensure that the orange stripe on the latch pin linkage aligns with the orange pointer in the observation window.Rotate the upper aft latch handle down until the black release button pops up. Check the security of this handle by attempting to pull it out and up without pressing the release button; it should not move. Close the access cover. Ensure the orange stripes on the four rotary latches (two on each side of the cargo door) align with the notch in the plate on the door frame.Finally, switch the battery switch on and verify that the CABIN DOOR annunciator extinguishes. The annunciator illuminates with: the battery switch on and either the airstair door or the cargo door

open the battery switch off and the airstair door closed but not securely

latched.Prior to the first flight of the day, perform the Cabin/Cargo Door Circuitry Check in the Expanded Normal Procedures chapter.




Warning Systems

WarningsWarning systems on the King Air include: landing gear warning system stall warning system annunciator system.

Landing Gear WarningIf the landing gear is not down and locked during certain power lever settings, flap positions, and airspeed (some aircraft), the landing gear warning system alerts the crew by illuminating a red light and sounding a horn. The landing gear warning system consists of: landing gear-in-transit light pitot/static system Q-switch (BB324 through BB-453 without SI1047) landing gear warning horn switches power lever warning horn switches flap warning horn switches warning horn horn silence button.

BB-2 to BB-323Flaps Airspeed Power SilenceUP N/A 75 to 77% Silence ButtonAPR N/A 75 to 77% Add Power/Raise FlapsPAST APR N/A 75 to 77% Lower GearBB-324 to BB-453 without SI 1047UP <140 Kts 86% Silence ButtonAPR <140 Kts 86% Add Power/Raise Flaps/ Lower GearPAST APR <140 Kts 86% Lower GearBB-453 and Subsequent; BL-1 and SubsequentUP N/A 79 to 82% Silence ButtonAPR N/A 79 to 82% Silence ButtonPAST APR N/A 79 to 82% Lower Gear

Table 5I-3: Landing Gear Warning Horn Operation

NTTc: SI-1047; Landing Gear – Modification of Landing Gear Warning System (BB-324 through BB-453).



Ghmu-tn-TumnstaThe red landing gear-in-transit light in the landing gear handle illuminates as a warning when: landing gear handle in UP position with weight on wheels (on the ground) one or both power levers below certain N1 speed and landing gear

not down-and-locked one or all landing gear not fully retracted or fully extended warning horn silenced and not operational.

Wmuntni wtaSchsEach landing gear (nose and main) has a warning horn switch on the brace wired in series with the switches on the other landing gear and the power lever warning horn switches. The power lever warning switches are on power lever linkages. A cam on each linkage activates the switch with the power lever positioned below a set N1% RPM.The flap warning switch works with the landing gear and power lever switches to activate the warning horn. The flap switch activates at the approach flap setting (40% flaps); with less than 40% flaps, the switch is off.On BB-324 through BB-453 without SI 1047, there is a Q-switch in the pitot/static system that works with the landing gear warning system. These aircraft have a landing gear warning trigger airspeed of below 140 Kts; SI 1047 removes the Q-switch.

NphumatonTriggering of the landing gear warning system varies with aircraft serial number and modification status. Also, depending on the serial number the N1 speed at which the system activates, in combination with other conditions, varies from 75 to 86%. Refer to Table 5I-3.There is a horn silence button next to the landing gear control handle (Figure 5I-18). In certain circumstances pressing the button silences the landing gear warning horn; the warning light does not extinguish.Adding power above the set point, raising the flaps, or lowering the gear deactivates and rearms the landing gear warning system.

Figure 5I-18: Horn Silence Button




On aircraft prior to BB-324, the landing gear warning system activates with the landing gear retracted under the following conditions: Flaps up and one or both power levers retarded below the 75 to 77%

N1 position. Pressing the silence button stops the intermittent horn; the light remains illuminated.

Flaps in approach position with one or both power levers retarded below the 75 to 77% N1 position. Adding power and raising the flaps silences the horn and extinguishes the light.

Flaps beyond approach position in any power lever position. Lowering the gear silences the horn and extinguishes the light.

On aircraft BB-324 to 453 without SI-1047, the landing gear warning system activates with the landing gear retracted under the following conditions: Flaps up and one or both power levers retarded below the 86% N1

position; the light illuminates. Flaps up, one or both power levers retarded below the 86% N1

position and airspeed below 140 Kts; the warning horn sounds and the light illuminates. Pressing the silence button quiets the horn.

Flaps in approach or beyond approach position; light illuminates. Flaps in approach or beyond approach position and airspeed below

140 Kts. The horn sounds and the light illuminates. On aircraft BB-453 and subsequent and BL-1 and subsequent, the landing gear warning system activates with the landing gear retracted under the following conditions:

Flaps in approach or beyond approach position and one or both power levers retarded below the 79 to 82% N1 position. Pressing the silence button quiets the horn.

Flaps beyond approach position; horn sounds, light illuminates and cannot be canceled without lowering gear. On aircraft with the hydraulic landing gear system, the landing gear warning system activates with the landing gear retracted under the following conditions:

Flaps in approach or beyond approach position and one or both power levers retarded below the 79 to 82% N1 position. Pressing the silence button quiets the horn.

Flaps beyond approach position; horn sounds, light illuminates and cannot be canceled without lowering gear.



Stall Warning SystemThe stall warning system provides an aural warning of an imminent stall approximately: 5 to 13 Kts above the stall with the flaps retracted 5 to 12 Kts above the stall with 40° flaps (approach) 8 to 14 Kts above the stall with the flaps extended fully.

The system consists of a lift transducer, lift computer, warning horn, squat switch, and a test switch. The transducer on the leading edge of the left wing has a vane that reacts to the air flowing across the wing. The vane connects to a switch whose limits are dependent on flap position; with different flap settings the switch activation point changes.As the wing’s angle-of-attack reaches a critical point and a stall begins, air pressure moving against the vane activates a switch; flap position varies the activation point. The switch sends a signal to the lift computer that sounds the stall warning horn.

Figure 5I-19: Transducer

The lift transducer has DC powered heating elements in the vane and mounting plate that prevent the formation of ice. Normally, the heating elements receive 26 to 29 VDC in flight. Through a landing gear squat switch, the elements receive only 12 VDC with weight on the wheels.The STALL WARN switch on the copilot subpanel simulates a stall condition by applying voltage to an electromagnet in the lift transducer.Moving the switch to TEST simulates a near-stall condition by moving the lift transducer’s vane to the stall position, the transducer then sends a signal to the lift computer that activates the stall warning horn.




Annunciator SystemThe annunciator system consists of a warning annunciator panel on the center of the glareshield, a caution/ advisory annunciator panel on the center subpanel, and a red MASTER WARNING and yellow MASTER CAUTION light above the pilot and copilot center instrument panels.The warning and caution/advisory annunciators and the MASTER CAUTION lights have an automatic dimming feature. Normally, these annunciators and lights illuminate at full intensity. The lights dim if: a generator is on-line OVERHEAD FLOOD LIGHTS switch is ON MASTER PANEL LIGHTS switch is ON PILOTS FLIGHT LIGHTS are ON cockpit ambient light level is above a preset level as sensed by a

photoelectric cell.The MASTER WARNING lights do not dim.Pressing the PRESS TO TEST button to the right of the warning annunciator panel tests all the bulbs in the warning and caution/advisory annunciators, the MASTER CAUTION lights, and the MASTER WARNING lights.

Wmuntni AnngnStmaou Pmnhl The warning annunciator panel contains either 12,16, or 20 individual red lamps with legends (depending on aircraft model) that illuminate when there is a condition that requires the immediate attention and action of the flight crew. Whenever a warning annunciator illuminates, the MASTER WARNING lights flash. Pressing a MASTER WARNING light extinguishes the light; correcting the fault extinguishes the warning annunciator.

Figure 5I-20: Warning Annunciator Panel



Cmgaton/Advtsouy AnngnStmaou PmnhlThe caution/advisory annunciator panel consists of 30 or 36 (depending on the aircraft) individual yellow and green lamps with legends. The green annunciators illuminate to indicate system operation.The yellow annunciators illuminate when a condition develops that requires immediate attention of the flight crew but not immediate action. When a yellow caution annunciator illuminates, the yellow MASTER CAUTION lights flash. Pressing a MASTER CAUTION light extinguishes it; correcting the fault extinguishes the yellow annunciator.BB-2 through BB-666 have a CAUTION toggle switch on the copilot sub-panel. Moving the spring-loaded switch momentarily to OFF extinguishes the illuminated yellow caution annunciators and illuminates the green CAUT LGND OFF annunciator. If another fault illuminates a caution annunciator, all annunciators turned off with the CAUTION switch re-illuminate. Momentarily selecting ON also re-illuminates a switch-extinguished annunciator(s).

Figure 5I-21: Caution/Advisory

Annunciator PanelFigure 5I-22: MASTER CAUTION

Lights Flash




Annunciator PanelsKING AIR 200(Prior to BB 1439; Prior to BL-139)

L FIRE ENG INVERTER CABIN DOOR ALT WARN R ENG FIRE


L OIL PRESS L GEN OVHT A/P TRIM FAIL R GEN OVHT R OIL PRESS

L CHIP DETECT L BL AIR FAIL A/P DISC R BL AIR FAIL R CHIP DETECT

L DC GEN

L IGNITION ON

L ICE VANE EXT L AUTOFEATHER

L ICE VANE

BRAKE DEICE ON

L BL AIR OFF

HYD FLUID LOW

BATTERY CHARGE

ELEC TRIM OFF LDG/TAXI LIGHT

RVS NOT READY

DUCT OVERTEMP

EXT PWR

AIR CND N1 LOW PASS OXY ON

FUEL CROSSFEED R BL AIR OFF

R DC GEN

R IGNITION ON

R ICE VANE EXT R AUTOFEATHER

R ICE VANE

PROP SYNC ON

KING AIR B200(BB 1439 and subsequent; BL-139 and subsequent)

L ENG FIRE INVERTER DOOR UNLOCK ALT WARN R ENG FIRE


L OIL PRESS L GEN OVHT

L BL AIR FAIL A/P FAIL

A/P TRIM FAIL

R BL AIR FAIL

R GEN OVHT R OIL PRESS

L DC GEN

L IGNITION ON

L ENG ANTI-ICE

L AUTOFEATHER

L ENG ICE FAIL

BRAKE DEICE ON

L BL AIR OFF

HYD FLUID LOW

BATTERY CHG

ELEC TRIM OFF

LDG/TAXI LIGHT

RVS NOT READY

DUCT OVERTEMP

EXT PWR

AIR CND N1 LOW

PASS OXYGEN ON

FUEL CROSSFEED R BL AIR OFF

R DC GEN

R IGNITION ON

R ENG ANTI-ICE

R AUTOFEATHER

R ENG ICE FAIL

L CHIP DETECT

ELEC HEAT ON

R CHIP DETECT







Annunciator Cross ReferenceAnnunciator DescriptionA/P DISC On Sperry autopilot equipped aircraft, illumination of the red A/P DISC (autopilot disconnect)

annunciator indicates an intentional or unintentional disconnection of the autopilot due to either the pressing of the copilot AP/YD & Trim Disc, autopilot Test, or go-around button, or a system malfunction.

A/P TRIM FAIL On Sperry autopilot equipped aircraft, the red A/P TRIM FAIL annunciator illuminates if there are improper trim commands, if there is a trim runaway, or if a trim system failure occurs. Do not re-engage the autopilot.

AIR COND N1 LOW The green AIR COND N1 LOW annunciator illuminates if the right engine N1 speed is too low (below 62% N1RPM) to power the Freon compressor without lugging the engine. An N1 speed sensor restricts electrical power to the compressor clutch until N1 speed is sufficient to prevent engine lugging with the compressor operating.

ALT WARN If cabin altitude exceeds 12,500 ft, the cabin altitude pressure warning switch closes and illuminates the red ALT WARN annunciator.

BATTERY CHG The yellow BATTERY CHG annunciator illuminates if the charging current exceeds 7.5 ±1.0 amps after a six second delay. After a battery assisted engine start, the BATTERY CHARGE annunciator illuminates for approximately five minutes as the battery recharges. If the annunciator illuminates in flight, turn the battery off and check the battery charging rate.

BRAKE DEICE ON (optional)

The green BRAKE DEICE ON annunciator illuminates when the brake deicing system bleed air shutoff valves are open and providing hot bleed air to the main wheelbrakes. After weight is off the wheels, the system automatically turns brake deicing off after 10 minutes; the annunciator extinguishes.

CABIN DOOR Illumination of the red or yellow CABIN DOOR or CABIN/CARGO DOOR annunciator indicates that the door is not closed, latched, and secured. The door warning system uses two switches: one actuated by the lower forward latch bolt and one actuated by the door latch arm. If either switch does not close, the annunciator illuminates. Do not check the door in flight. Land as soon as possible.

CAUT LGND OFF On BB-2 through BB-666, illumination of the green CAUT LGND OFF annunciator indicates that an uncorrected fault condition exists with the caution annunciator (caution/advisory panel) turned off. Momentarily selecting ON with the CAUTION switch re-illuminates the caution annunciator.

DUCT OVERTEMP The yellow DUCT OVERTEMP annunciator illuminates when temperatures in the cabin heating duct system reach 300°F. Select a lower temperature with the CabinTemp knob, select the HI position with the Vent Blower switch, and push the Cabin Air knob to increase airflow.

ELEC HEAT ON Advisory light indicates power to the electrical heating elements. The light must be extinguished before the blowers are turned off.

ELEC TRIM OFF On Sperry autopilot equipped aircraft, the green ELEC TRIM OFF annunciator illuminates when the electric elevator trim system disconnects after pressing the trim disconnect switch with the ELEV TRIMs witch ON (power on). Cycling the ELEV TRIM switch from ON to OFF, then to ON resets the system.

EXT PWR The yellow EXT PWR (external power) annunciator illuminates with a ground power unit (GPU) connected to the aircraft.

FUEL CROSSFEED

The green FUEL CROSSFEED annunciator illuminates to indicate operation of the fuel crossfeed system. Moving the Crossfeed Flow lever lock switch out of the center (OFF) position energizes the standby pump on the side supplying fuel, closes the motive flow valve on the side being supplied, and illuminates the annunciator.

HYD FLUID LOW The yellow HYD FLUID LOW annunciator indicates a low hydraulic fluid level in the hydraulic power pack supplying the landing gear system. Pressing the HYDFLUID SENSOR TEST button tests the annunciator (BB-1158, BB-1167, BB-1193 and subsequent; BL-73 and subsequent; prior aircraft with BeechKit P/N 101-8018-1).

INVERTER INST INV

Illumination of the red INVERTER or INST INV annunciator indicates loss of 115 VAC to the inverter warning relay from the selected inverter; the selected inverter is inoperative. Check inverter voltage and frequency; normal indications are 107 to 120 VAC, and 396 to 404 Hz. Select the opposite inverter to regain AC power.

L AUTOFEATHER R AUTOFEATHER

Illumination of the green L/R AUTOFEATHER annunciators indicates arming of the propeller automatic feathering system with the power levers advanced above 90% N1. The system de-arms with the throttles retarded below 90% N1.



L BL AIR FAIL R BL AIR FAIL

The red L/R BL AIR FAIL annunciators illuminate if there is a failure (i.e., rupture) of the associated bleedair supply line from the engine to the cabin. The system uses a pressurized polyflow tubing that parallels the bleed air supply lines. A normally open switch in the polyflow tubing closes to illuminate the associated BLAIR FAIL annunciator when hot air from a failed bleedair supply line melts the tubing. On the affected engine, select INSTR & ENVIR OFF with the BLEED AIRVALVE and monitor the engine instruments.

L BL AIR OFF R BL AIR OFF

Illumination of a green BL AIR OFF annunciator indicates the respective bleed air shutoff valve is closed.

L CHIP DETECT R CHIP DETECT

The red L/R CHIP DETECT annunciators illuminate if sufficient ferrous (magnetic) particles accumulate to bridge the contacts on the engine chip detector.

L DC GEN R DC GEN

Illumination of the yellow L DC GEN or R DC GEN annunciator indicates that the respective generator is offline. Depending on the aircraft serial number, turn the respective generator switch OFF then ON (prior to BB-88) or OFF, RESET for 1 second, then ON to reset the generator. If the generator does not reset, turn the generator OFF.

L ENG FIRE R ENG FIRE

Illumination of a red ENG FIRE annunciator indicates a possible fire in the respective engine nacelle. The system uses three infrared radiation sensors or loop system (see Fire Protection) that cover the entire engine nacelle. (Moisture or reflected sunlight can cause erroneous warnings on infrared sensors.)


The respective red FUEL PRESS annunciator illuminates when fuel pressure decreases below 9 to 11 PSI. The annunciator extinguishes when fuel pressure exceeds 9 to 11 PSI. During engine start the annunciator illuminates then extinguishes once fuel pressure builds. If the annunciator illuminates in flight, turning on the standby fuel pump increases fuel pressure and extinguishes the light.

L GEN OVHT R GEN OVHT (optional)

Illumination of a red GEN OVHT annunciator indicates generator frame temperature above 315°F. If an annunciator illuminates in flight, turn the affected generator switch OFF. If the annunciator fails to extinguish, shut down the engine. If the annunciator illuminates on the ground, reduce the electrical load, turn the affected generator OFF, or increase engine N1 RPMto 70% or above.

L ICE VANE R ICE VANE

If an ice vane does not reach the extended position within 15 seconds after placing the respective ICEVANE switch in EXTEND, the respective yellow ICEVANE annunciator illuminates. If the vane fails to extend, pull the respective ICE VANE CONTROL CB then pull the ICE VANE EMERGENCY – MANUAL EXTENSIONT-handle to extend the ice vane and bypass door; do not use the switches to retract or extend the ice vane and bypass door after manual extension. Continue to operate the ice vane doors manually until maintenance corrects the fault.

L ICE VANE EXT R ICE VANE EXT

The respective green L/R ICE VANE EXT annunciator illuminates when the associated engine inlet ice vane and bypass door (inertial separator system) reach the extended position through electric motor or manual extension.

L IGNITION ON R IGNITION ON

The green IGNITION ON annunciators illuminate to indicate power to the engine ignition system. The annunciator (s) illuminate when the IGNITION AND ENGINE START switch is in ON and the ignition power relay energizes or the auto ignition system is on and energizing the ignition system.

L OIL PRESS R OIL PRESS (optional)

Illumination of a red OIL PRESS annunciator indicates that engine oil pressure is below 40 PSI. Shut down the engine and land as soon as possible.

LANDING LIGHT LDG/TAXI LIGHT

Illumination of the green LANDING LIGHT or LDG/TAXI LIGHT annunciator indicates that the landing and/or taxi lights are on with the landing gear retracted; turn the lights OFF.

MASTER CAUTION

The yellow MASTER CAUTION lights flash whenever a yellow annunciator in the caution/advisory annunciator panel illuminates indicating a fault or condition that requires the pilot’s attention but not immediate action. To extinguish the MASTER CAUTION lights, press either of the lights.

MASTER WARNING

The red MASTER WARNING lights flash whenever a red annunciator in the warning annunciator panel illuminates to indicate a fault requiring immediate attention. Pressing either light extinguishes them.

PASS OXYGEN ON The green PASS OXYGEN ON annunciator illuminates when a switch in the passenger oxygen supply line senses pressure in the lines. On aircraft with the auto-deployment oxygen system, a barometric pressure switch automatically activates the passenger oxygen system by supplying pressure to system, illuminating the PASS OXY ON annunciator, and deploying the passenger oxygen masks. If the annunciator does not illuminate after the ALT WARN annunciator illuminates, pull the PASSENGER MANUAL O’RIDE valve to deploythe oxygen masks.

PROP SYNC ON (optional)

Illumination of the yellow PROP SYNC ON annunciator indicates that the propeller synchrophaser is ON with the landing gear extended; turn the synchrophaser OFF.

RVS NOT READY The yellow RVS NOT READY (reverse not ready) annunciator illuminates when the propeller controll ever (s) are in the low RPM, high pitch position with landing gear down. The annunciator extinguishes when both propeller levers are in the high RPM, low pitch position.

Pneumatic Systems 5J

King Air 200 5J-1December 2011


ContentsPneumatic Systems

SchematSc: Pneumatic System .....................................................5J-5Pneumatic System

SchematSc: Bleed Air Valve Switches ...........................................5J-8Bleed Air Valve Switches ........................................................................5J-9Pneumatic Instrument Air .......................................................................5J-9Bleed Air Distribution ............................................................................ 5J-10

Vacuum System ................................................................................. 5J-10Flight Hour Meter ............................................................................... 5J-12Rudder Boost System ........................................................................ 5J-12

Bleed Air Warning System .................................................................... 5J-12Pressurization System

Distribution ............................................................................................ 5J-15Heat Exchanger ..................................................................................... 5J-15Flapper Valve Unit ................................................................................. 5J-15Muffler/Mixing Plenum .......................................................................... 5J-16Controller/Outflow Valve ....................................................................... 5J-16Safety Valve ............................................................................................ 5J-17CABIN PRESS Switch ........................................................................... 5J-17Fresh Air Ventilation .............................................................................. 5J-18Cabin Pressurization Indicating System ............................................. 5J-19Cabin Pressurization Failure Warning System ................................... 5J-19

SchematSc: Air Distribution System ............................................ 5J-21SchematSc: Air Distribution System S/Ns BB-1439, 1444 and Subsequent; BT-35 and Subsequent ........................... 5J-23

Air Conditioning SystemDistribution ............................................................................................ 5J-25

SchematSc: Air Conditioning System .......................................... 5J-26Environmental Control Panel ............................................................... 5J-27

CABIN TEMP Controls ...................................................................... 5J-27Vent Blower Control ........................................................................... 5J-28Cockpit Comfort Controls ................................................................... 5J-29

Radiant Heating System ....................................................................... 5J-30

King Air 2005J-2December 2011


Electric Heating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5J-30Freon Air Conditioning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5J-31

Compressor ......................................................................................5J-31Condenser ........................................................................................5J-32Limit Switches ...................................................................................5J-32Optional Evaporator Unit ..................................................................5J-33

Servicing and ProceduresPreflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5J-35Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5J-35

Cabin Pressure Too High (Overpressurization) ................................5J-35Under Pressurization ........................................................................5J-35Loss of Pressurization ......................................................................5J-36DUCT OVERTEMP Annunciator Illuminated ....................................................................5J-36BL AIR FAIL Annunciator Illuminated ....................................................................5J-36Crack in Any Side Window or in Windshield .....................................5J-36Smoke and Fume Elimination ...........................................................5J-37Rapid Decompression/Emergency Descent .....................................5J-37

Data SummariesBleed Air System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5J-39Air Conditioning/Heating System . . . . . . . . . . . . . . . . . . . . . . . . . . . 5J-40Pressurization System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5J-40

Pneumatic Systems



Pneumatic SystemsThis chapter describes the systems that extract, distribute, and control engine bleed for aircraft pressurization, cooling and heating, and deice protection (see Ice and Rain Protection chapter).The Pneumatic System extracts bleed air from the engine’s fourth stage compressor section (P3) and transfers it to various other systems (i.e., surface deice, rudder boost, brake deice, bleed air warning, and door seal). A venturi-ejector in the system creates a vacuum source for the air-driven gyros, pressurization control, and deflation of the deice boots.The Pressurization System controls cabin altitude through operating outflow valves. This system controls the outflow of conditioned air through outflow valves to produce cabin pressure.The Environmental System provides cabin heating by regulating engine bleed air temperatures. Cabin cooling, when required, utilizes an automotive-type Freon system. During unpressurized flight, a ram air scoop provides fresh air ventilation. The air conditioning section discusses aircraft heating and cooling.





5J-5For Training Purposes Only

Pneumatic Systems

Pneumatic System

DEICE DISTRIBUTOR

VALVE

TIMER



RUDDER BOOST

OFF

OFF

MANUAL

SINGLE CYCLE

TO TAIL BOOTS

TO WING BOOTS

18 PSI PRESSURE

REG

ENVIR OFF

INSTR & ENVIR OFF


LEFT RIGHT

RIGHT P INSTRUMENT AIR

3

N. O. INSTRUMENT AIR VALVE

RIGHT BRAKE DEICE VALVE

N. O.

INSTRUMENT AIR VALVE

LEFT BRAKE DEICE VALVE

LEFT P INSTRUMENT AIR

3

LEFT BRAKE DEICE MANIFOLD

RIGHT BRAKE DEICE MANIFOLD

15 PSI PRESSURE

REG

TIMER

BRAKE DEICE ON

R BL AIR FAIL

20 20

1010

L R

AIR

23

25 30

35

2119

17

15

3

1

2 4

5

6 7 8 9

10 11

13

0

CABIN

ALT

1000 FT

RATE CABIN ALT

ACFT ALT 100 FT

15

M

N

M A X

FLIGHT

0 4 5 6 2 HOURS 1/10

GYRO SUCTION

INCHES OF MERCURY

3 4 6

PNEUMATIC PRESSURE

0

10

20

PNEUMATIC PRESSURE

GAUGE

COPILOT'S ATTITUDE

INDICATOR

COPILOT'S TURN & SLIP INDICATOR

SUCTION GAUGE

VACUUM REGULATOR

PRESSURIZATION CONTROLLER

VACUUM AIR FILTER

CABIN AIR FILTER

4 PSI REG

N. O.

TO DOOR SEAL

N.O. PRESET SOLENOID

N.C. DUMP SOLENOID

TO DUMP VALVE

FLT HOUR METER

N. C.

N. C.

N. C.

N. C.

L BL AIR FAIL

FILTER

UP

DNGEAR UPLOCK

P SWITCH LEFT RUDDER BOOST SERVO

RIGHT RUDDER BOOST SERVO

UNREGULATED

REGULATED

VACUUM

RUDDER




5A

5A

5A

5A

5A

5A

5A




AIR LEAK

TO OUTFLOW VALVE


5J-6 For Training Purposes Only


Pneumatic Systems



Pneumatic SystemEngine bleed air drawn from the fourth stage of each engine compressor is referred to as P3 air. Since engine speeds can vary from an idle RPM of 52–56% to takeoff RPM of 100%, bleed air pressures and temperatures also vary widely because of compression.Engine compressor efficiency at 100% RPM is approximately 9:1. As a result, P3 pressure at 100% RPM at sea level is approximately 130 PSI; bleed air temperature is approximately 650°F. P3 pressure at 100% RPM at 31,000 ft where atmospheric pressure is only 4 PSI, however, is approximately 36 PSI with temperature ranging between 100 and 200°F depending on outside temperature at that altitude. On the ground at idle, P3 pressure is about 20 PSI.These wide ranges of pressures and temperatures are regulated and controlled to provide needed cabin pressurization and environmental heating and cooling. In addition, certain subsystems within the aircraft operate pneumatically.After P3 air leaves the engine, it divides into two separate systems: Environmental (ENVIR) and pneumatic Instrument (INSTR) air. The larger (ENVIR) of these two bleed air lines provides the proper volume of air required for heating, cooling, and pressurization.The smaller line (INSTR) provides pneumatic pressure for various subsystems and heat for brake deicing, if installed.

5J-8 For Training Purposes Only King Air 200December 2011

Bleed Air Valve Switches

R BLEED AIR FAIL

ENVIRONMENTAL MIXING PLENUM

BLEED AIR WARNING SWITCHES

ENGINE BLEED AIR

AMBIENT AIR

AFT FIREWALL

FLOW CONTROL UNIT

ENVIRONMENT BLEED AIR SHUTOFF VALVE (N.C.)

PLUGS

PNEUMATIC BLEED AIR SHUTOFF VALVE (N.O.) AIR INLET

AIR-TO-AIR HEAT EXCHANGER

CABIN HEAT CONTROL VALVE

LEFT BLEED AIR WARNING LINE (POLY FLOW TUBING)

RIGHT BLEED AIR WARNING LINE (POLY FLOW TUBING)

BLEED AIR WARNING LINE

PRESSURIZATION BLEED AIR

PNEUMATIC BLEED AIR

AMBIENT

COOLED BLEED AIR

L BLEED AIR FAIL

18 PSI

Pneumatic Systems



Bleed Air Valve SwitchesThe BLEED AIR VALVE switches on the copilot instrument subpanel control the flow of bleed air entering the cabin via the environmental and instrument air valves. Switch positions are OPEN/ ENVIR OFF/ INSTR & ENVIR OFF.The OPEN position opens both the environmental and the instrument air valves for the selected side (LEFT or RIGHT). The ENVIR OFF position closes the environmental air valve and opens the instrument air valve for the selected side.The INSTR & ENVIR OFF position closes both the environmental air valve and the instrument air valve for the selected side. Because this selection shuts off the flow of bleed air to the cabin, it also provides maximum cooling for ground operation.The ENVIR valve is a normally closed valve while the INSTR valve is a normally open valve.

Figure 5J-1: BLEED AIR VALVE Switches

Pneumatic Instrument AirBleed air flows through the instrument air valve and check valve to a common tee in the fuselage. The check valve in the left and right lines prevents reverse bleed air flow if an engine becomes inoperative.The tee distributes bleed air to a pressure regulator valve, under the right seat deck, to provide 18 ±1 PSI to the sub-systems. The regulator contains a 21 PSI relief valve to limit bleed air flow if a failure of the regulator occurs. Bleed air temperature drops to approximately 21°C (70°F) above ambient air at the pressure regulator valve because of minimal air flow and heat loss in the pneumatic plumbing. A PNEUMATIC PRESSURE gauge on the copilot right subpanel monitors the system; it displays pressure from 0 to 21 PSI.


Bleed Air Distribution

VmSuue ysaheBleed air from the 18 PSI pressure regulator passes through a venturi-ejector to produce vacuum for the following items: pressurization controls (see Pressurization, this chapter) air-driven flight instruments (see Avionics chapter) deflation of the deice boot.

Unregulated vacuum produced by the ejector continuously pulls the wing deicing boots smoothly against the underlying wing airfoil.In addition, vacuum is pulled from the unpressurized nose compartment filter through a vacuum regulator to the ejector. This regulated vacuum operates air-driven instrument gyros and the cabin outflow and safety valve of the pressurization system.A GYRO SUCTION gauge on the copilot right subpanel monitors regulated vacuum in INCHES OF MERCURY. Normal indicates range from 2.8 to 4.3 inches Hg above 15,000 ft and 4.3 to 4.9 inches Hg below 15,000 ft.

Figure 5J-2: GYRO SUCTION gauge

Pneumatic Systems



urfmSh DhtSh ysahePressure-regulated bleed air through the instrument air valve supplies pressure to inflate the pneumatic deice boots (see Ice and Rain Protection). Bleed air at 18 PSI passes through a deice distributor valve to inflate the boots. The venturi-ejector creates vacuum to deflate and hold the deice boots tightly down when not in use.The DEICE CYCLE switch on the pilot inboard subpanel controls the deice distributor valve. The 5A SURF DEICE CB on the copilot side panel receives power from the No.1 Dual-Fed bus to supply power to the boot system.

Figure 5J-3: EICE CYCLE switch

Cabin Door Seal Bleed air tapped downstream of the 18 PSI pressure regulator passes through a 4 PSI regulator. From this regulator, the air passes through a normally open door valve controlled by the left landing gear squat switch and pressure switch. As the aircraft lifts off, the squat switch opens to remove power from the normally open door seal valve; the seal inflates. Bleed air at 4 PSI inflates the door seal.

Figure 5J-4: Left Landing Gear Squat Switch


Fltgca Hour MhahrThe flight hour meter is activated by 28 VDC power through the RH landing gear squat switch and a ground provided by a regulated 18 PSI pressure switch.

Ruddhr Boosa ysaheLittle P3 from each engine nacelle continues into the rear cabin floor area of the aircraft. Prior to reaching the isolation check valves and the 18 PSI regulator, pressure is teed off and routed to a pressure differential switch. Engines producing takeoff power exert equal (unregulated) pressure on each side of this switch.Should either engine fail, P3 pressure on the failed engine reduces quickly. Differential P3 pressure of 60-65 PSI depresses the pressure differential switch to the failed side; this activates 28 VDC power to a normally closed air valve. When this valve opens, 15 PSI regulated air dumps into a pneumatic cylinder (servo) in the tailcone; the cylinder is yoked to the applicable rudder control cable to assist pilot’s rudder input for directional control.The ON/OFF control switch for the rudder boost system is immediately aft of the power quadrant.The No.2 Dual-Fed bus provides power for the rudder boost system. A failure of either bleed air source that requires closing the applicable bleed air valve deactivates the rudder boost system; i.e., either bleed air switch in OFF disables rudder boost.

Figure 5J-5: ON/OFF Control Switch

Bleed Air Warning SystemPolyflow tubing parallels the bleed air lines from the engines to the cabin; 18 PSI bleed air supplies the line with pressure. If a bleed air line should rupture, excessive heat 93°C (200°F) melts the adjacent tubing to release pressure. A normally open switch in the line under the copilot floor closes when pressure is reduced. The closed switch completes a circuit to illuminate the respective BL AIR FAIL annunciator.

Pneumatic Systems



Pressurization SystemP3 bleed air pressurizes the aircraft. A Flow Control Unit (FCU) in the nacelle of each engine mixes ambient air with the hot engine bleed air and regulates the flow mixture and volume into the cabin. Output volume of FCU is 8-12 lbs/min.The mixing of ambient air is a function of atmospheric pressure and Outside Air Temperature (OAT).As the aircraft climbs, the percentage of bleed air flowing to the cabin increases while the percentage of ambient air decreases. Above approximately 18,000 ft, the air flow to the cabin environmental system consists of bleed air only. The bleed/ambient air mixing schedule ensures that the cabin environmental system receives air at temperatures adequate for cabin heating. Ambient air mix ceases completely when OAT reaches -30°F.Prior to BB-1180, FCUs are pneumatically controlled. BB-1180 and subsequent have electronic FCUs that control ambient air mixing through OAT sensing only. Ambient air ceases below 10°C (50°F) in the latter system.Aircraft equipped with PT6A-41 engines (200) maintain a6.1 ±0.1 PSI differential. The cabin pressure altitude is 3,900 ft at an altitude of 20,000 ft; and 9,900 ft at 31,000 ft.Aircraft equipped with PT6A-42 engines (B200) maintain a6.5 ±0.1 PSI differential. The cabin pressure altitude is 2,800 ft at an altitude of 20,000 ft; 8,600 ft at 31,000 ft; and 10,400 ft at 35,000 ft.

Figure 5J-6: Flow Control Unit (FCU)


Pressurization System

NO1

DUAL

FED

BUS

FILTER

MOISTUREDRAIN OUTFLOW

VALVE

SAFETYVALVE

N.O.PRESETSOLENOID N.C.

DUMPSOLENOID

ALTITUDE RATE

UP

DOWN

L.G.SAFETYSWITCH

TEST

DUMPPRESS

PRESSCTRL

CABIN PRESSURE CONTROL SWITCH

TODOORSEAL

N.O.SOLENOID

PRESSIN

4 PSI

STATIC

SUCTION

CABIN AIR MIXED CABIN AND SUCTION

5A

TOVACUUM SOURCE

STATIC AIR

CABINAIR INPUT

RAM AIR DOOR LATCH

ENVIRONMENTAL BLEED AIR

Pneumatic Systems



DistributionAir from the flow control unit flows to the air-to-air heat exchanger. Integral to the FCU is the normally closed environmental air valve. It is powered open when the BLEED AIR VALVE switch is in the OPEN position by 28 VDC (through the LH BLEED circuit breaker on the No.1 Dual-Fed bus and the RH BLEED circuit breaker on the No.2 Dual-Fed bus).A bypass valve regulates the amount of air mixture routed through or around the respective heat exchanger. This conditioned air flows through flapper valves and a muffler within the pressure vessel before it is circulated further.

Heat ExchangerThe radiator-type heat exchanger in each wing root uses ram air to further cool the bleed air passing through it. The cooled air mass re-mixes with the air that bypassed the exchanger and proceeds to the flapper valve unit. Ram air passing over the heat exchanger is directed overboard through louvers on the under side of the wing root.

Figure 5J-7: Radiator-Type Heat Exchanger

Flapper Valve UnitThe air masses from both engines unite underneath the cabin floor at the flapper valve unit. One-way check valves in the unit isolate environmental air from either the left or right engine should one bleed air source fail.

Pressurization System Limitations

Maximum Cabin Pressure Dif-ferential – 6.1 PSI (200); 6.6 PSI (B200).


Muffler/Mixing PlenumThe muffler reduces airflow noise and directs the air to the mixing plenum. As a final temperature adjustment, the bleed air mixes in the mixing plenum with recirculated cabin air provided by the forward vent blower. A partition divides the mixing plenum; one section supplies the floor-outlet ducts, and the other the ceiling-outlet ducts.Warm air, tapped from the mixing plenum, is routed to the crew foot outlets and the windshield defroster. Controls for the pilot and copilot foot outlets and defrosters are on the panels immediately left and right of the respective control yokes (see Air Conditioning section, this chapter, for discussion on controls).

Controller/Outflow ValveThe pressurization controller on the center pedestal uses vacuum to maintain selected cabinaltitude and controls the cabin rate of climb/descent by regulating the position of the outflow valve. The outflow valve acts also as a safety valve and as a negative pressure valve.Air flows out of the aircraft through the outflow valve on the aft pressure bulkhead. The valve establishes a metered flow of air from the cabin to maintain the selected cabin altitude or to establish a pressure rate of change that corresponds to the aircraft’s rate of climb or descent. The outflow valve responds to signals from the pressurization controller.Before flight, set the pressure controller to indicate either 500 ft above the field elevation on the outer CABIN ALT scale or 1,000 ft above the cruise altitude on the inner ACFT ALT scale, whichever provides the higher cabin altitude. During descent, reset the controller to 500 ft above destination pressure altitude to ensure an unpressurized cabin prior to touchdown. Use the chart in the Operating Handbook.

Figure 5J-8: Pressurization Controller

Pneumatic Systems



Safety ValveA vacuum-operated safety valve on the aft pressure bulkhead and adjacent to the outflow valve provides pressure relief if the normal outflow valve fails or DUMP is selected to depressurize the aircraft.In addition, the safety valve keeps the aircraft unpressurized while on the ground and acts as a negative pressure relief valve during rapid descents (limits negative pressure to 0.1 PSI).A solenoid dump valve allows vacuum pressure to open the safety valve when actuated by 28 VDC from either the left landing gear squat switch or the CABIN PRESS DUMP switch. At lift-off, the squat switch opens to close the safety valve and activate the outflow valve.

CABIN PRESS SwitchThe CABIN PRESS/DUMP/PRESS/TEST switch is left (200) or forward (B200) of the pressurization controller on the pedestal. The center position, PRESS, closes the safety valve so that the pressurization controller regulates the outflow valve during flight. The cabin pressurizes in this position.The aft TEST position is for ground testing of the pressurization system. This position closes the safety valve by removing power and bypassing the landing gear squat switch. The DUMP, locked position opens the safety valve so the cabin depressurizes.

Figure 5J-9: CABIN PRESS/DUMP/PRESS/TEST Switch


Fresh Air VentilationOn S/N 001 and 002 with the cabin unpressurized and the CABIN PRESS switch in DUMP, ambient fresh air enters the ram air scoop on the left side of the unpressurized nose section. On subsequent aircraft, the air enters through condenser air-inlet louvers on the right side of the nose. The fresh air travels to a fresh air (ram air) door. The ram air door is held closed by a spring assembly, cabin pressure, or an electro-magnetic latch, which is activated with DC power. The pressurization switch is either in the PRESS or TEST positions. When the pressurization switch is placed in the DUMP position, the preset solenoid closes,removing vacuum from the outflow valve and it closes. The door seal solenoid closes, the door seal deflates, the dump solenoid opens, vacuum pulls the safety valve open, and power is taken away from the ram air door magnetic latch. As soon as the cabin differential reduces below approximately 1-1.5 PSID, ram air entering from the condenser cross-over duct pushes the door open and assists in the removal of contaminated air from the cabin.A vent blower delivers ram air to the cabin. The ambient air mixes with recirculated cabin air, flows through the vent blower, then through the forward evaporator, which, if operating,cools the air. The air then travels into the mixing plenum for delivery to the ceiling and the floor outlet ducts. An eyeball to each ceiling outlet controls the airflow and direction.A second source of fresh air ventilation is available during pressurized and non pressurized operation. The bleed air heating system mixes air with recirculated cabin air that enters the cabin through the floor registers.Prior to the 1979 models, regulate the volume of air with the sliding handle at the side of each inboard facing register. On subsequent models, the CABIN AIR control knob on the copilot subpanel controls the volume of air from the floor registers (Figure 5J-10).

Figure 5J-10: The CABIN AIR Control Knob

Pneumatic Systems



Cabin Pressurization Indicating SystemThe cabin pressurization indicating instruments on the center pedestal monitor operation of the cabin pressurization system. These include: CABIN ALT/differential pressure indicator that displays actual cabin

pressure altitude and differential pressure. CABIN CLIMB (vertical speed indicator) that shows rates of climb/

descent of cabin pressure in ft per minute.

Figure 5J-11: Vertical Speed Indicator

Cabin Pressurization Failure Warning SystemA pressurization warning system alerts the crew when cabin altitude reaches 12,500 ft. The system consists of a pressure-sensing switch to illuminate the ALT WARN annunciator.In addition, an auto-deployment oxygen system sensing switch closes to pressurize the passenger oxygen line and deploy the oxygen masks (see Miscellaneous chapter).





Pneumatic Systems

Air Distribution System

REFRIGERANTCOMPRESSOR

PNEUMATIC THERMOSTAT

MODULATING VALVE

FLOW CONTROL UNIT

ENVIRONMENTAL BLEED AIR SHUTOFF VALVE (N.C.)

PNEUMATIC/ INSTRUMENT AIR SHUTOFF VALVE (N.O.)

FIREWALL AIR TO AIR HEAT EXCHANGER

CABIN - HEAT CONTROLVALVE

FLAPPERVALVES

VENT BLOWER

FRESH AIR\RAM AIR DOOR (CLOSED WHEN PRESSURIZED)

RAM AIR SCOOP

FWD EVAPORATOR

MIXING PLENUM

CEILING DUCT/ FLOOR DUCT DIVIDER

DUCTOVERTEMP SENSOR

CONDENSER

RECEIVER DRYER

CONDENSER BLOWER OUTLETAIR

FWD PRESSURE BULKHEAD

CREW HEAT DUCT

WINDSHIELD DEFROSTER (ON GLARESHIELD)

CABINAIR CONTROL VALVE

AIR INLET SCOOP

ENVIRONMENTAL BLEED AIR SHUTOFF VALVE (N.C.)

PNEUMATIC THERMOSTAT

AMBIENT AIR MODULATING VALVE FIREWALL

PNEUMATIC/ INSTRUMENT AIR SHUTOFF VALVE (N.O.)

AFT EVAPORATOR

AFT EVAPORATOR BLOWER

CABIN-HEAT CONTROLVALVE AND 30° POSITION SWITCH


FLOOR OUTLET

CEILING OUTLET


NORMALOUTFLOW VALVE

SAFETY/DUMPVALVE

DOOR (COOLED AIR TO FLOOR OUTLETS)

HOT ENGINE BLEED AIR

COOLED BLEED AIR

ENVIRONMENTAL BLEED AIR

RECIRCULATED CABIN AIR (AIR-CONDITIONED WHEN EVAPORATOR IS ON)

AMBIENT AIR

PRESSURE VESSEL

MUFFLER

AIRINLET

DUCTSENSOR

1

AMBIENT AIR

DOOR

AFT FLOOR OUTLETSFLOOR DUCT

AIR CONDITIONED AIR FROM AFT EVAPORATOR

CEILINGOUTLETS

SIDEVIEWFWD1






Pneumatic Systems

Air Distribution System S/Ns BB-1439, 1444 and Subsequent; BT-35 and Subsequent

FIREWALL AMBIENT AIR

FRESH AIR/ RAM AIR

DOOR

FORWARD VENT

BLOWER FORWARD EVAPORATOR

AIRDUCTTEMP SENSE ELEMENT

FORWARD OVERHEAT SENSOR

FORWARD ELECTHEATER

RIGHT BLEED AIR SHUTOFF VALVE (N.C.)

DUCTOVERTEMP SENSOR

CABIN TEMP CONTROL

CHECKVALVES

CABIN HEAT CONTROL VALVE

R BLEED AIR FLOW CONTROL VALVE

MODULATING

AFT EVAPORATOR

AFTELECTHEATER

PRESSURE BULKHEAD AMBIENT AIR SHUTOFF VALVE

LEFT BLEED AIR FLOW CONTROL VALVE FIREWALL

AIR INLET SCOOP

THERMISTER LEFT ENGINE P3 AMBIENT AIR PNEUMATIC BLEED

BLEED AIR SHUTOFF VALVE (N.C.)


CABIN HEAT CONTROL VALVE 30° POSITION SWITCH

ELECTRONIC MODULE BOX ELEC HEAT CONT AIR CONDITIONING CONT

AFT VENT BLOWER

AFTOVERHEAT SENSOR

PRESSUREBULKHEAD

AMBIENT AIR

COOLED AIR

HEATED AIR

PNEUMATIC BLEED AIR SHUTOFF VALVE (N.O.)

AMBIENTAIR

BLEED AIR MODULATING VALVE

AIR SHUTOFF VALVE (N/O)




Pneumatic Systems



Air Conditioning SystemThe air conditioning system provides conditioned air to the cabin and cockpit. The system uses engine bleed air for heating and pressurization; the right engine drives an automotive-type Freon system for cabin cooling. During unpressurized flight, a ram air scoop provides fresh air ventilation.

DistributionAt maximum takeoff power, bleed air directed from the engine compressor section arrives at the environmental Flow Control Unit (FCU) at approximately 343°C (650°F) and 120 PSI. The FCU directs the regulated and mixed bleed air to the cabin-heat control valve, which determines the amount of air that passes through the air-to-air heat exchanger. As the valve closes, more of the air mass passes through the exchanger to decrease the temperature of the air directed to the cabin. The environmental controls in the cockpit position the valve.

Figure 5J-12: Cabin-Heat Control Valve



Air Conditioning System

47 PSI PRESSURE

SWITCH INPUT

COMPRESSOR UNDER PRESSURE/ OVER PRESSURE

SWITCH

N1 SENSOR SWITCH

60% N1, FOR 200 62% N1, FOR B200

10 SEC TIME DELAY

FROM DEACTIVATION

TO REACTIVATION

AUTOMATIC TEMP

CONTROLLER

CABIN HEAT CONTROL

VALVE POSITION SWITCH

CONDENSER BLOWER

CONDENSER

50°F OAT/47 PSI PRESSURE SWITCH INPUT

VENT BLOWER

IN

OUT

RECEIVER DRYER

BY-PASS VALVE

EXPANSION VALVE

FWD EVAPORATOR

COMPRESSOR CLUTCH

UNDER PRESSURE

SWITCH

OVER PRESSURE SWITCH

OUT

IN

AFT EVAPORATOR

AFT VENT BLOWER

EXPANSION VALVE

BYPASS VALVE SENSE SWITCH (33°F)

SUCTION – GASEOUS FREON

PRESSURE – GASEOUS FREON

PRESSURE – LIQUID FREON

COMPRESSOR

EXPANSION VALVE TEMPERATURE BULB

EXPANSION VALVE TEMPERATURE BULB

FWD BULKHEAD

7.5A

NO.

1

D U A L

F E D

B U S

Pneumatic Systems



Environmental Control PanelThe ENVIRONMENTAL control panel on the copilot subpanel provides both automatic or manual control. The panel contains: BLEED AIR VALVES switches (see Pneumatic section, this chapter) CABIN TEMP MODE selector CABIN TEMP control selector MANUAL TEMP switch forward VENT BLOWER switch AFT BLOWER switch.

Figure 5J-13: ENVIRONMENTAL Control Panel

CABIN TEMP ConarolsThe CABIN TEMP MODE selector on the copilot subpanel controls the mode of operation: OFF/AUTO/MAN HEAT/MAN COLD. The selector connects to a control box through a balanced bridge circuit; the control box is the PC board controller of the automatic temperature system.In AUTO, the heating and air conditioning systems operate automatically. The CABIN TEMP knob modulates the cabin heat control valves to maintain the proper temperature of the incoming bleed air.A temperature-sensing unit in the cabin along with the requested setting initiates a heat or cool command to the temperature controller. A duct anticipator temperature probe (duct stat) allows the system to anticipate changes in temperature of inlet air to provide for constant temperature control. The temperature probe and controller are in the ceiling aft of the forward bulkhead.



Selecting a warmer cabin (toward INCR) signals the automatic temperature control to modulate the cabin heat control valves one at a time and allow heated air to bypass the heat exchangers in the wing center sections. The warm bleed air enters the cabin to mix with recirculated cabin air in the floor ducting under the copilot floor.Selecting a cooler mode signals the environmental system to move from a heating mode to a cooling mode. The cabin heat control valves move towards the cool position and pass bleed air through the air-to-air heat exchanger. If this position does not satisfy the thermostat, then the air conditioner comes on. When the controller modulates the cold air, it opens the LH heat valve. When the valve reaches 30° open, the air conditioning goes off. As the temperature warms, the valve cycles closed and the air conditioning resumes.In the MAN HEAT or MAN COOL position, the temperature control system permits manual control of the cabin temperature with the MANUAL TEMP switch.Momentarily holding the MANUAL TEMP spring-loaded center return switch to INCR (hot) or DECR (cold) results in modulation of the cabin-heat control valves in the bleed air lines. Only one valve moves at a time. Allow approximately 30 seconds per valve (one minute total) for the valves to move to the full heat or full cold travel.Movement of the valves varies the amount of bleed air routed through the heat exchanger to control the temperature of the incoming bleed air. The bleed air mixes with recirculated cabin air (air conditioned air with the refrigeration system operating) in the mixing plenum. Cool air comes out of the overhead vents while the majority of warm air comes out of the floor vents. MAN COOL directs the air conditioner system to operate if the right engine speed is above 60% (200) to 62% (B200) N1.

Vhna Blowhr ConarolThe VENT BLOWER/HI/LO/AUTO switch controls the forward vent blower. With the CABIN TEMP MODE selector switch in AUTO, the blower operates at low speed but can be switched to HI. Selecting the VENT BLOWER switch to LO or HIGH operates the blower at that speed regardless of the CABIN TEMP MODE selection.If the optional aft evaporator unit is installed in the air conditioning system, an aft blower is installed under the floor to draw cabin air across the aft evaporator and to the aft floor and ceiling outlets. The AFT BLOWER switch controls the blower (high speed only).Do not use the aft blower during heating mode. The door between the aft blower duct and floor outlet duct opens when the aft blower is in operation; this stops bleed air flow to the aft floor registers and delivers recirculated air. Some aircraft serial numbers prior to BB-39 have a two-speed aft blower controlled by the VENT BLOWER switch. Operation was entirely automatic.

Pneumatic Systems



CoSkpta Coefora ConarolsFour manual controls on the main instrument subpanels regulate cockpit comfort with the cockpit partition door closed and the cabin comfort level satisfactory. The controls are: PILOT AIR/COPILOT AIR knobs DEFROST AIR knob CABIN AIR knob.

Figure 5J-14: PILOT AIR/COPILOT

AIR Knobs

Figure 5J-16: CABIN AIR Knob

Pulling the controls full out provides maximum heating to the cockpit. Pushing the controls full in provides maximum heating to the cabin.The CABIN AIR knob controls the cabin air valve. Pulling the knob out allows only a minimum amount of air to pass through the valve; this increases the amount of air available to the crew outlets.

Figure 5J-15: DEFROST AIR Knob



The PILOT AIR and COPILOT AIR knobs controls air flow volume through a mechanically controlled damper in each outlet. The DEFROST AIR knob controls a valve for defrosting at the forward side of the pilot and copilot heat duct. It also controls air flow to the eyeball outlets through the air plenum in the glareshield.

Radiant Heating SystemThe optional radiant heat system preheats the cabin prior to engine start or supplies a supplemental heat source during flight. On S/N 002 to 414, the system consisted of two heating panels bonded to the headliner with an overheat protection provided by a thermal switch in each panel.On S/N 415 and subsequent, the aircraft use five heating panels above the windows with overheat protection from a thermostat and 90°C (194°F )thermal fuse on the back of each panel.The RADIANT HEAT/OFF switch on the copilot subpanel controls the operation of the panels.A radiant heat element in the cargo door is controlled by the CABIN TEMP MODE switch; it operates in all heating modes. This unit provides supplemental heat to the cabin for additional passenger comfort.

Electric Heating SystemOn S/N BB-1139, BB-1444 and subsequent, a supplemental electric heating system is operated by the ELEC HEAT-OFF switch in the ENVIRONMENTAL group on the copilot’s left subpanel. It should be used in conjunction with the manual heat or auto temp control mode only. This system can be used in conjunction with an auxiliary power unit for cabin heating prior to starting the engines, and it can provide supplemental heating for ground operation only.This system uses one forward heating element located in a forward duct and one aft heating element located in the aft evaporator plenum. Both the forward and the aft blower must be operating during electric heat operation and for 15-30 seconds minimum after switch is turned OFF.An ELEC HEAT ON advisory annunciator indicates that the power relays are in the closed position for applications of electrical power to the heating elements. When the electric heat system is selected to OFF, the ELEC HEAT ON annunciator must be extinguished to indicate power is removed from the heating elements before the blowers are switched OFF.

NNTEc: The electric heat system will draw approximately 300 amps.

Pneumatic Systems



Freon Air Conditioning SystemThe air conditioning system uses Freon (refrigerant 12) to cool the cabin and cockpit on the ground and in the air. With the Freon system operating, recirculation occurs as the blower fan draws recirculated warm cabin air across the evaporator to cool it and then delivers the cooler air back to the cabin and cockpit distribution system.If the air conditioning system is not operated for an extended period of time, moisture may condense and settle in the system low spots, resulting in corrosion of the refrigerant lines. Additionally, system seals may dry, shrink and crack, due to inadequate lubrication. To protect the integrity of the system, the air conditioner should be operated at least 10 minutes every month. Since the air conditioner can not be operated at ambient temperatures below 10°C (50°F), the recommended monthly interval can be extended in conditions where minimum ambient temperatures cannot be attained.

NNTEc: SI-0968-II; Air Conditioning – Installation of High Pressure Plumbing and a New High Pressure Switch (S/N 002 to 344 except 001, 003 to 005, 034, 123, 186, 203, and 270).

CoeprhssorThe Freon compressor is an automotive-type compressor mounted on and belt-driven by the right engine. Plumbing routes Freon through the right wing inboard leading edge to the fuselage and then forward to the condenser coil, receiver dryer, expansion valve, bypass valve, and evaporator before returning to the compressor.

Figure 5J-17: Freon Compressor



CondhnshrAmbient air across the condenser cools and condenses high-pressure and high temperature gaseous Freon into a liquid. On the ground, the condenser blower draws air across the condenser. During flight, ram air enters the ram air scoop and passes through the condenser to cool the gas. The ram air exits through the outlet air vent. The liquid Freon then flows to the receiver-dryer.On S/N 345 and subsequent and aircraft with SI-0968-II, the condenser blower shuts off when nose gear drag brace unlocks on gear retraction.A receiver-dryer is in the upper portion of the condenser compartment. The reservoir incorporates a sight glass that allows maintenance to check for air in the system with the Freon system operating. The flow of bubbles across the sight glass indicate low Freon charge or air in the system.The reservoir contains a desiccant that absorbs traces of water. This absorption prevents moisture damage to the system.

EvaporatorThe Freon enters the evaporator through an expansion valve where it rapidly expands into a cold vapor because of the low pressure level established by compressor suction. The gaseous Freon in the evaporator absorbs heat from the cabin air as it passes through the evaporator core. A forward vent blower bows recirculated cabin air across the forward evaporator. The cooler air passes into the mixing plenum, and into the ceiling outlet ducts.Between the expansion valve and the evaporator coil is a 0.5°C (33°F) thermal sensor switch. This switch activates a hot gas bypass valve to mix hot gas with the cold Freon entering the evaporator; this prevents ice forming on the cooling coils.

Lteta wtaSchsThe high and low pressure limit switches and the N1 speed switch prevent damage to the compressor from too low or too high pressure operation. High and low pressure switches attach to the refrigerant lines under the right center section leading edge. On early model 200s, activation of the pressure switches opens an adjacent 7.5A fuse; on later model 200s, a reset switch in the nosewheel well deactivates the system. Both systems shut the compressor and condenser blower down.An additional limiter is the 47 PSI (50°F) switch downstream of the condenser in the nose. Condenser pressure below 47 PSI removes power from the compressor clutch and condenser blower. Pressures above 52 PSI (55°F) reactivate the clutch and blower.The N1 engine speed switch removes electrical power from the compressor clutch if the right engine speed drops below 60% (200) to 62% (B200) N1. In addition, the switch illuminates the green AIR CND N1 LOW annunciator to signal current interruption to the compressor clutch.

Pneumatic Systems



Npatonml Evmpormaor UntaAn aft evaporator and blower may be installed in the fuselage center aisle equipment bay to provide additional cooling. The unit works when the aft blower sends recirculated cabin air through the aft evaporator and delivers it to the aft floor and ceiling outlets.




Pneumatic Systems




PreflightInspect the pneumatic, air conditioning, and pressurization systems in accordance with the Preflight chapter of this manual. Accomplish normal operation of these systems in accordance with the Expanded Normals and Maneuvers chapters of this manual.

Emergency ProceduresThis section discusses what happens within the pneumatic, pressurization, and air conditioning system during emergency situations. For a list of specific procedural steps, please refer to your CAE Operating Handbook.

Cmbtn Prhssurh Too Htgc (Nvhrprhssurtzmaton)Monitor cabin pressure and cabin vertical speed indicators. When too high cabin pressure is indicated, select a higher altitude on the pressurization controller to increase the cabin altitude reference for the pressurization system. This lowers the cabin pressure. If this action is successful, continue the flight with the higher altitude.If the cabin pressure does not decrease, don oxygen masks as required. Select ENVIR OFF to close the environmental FCU and open the pneumatic instrument air valves. After the cabin depressurizes, select the CABIN PRESS switch to DUMP to open the safety valve.Descend to the required altitude to allow for safe operation and move BLEED AIR VALVES to OPEN to open the environmental FCU and pneumatic instrument air valve.

Undhr PrhssurtzmatonShould the aircraft gradually loose pressurization or fail to pressurize after liftoff, verify that the cabin pressurization switch is in the PRESS position, the bleed air valves are OPEN, and the controller is properly set for the particular phase of flight. To regain pressurization control, move the pressurization control switch to the TEST position, thereby removing any electrical power from the control solenoids. If pressurization returns, allow approximately 1.0 to 1.5 PSI differential to build before pulling the PRESS CONT Circuit Breaker (CB). Release the test switch.The controller must now be properly adjusted to ensure the outflow valve will open prior to touchdown. There will be no safety relief available until the Control CB is reset. The Control CB should be reset prior to touchdown to guarantee a depressurized cabin upon landing. Failure to reset the CB could result in landing pressurized, preventing safe cabin egress.



Loss of PrhssurtzmatonDuring a loss of pressurization at high altitude use oxygen and descend as required. See Table 5J-1 for information concerning useful consciousness time without supplemental oxygen. The ALT WARN annunciator illuminates indicating cabin is 12,500 ft. Confirm that passenger masks have been deployed.

NNTEc: Cabin lighting goes to full bright when passenger oxygen masks are deployed.

DUCT NVERTEMP AnnunStmaor IlluetnmahdIllumination of the DUCT OVERTEMP annunciator indicates an air temperature of 149°C (300°F) inside the duct. Select a lower temperature on the CABIN TEMP knob and adjust the vent blower to HI. Push the CABIN AIR knob in.If the annunciator remains illuminated, select the CABIN TEMP MODE selector to MAN HEAT and the MANUAL TEMP switch to decrease.

BL AIR FAIL AnnunStmaor IlluetnmahdIllumination of the red L or R BL AIR FAIL annunciator indicates the polyflow bleed air failure warning line has ruptured (failed); this indicates a possible bleed air leak. Check engine gauges for secondary indications.Until the repair of the polyflow line, the respective L or R BL AIR FAIL annunciator remains illuminated.

CrmSk tn Any tdh Wtndow or tn Wtndscthld

CAAUTIO Prior to next flight, maintenance actions on a cracked window or windshield are required. Refer to the Airworthiness Limitations in Chapter 4 of the Super King Air 200 Series Maintenance Manual.

If it has been determined that a crack has developed in any side window or windshield, maintain altitude at 25,000 ft or less and reset the pressurization controller to maintain 4.0 PSI or less as required.

NNTEc: Visibility through a cracked windshield may be significantly impaired.

NNTEc: Heating elements may be inoperative in the area of a crack.

Pneumatic Systems



eokh mnd Fueh EltetnmatonIdentify the source of the smoke or fumes. Gray or tan smoke along with irritation of the nose and eyes identifies an electrical failure. An environmental system failure produces a white smoke and is much less irritating to the nose and eyes.Don oxygen masks and select communications before attempting to troubleshoot the source.

NNTEc: Opening a storm window after depressurization facilitates smoke and fume removal from the aircraft.

Rmptd DhSoeprhsston/EehrghnSy DhsShnaDon crew oxygen masks, select communications, and execute a maximum performance descent. Confirm passenger oxygen masks deployed, turn ignition on to prevent engine flameouts, advise ATC, and set the transponder to emergency squawk. The aircraft descent must be rapid to minimize the risk of losing consciousness at high altitude.

Altitude Time35,000 ft 1/2 to 1 minute 30,000 ft 1 to 2 minutes 28,000 ft 21/2 to 3 minutes 25,000 ft 3 to 5 minutes 22,000 ft 5 to 10 minutes

12,000 ft to 18,000 ft 30 minutes

Table 5J-1: Useful Consciousness Time without Supplemental Oxygen




Pneumatic Systems



Data Summaries

Bleed Air System

Power Source Bleed air (each engine – station P3) Distribution Prior to 18 PSI regulator:

Brake deice Rudder boost ÄP switch After 18 PSI regulator Bleed air warning system Rudder boost servos Flight hour meter Door seal (if installed) Vacuum Deice boots

Control BLEED AIR VALVE switches Monitor PNEUMATIC PRESSURE gauge GYRO

SUCTION gauge Protection BL AIR FAIL annunciators Bleed air

shutoff valves Relief valves Check valves Power Source Bleed air (each engine – station P3) Distribution Prior to 18 PSI regulator: Brake deice

Rudder boost AP switch After 18 PSI regulator Bleed air warning system Rudder boost servos Flight hour meter Door seal (if installed) Vacuum Deice boots

Control BLEED AIR VALVE switches Monitor PNEUMATIC PRESSURE gauge GYRO

SUCTION gauge Protection BL AIR FAIL annunciators Bleed air

shutoff valves Relief valves Check valves



Air Conditioning/Heating System

Power Source Engine-bleed air – heating Right engine – Freon system

Distribution Engine compressor bleed air Environmental control unit Cabin Cockpit

Control Switches BLEED AIR VALVES MANUAL TEMP VENT BLOWER AFT BLOWER RADIANT HEAT (prior to BB-1439) ELEC HEAT (BB-1439 and subsequent) CABIN TEMP MODE selector CABIN TEMP control selectorRight engine RPM above 60%

Monitor CABIN AIR gauge Thermostat AIR CND N1 LOW annunciator

Protection High and low pressure switches N1 speed switch 47 PSI pressure switch

Pressurization System

Power Source Bleed air (each engine – station P3) Distribution Flow control unit

Air-to-air heat exchangers Cockpit Cabin

Control Switches BLEED AIR VALVES CABIN PRESSPressurization controller

Monitor Cabin altitude and differential pressure gauge Cabin VSI

Protection BL AIR FAIL annunciators Bleed air shutoff valves CABIN ALT annunciator Squat switch Outflow/safety valves (negative/maximum differential relief)Passenger oxygen mask auto deployment system

Powerplant 5K

King Air 200 5K-1December 2011


ContentsPowerplant

SchematSc: PT6A Engine ............................................................5K-7Turboprop Engine

Components ...........................................................................................5K-9Air Inlet ............................................................................................. 5K-10Inertial Separation System ............................................................... 5K-10Compressor Inlet Case .................................................................... 5K-12Gas Generator Case and Compressor Section ............................... 5K-12Combustion Section ......................................................................... 5K-13Compressor Turbine and Guide Vanes ............................................ 5K-13Accessory Gearbox ......................................................................... 5K-14Power Section .................................................................................. 5K-14Exhaust Duct ................................................................................... 5K-15Reduction Gearbox and Propeller Shaft .......................................... 5K-15

Engine Indicating ................................................................................. 5K-16ITT Indicator ..................................................................................... 5K-16Torquemeter .................................................................................... 5K-17Gas Generator Tachometer (N1) ...................................................... 5K-17

Operating Limitations ......................................................................... 5K-19SchematSc: Lubrication System ................................................ 5K-23

Powerplant SystemsLubrication ........................................................................................... 5K-25

SchematSc: Fuel Control System .............................................. 5K-26Oil Tank ............................................................................................ 5K-27Centrifugal Breather ......................................................................... 5K-27Chip Detector ................................................................................... 5K-27Pressure Pump ................................................................................ 5K-27

SchematSc: Ignition System ...................................................... 5K-28Oil Pressure Relief/Pressurizing Valve ............................................ 5K-29Oil Filter ........................................................................................... 5K-29Oil Cooler ......................................................................................... 5K-29Fuel Heater ...................................................................................... 5K-29Scavenge Pumps ............................................................................. 5K-29

King Air 2005K-2December 2011


Oil Pressure Gauge ..........................................................................5K-30Oil Temperature Gauge ....................................................................5K-30Distribution ........................................................................................5K-30

Engine Fuel and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-31Boost Pump ......................................................................................5K-31Oil-to-Fuel Heater .............................................................................5K-31Fuel Pump ........................................................................................5K-31Fuel Control Unit ...............................................................................5K-32Pneumatic Section ............................................................................5K-33Governing Section ............................................................................5K-33Torque Limiter ...................................................................................5K-34Flow Divider and Dump Valve ..........................................................5K-34Fuel Manifold and Nozzles ...............................................................5K-34Fuel Drain Valves .............................................................................5K-35Fuel Flow Gauge ..............................................................................5K-35FUEL PRESS Annunciators ..............................................................5K-35Ignition ..............................................................................................5K-36Ignition Exciter ..................................................................................5K-36Igniter Plugs ......................................................................................5K-36Ignition Switches ...............................................................................5K-37Auto-Ignition System .........................................................................5K-37

Engine Power Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-38Power Lever .....................................................................................5K-38Condition Lever .................................................................................5K-38Bleed Air ...........................................................................................5K-39Compressor Bleed Valves ................................................................5K-39Cooling and Sealing ..........................................................................5K-39Airframe Bleed Air .............................................................................5K-40

Propeller SystemsSchematSc: Propeller Systems ................................................. 5K-42SchematSc: Autofeather System ............................................... 5K-43SchematSc: Type I and Type II Sychrophaser ........................... 5K-44

Propeller De-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-45Low Pitch Stop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-45Primary Governor and Beta Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-46Overspeed Governor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-47Autofeather System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-48Propeller Synchrophaser . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-49

Type I Synchrophaser ......................................................................5K-49Type II Synchrophaser ......................................................................5K-49

Powerplant



Propeller Synchroscope . . . . . . . . . . . . . . . . . . . . . . . . . . . .5K-50Prop Tachometer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5K-51Power Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5K-52Propeller Control Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . .5K-52

Preflight and ProceduresPreflight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5K-53Emergency Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . .5K-53

Emergency Engine Shutdown . . . . . . . . . . . . . . . . . . . . . . 5K-53Engine Fire On Ground . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-54Engine Failure During Ground Roll . . . . . . . . . . . . . . . . . . 5K-54Engine Failure After Lift-Off . . . . . . . . . . . . . . . . . . . . . . . . 5K-54Failure in Flight Below Minimum VMCA . . . . . . . . . . . . . . . . 5K-55Engine Flameout (Second Engine) . . . . . . . . . . . . . . . . . . 5K-55Low Oil Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-55Airstart/Starter Assist . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-55Windmilling Airstart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5K-56Propeller Overspeed (Up to 2,080 RPM) . . . . . . . . . . . . . . 5K-56Propeller Overspeed (Above 2,080 RPM) . . . . . . . . . . . . . 5K-56

Data SummaryPowerplant System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5K-57




Powerplant



PowerplantThe Powerplant chapter includes information on several areas: the turboprop engine (components, instrumentation and operation) engine oil and lubrication ignition engine fuel and fuel control power control propeller (components, instrumentation and operation).

Two Pratt & Whitney Canada PT6A-41 (for 200 Series) or PT6A-42 engines (for B200 Series) power the Super King Air. The PT6A is a lightweight, reverse flow, free turbine engine that drives a three or four-bladed constant speed, full-feathering reversible propeller.Each engine produces 850 shaft-horsepower, 2,230 foot-pounds of torque, and 134 to 135 lbs of jet thrust (through engine exhaust). The PT6A-41 and PT6A-42 engines are essentially the same; the PT6A-42 differs in that it has an improved compressor section that provides a 10% increase in performance over the PT6A-41.





5K-7For Training Purposes Only

Powerplant

PT6A Engine

PROPFLANGE

PROPGOVERNORPAD

REDUCTIONGEARS

POWERTURBINES

FUELNOZZLE

IGNITOR

INTERSTAGETURBINETEMPERATUREPROBE

COMPRESSORTURBINE

CENTRIFUGALCOMPRESSOR

HIGH PRESSUREBLEED VALVE

AIR INLETSCREEN

ACCESSORY SECTION HIGH PRESSURE FUEL PUMP BOOST PUMP OIL PRESS PUMP N1 GOVERNOR A/C COMPRESSOR (RIGHT ENG) OIL SCAVENGE PUMPS STARTER GENERATOR N1 TACH GENERATOR

CHIP DETECTOR EXHAUST OUTLET

FUELNOZZLE

IGNITOR

ANNULARCOMBUSTIONCHAMBER

3 STAGEAXIALCOMPRESSORLOW PRESSURE

BLEED VALVE

COMPRESSORINLET

P BLEED AIR TAPOFF FOR: ENVIRONMENTAL SYSTEM PRESSURIZATION PNEUMATICS

3

TRANSFERVALVE

OIL RESEVOIR


5K-8 For Training Purposes Only


Powerplant



Turboprop EngineThe free-turbine turboprop engine compresses air, mixes it with fuel, and ignites the mixture to produce a hight emperature, high-speed gas that drives a power turbine connected to a reduction gearbox. The reduction gearbox converts high-speed, low torque energy from the power turbine into low-speed, high-torque shaft horsepower for the propeller.The combustion cycle begins with the induction of air through an annular (ring-shaped) plenum chamber formed by the compressor inlet case. The air directed forward flows through the compressor where each successive compressor stage (stator and rotor pair) converts air velocity into increasing air pressure. After exiting the compressor section, vanes straighten the airflow before it reaches the combustion section.As the high pressure air enters the annular combustion chamber, it changes direction 180° before it mixes with fuel. A circular arrangement of 14 simplex atomizers introduce a fine spray of fuel into the combustion chamber where the air and fuel mix. Two igniters protruding into the combustion chamber spark to ignite the fuel/air mixture. Once the engine reaches operating speed, the igniters are no longer required as the combustion process is self-sustaining.The rapidly expanding, high-temperature gases flow from the combustion chamber and change direction 180° as they travel through the exit zone. Inlet guide vanes straighten the gas flow before it reaches the single-stage compressor turbine. The flow of high temperature, high-speed gases drives the compressor turbine; the turbine, in turn, drives the compressor through a shaft at the rear of the engine.After passing through the compressor turbine, the gas flow drives the two stage power turbine connected to a reduction gearbox. The gearbox-drives the propeller shaft. Finally, the exhaust gases exit the engine through the exhaust duct and stacks.

ComponentsTo assist in the understanding of the turboprop engine, this discussion examines the components of the engine in the same manner as the engine produces power: from air intake to propeller shaft rotation. The PT6A is a modular engine that consists of a gas generator section and a power section.



Atr InlhaAir enters through the air inlet at the bottom of the engine nacelle. For anti-icing the inlet uses engine exhaust diverted by a scupp carries hot exhaust air through a tube that encircles the engine air inlet. An opening at the bottom of the nacelle exhausts the hot air overboard. Later model aircraft do not have the exhaust on the lower nacelle; a scupper on one stack supplies hot exhaust to the tube and exhausts it through the opposite stack.

Figure 5K-1: Air Inlet

Inhratml hpmrmaton ysaheAn inertial separator system in the engine air inlet duct prevents the ingestion of moisture and ice particles into the engine (see Ice and Rain Protection chapter). When extended, an electrically or mechanically operated ice vane creates a sharp turn in the air flow toward the engine. The inertial force of the heavier moisture and ice particles carries them past the sharp turn required to enter the engine inlet and out through the bypass door in the lower cowling.

Powerplant



The ice vane and bypass door are simultaneously electrically-extended through a chain and sprocket drive controlled by the ICE VANE switches on the pilot subpanel. The system employs a timer circuit and microswitches that illuminate a yellow ICE VANE annunciator if the doors fail to reach the fully-extended position in 15 seconds. If the electrical system fails there are mechanical ICE VANE EMERGENCY MANUAL EXTENSION T-handles below the left subpanel. After pulling the system CBs, pulling the handles extends the ice vane and bypass doors; pushing the handles in retracts the vanes and doors. There is no manual backup system on later aircraft.

Figure 5K-4: Ice Vane Emergency Manual Extension T-Handles

Figure 5K-2: Ice Vane Switches Figure 5K-3: Anti-Ice System Control Panel



Coeprhssor Inlha CmshAir flows from the engine inlet duct through a wire mesh screen into the annular (ring-shaped) plenum chamber. The mesh screen prevents engine ingestion of foreign objects. The front section of the compressor inlet case forms the plenum chamber and the aft section forms the engine oil tank where the accessory gearbox bolts to the engine. The compressor inlet case also contains the No.1 bearing, bearing support housing and compressor rear stator air seal.The air inlet directs air forward (reverse flow) to the compressor section.

Figure 5K-5: Compressor Inlet Case

Gms Ghnhrmaor Cmsh mnd Coeprhssor hSatonThe gas generator case, between the compressor inlet case and the exhaust duct, houses the compressor section, No.2 bearing, and the combustion section. The front section of the case forms the outer housing for the combustion chamber lining. Two compressor bleed valves (see Bleed Air) on the conical outer section dump excess interstage compressor air (P2.5) overboard at lower power settings.The compressor section consists of a three-stage axial compressor and a single- stage centrifugal impeller.Each axial compressor stage consists of airfoil-shaped rotor blades secured to a disk followed by a stator assembly. The stator assembly consists of airfoil-shaped vanes arranged radially within the gas generator case. The stator vanes straighten the airflow exiting the axial compressor before it reaches the next stage.The first-stage rotor has 18 (PT6A-41) or 16 (PT6A-42) titanium rotor blades; the second and third stages each have 36 cadmium-plated stainless steel blades. Each blade has dovetail roots that loosely fit into grooves in the rotor disks. Between each rotor disk there are interstage spacers. Together, the disks, spacers, and centrifugal impeller bolt together with six tie-rods. The compressor stubshaft and the No.2 bearing support the front of the compressor assembly. The aft end is supported by the No.1 bearing and hub.

Powerplant



Once the airflow enters the compressor section, the rapidly rotating axial rotors transfer energy to the airflow. The stator blades straighten and diffuse the airflow to partially convert high velocity airflow into pressure.After exiting the three axial compressor stages, the centrifugal impeller picks up and accelerates the airflow outward toward the diffuser tubes. The diffuser tubes, brazed to the inside of the gas generator case, increase the static pressure of the compressor discharge air (P3) and direct it through straightening vanes before it reaches the combustion section.The engine uses approximately 25% of the bleed air developed by the compressor section for combustion. It uses the remaining 75% for: cabin pressurization bleed valve operation fuel control unit regulation turbine, combustion chamber, and nozzle cooling bearing sealing.

Coebusaton hSatonAfter exiting the diffuser tubes and straightening vanes, the airflow enters the combustion section. The combustion section, contained within the front section of the gas generator case, consists of an annular stainless steel liner and the small and large exit ducts. Fourteen fuel nozzles arranged radially around the liner introduce an atomized spray of fuel into the combustion liner. Two igniters extending into the liner provide an ignition source for engine starting. The fuel manifold nozzle adapters and the two igniter plugs secure the liner in the gas generator case.Holes in the inner and outer walls of the liner direct compressor discharge air into the liner combustion zone where it mixes with fuel. The igniters provide the initial spark for combustion. The fuel/air mixture ignites and the hot expanding gases flow to the rear and into the annular passage between the small and large exit ducts. The passage forces the air through a 180° turn before it reaches the compressor turbine inlet guide vanes.

Coeprhssor Turbtnh mnd Gutdh VmnhsThe fourteen inlet guide vanes, arranged radially within the compressor turbine guide vane ring, guide the expanding gases from the combustion liner toward the compressor turbine. The compressor turbine, rotating within the compressor turbine shroud assembly, consists of a balanced turbine disk with 53 turbine blades and an integral shaft extension. The compressor turbine shroud assembly incorporates shroud segments that act like a seal and provide running clearance for the turbine bladesThe compressor turbine converts energy (velocity) from the rapidly expanding gases exiting the combustion section into mechanical energy. The rapidly rotating compressor turbine then drives the compressor section and accessory gearbox.



ASShssory GhmrboxThe accessory gearbox, driven by the compressor shaft through a coupling shaft, uses a system of gears to reduce shaft speed and power accessory drive pads. Drive pads on the rear of the accessory gearbox drive the starter/generator, pressure oil pump, scavenge oil pumps, pressure fuel pump and fuel control unit, fuel boost pump, and N1 tachometer/generator.Additionally, the right engine accessory gearbox drives the air conditioning system compressor.

Powhr hSatonAfter passing through the compressor turbine the exhaust gases flow into the power section. The power section consists of the power turbine stator housing, power turbine first stage vane ring, interstage baffle, and two power turbine disks.The power turbine stator housing is a conical casting with open ends. The housing accommodates the power turbine stator and shroud assemblies. Eight or 10 interstage turbine temperature (T5) thermocouples mount in bosses on the rear end of the housing. The first-stage power turbine stator assembly consists of 20 vanes mounted to an inner and outer ring. It directs combustion gases to the first-stage power turbine blades. The first-stage power turbine has 47 blades on a turbine disk. The second-stage stator vane ring has 41 vanes mounted between an inner and outer ring. It directs the gas flow exiting the first-stage turbine. The second-stage power turbine consists of 43 blades on a turbine disk. Both the first and second-stage turbines mount to the power turbine shaft that is supported by the No.3 and No.4 bearings. Together, the turbine and stator assemblies convert the high speed combustion gases into mechanical energy to rotate the power turbine shaft.

Powerplant



Excmusa DuSaAfter passing through the power section, the combustion discharge gases enter the exhaust duct. The exhaust duct directs the exhaust gases out of the engine through an exhaust stack on either side of the engine. The exhaust duct also supports the reduction gearbox.

Figure 5K-6: Exhaust Duct

RhduSaton Ghmrbox mnd Prophllhr cmfaThe PT6A uses a Reduction Gearbox (RGB) to convert the high-speed, low torque energy from the power turbines into low-speed, high-torque energy to drive the propeller. The reduction gearbox reduces power turbine speed of approximately 33,000 RPM to 2,000 RPM at the propeller shaft (15:1).The reduction gearbox consists of a front and rear case that contains a two stage planetary gear train, accessory drives, torquemeter, and a propeller shaft. As the power turbine shaft and turbine coupling rotate they transmit torque through a sun gear and three planet gears to a stationary ring gear; this action turns the planet gear carrier. Rotation of the first-stage planet gear carrier, in turn, drives the second-stage sun gear through a flexible coup-ling. Rotation of the second-stage sun gear drives five planet gears in the second stage planet gear carrier. The second stage planet gear carrier drives the propeller shaft..A bevel gear at the rear of the propeller drive shaft bearing provides power through a drive shaft to three accessory pads on the front case of the reduction gearbox. The drive pads power the propeller governor, propeller tachometer generator, and the propeller overspeed governor.



Engine IndicatingEach engine has its own set of engine instruments on the left side of the center instrument panel. These include: Interstage Turbine Temperature (ITT) Torquemeter Gas generator speed (N1) indicator.

The fuel flow, oil pressure and temperature, and propeller speed indicators are found in the discussions concerning their systems.

ITT IndtSmaorThe ITT indicator uses eight or 10 alumel/chromel thermocouple probes to provide an indication of interstage temperature (T5). Each probe protrudes in front of the first-stage turbine guide vanes into the gas stream. The probes, arranged radially around the power turbine stator assembly, connect in parallel through junction boxes to create an averaging circuit. A wiring harness from a central junction box carries probe signals from the system to the ITT indicator in the cockpit.

Figure 5K-7: ITT Indicator

The ITT indicator displays temperature in degrees centigrade (°C) from 200 to 1,200°C. A green arc, a red radial line, and a dashed red radial line indicate the normal and maximum continuous operating and maximum starting temperatures for the engine. The ITT indicating system does not operate on aircraft power; it is self powered.On aircraft with PT6A-41 engines, the green arc covers 400 to 750°C and the red radial line is at 750°C. On aircraft with PT6A-42 engines, the green arc covers 400 to 800°C and the red radial line is at 800°C. The ITT indicators for both engines have a red dashed radial line at 1,000°C to indicate the maximum starting temperature.

Torque Limitations

Takeoff . . . . . 2,230 FT-LBSMax Continuous. . . . 2,230 FT-LBSCruise Climb 2,230 FT-LBS

Powerplant



TorquhehahrA hydro-mechanical torquemeter accurately measures the torque applied to the propeller shaft. The torquemeter, in the first stage of the reduction gearbox, uses the rotational force of the first stage ring gear to compress oil in a chamber. The system then senses the difference between torquemeter pressure and reduction gear internal pressure to derive a measurement of torque being applied. A torquemeter transmitter provides this measurement as a 26 VAC electrical signal (28 VDC on S/Ns BB-1096, BB-1098 and subsequent to the TORQUE indicator on the center instrument panel.

Figure 5K-8: Torquemeter

The indicator, marked from 0 to 25, indicates engine torque from 0 to 2,500 ft-lbs. A green arc and a red radial line indicate the normal and maximum operating ranges. The green arc covers 400 to 2,230 ft-lbs and the red radial line is at 2,230 ft-lbs.On aircraft S/N BB-1439, BB-1444 and subsequent, BN-5 and subsequent, and BL-139 and subsequent, the torque indicator is a single needle with LED readout. The green arc extends to 2,230 ft-lbs, where the red line is displayed.

Gms Ghnhrmaor TmScoehahr (N1)A tachometer indicates the speed of the engine compressor shafts as a percentage of maximum operating speed (100% = 37,500 RPM). The engine accessory gearbox drives a tachometer-generator that provides an electrical signal proportional to engine rotation. The signal drives the gas generator tachometer (N1) in the cockpit.

N1 Limitations

Low Idle (-41). . . . . . . 52%Low Idle (-42). . . . . . . 56%Takeoff . . . . . . . . . .101.5%Max Continuous. . .101.5%Cruise Climb . . . . .101.5%



The indicator uses two pointers moving over separate scales to indicate %RPM. The larger scale, graduated from 0 to 100, indicates to the nearest 10% RPM. The smaller scale, graduated from 0 to 9 indicates RPM to the nearest 1%.A green colored arc on the larger scale from 52 or 56% to 101.5% indicates the normal operating range of the engine. A red radial at 101.5% indicates the maximum operating speed of the engine.

Figure 5K-9: Gas Generator Tachometer

For aircraft S/N BB-1439, BB-1444 and subsequent, BN-5 and subsequentand BL-139 and subsequent, the N1 percent of RPM indicator is a single needle with LED readout. The normal (green arc) is from 62% to 101.5%. The red arc reading is 101.5%.

Figure 5K-10: RPM Indicator

Powerplant



Operating LimitationsTables 5K-1, 5K-2, 5K-3 and 5K-4 include operating limitations for the Pratt and Whitney Canada PT6A-41 and PT6A-42 engines.

Operating Condition

SHP Torque(ft-lbs)1

MaxObservedITT (°C)

N1 RPM N1 % PropRPM

N2

Oil Press(PSI)2

Oil Temp°C

Starting ___ ___ 10003 ___ ___ ___ ___ -40 (min) Low Idle ___ ___ 6604 19,500 52 (min) ___ 60 (Min) -40 to 99 High Idle ___ ___ ___ ___ 5 ___ ___ -40 to 99 Takeoff 9 850 2230 750 38,100 101.5 2000 105 to 135 10 to 99 Max Continuous and Cruise

850 22306 750 38,100 101.5 2000 105 to 135 10 to 99

Cruise Climb and Rec Cruise

850 22306 725 38,100 101.5 2000 105 to 135 0 to 99

Max Reverse7 ___ ___ 750 ___ 88 1900 105 to 135 0 to 99 Transient ___ 27503 850 385,008 102.68 22003 ___ 0 to 1049

Table 5K-1: King Air 200 Engine Operating Limits (PT6A-41)1. Torque limit applies within range of 1,600 to 2,000 propeller RPM (N2). Below 1,600 RPM, torque

is limited to 1,100 ft-bs.2. When gas generator speeds are above 27,000 RPM (72% N1) and oil temperatures are between

60 and 71°C,normal oil pressure are: 100 to 135 PSI below 21,000 ft and 85 to 135 PSI at 21,000 ft and above. During extremely cold starts, oil pressure may reach 200 PSI. Oil pressure between 60 and 85 PSI is undesirable; it should be tolerated only for the completion of the flight, and then only at a reduced power setting not exceeding 1,100 ft-lbs torque, Oil pressure below 60 PSI is unsafe; it requires that either the engine be shut down, or that a landing be made as soon as possible with minimum power to sustain flight. Fluctuations of ±10 PSI are acceptable.

3. These values are time limited to 5 seconds.4. High ITT at ground idle may be corrected by reducing accessory load and/or increasing N1 RPM.5. At approximately 70% N1.6. Cruise torque values vary with altitude and temperature.7. This operation is time limited to one minute.8. These values are time limited to 10 seconds.9. These values are time limited to 5 minutes.



Operating Condition

SHP Torque(ft-lbs)1


N1 RPM N1 % PropRPM

N2

Oil Press(PSI)2

Oil Temp°C3,4

Starting ___ ___ 10005 ___ ___ ___ ___ -40 (min) Low Idle ___ ___ 7506 21000 56 (min) ___ 60 (Min) -40 to 99 High Idle ___ ___ ___ ___ 7 ___ ___ -40 to 99

Takeoff 9 850 2230 800 38,100 101.5 2000 105 to 135 10 to 99 Max Continuous and Cruise

850 22308 800 38,100 101.5 2000 105 to 135 10 to 99


850 22308 770 38,100 101.5 2000 105 to 135 0 to 99

Max Reverse8 ___ ___ 750 ___ 88 1900 105 to 135 0 to 99 Transient ___ 27503 850 38,50010 102.610 22005 2002 0 to 10411

Table 5K-2: King Air 200 Engine Operating Limits (PT6A-41)

S/Ns BB-743, 793, 829, 854 to 870, 874 to 891, 894, 896 to 911,913 to 1438, 1443; BL-37 to 1381. Torque limit applies within range of 1,600 to 2,000 propeller RPM (N2). Below 1,600 RPM, torque is limited to 1,100 ft-bs2. When gas generator speeds are above 27,000 RPM (72% N1) and oil temperatures are between 60 and 71°C, normal oil pressures

are: 100 to 135 PSI below 21,000 ft and 85 to 135 PSI at 21,000 ft and above. During extremely cold starts, oil pressure may reach 200 PSI. Oil pressure between 60 and 85 PSI is undesirable; it should be tolerated only for the completion of the flight, and then only at a reduced power setting not exceeding 1,100 ft-lbs torque. Oil pressure below 60 PSI is unsafe; it requires that either the engine be shut down, or that a landing be made as soon as possible with minimum power to sustain flight. Fluctuations of ±10 PSI are acceptable.

3. A minimum oil temperature of 55°C is recommended for fuel heater operation at takeoff power.4. Oil temperature limits are -40°C and 99°C. However, temperature of up to 104°C are permitted for a maximum time of 10 minutes.5. These values are time limited to 5 seconds.6. High ITT at ground idle may be corrected by reducing accessory load or increasing N1 RPM.7. At approximately 70% N1.8. Cruise torque values vary with altitude and temperature.9. This operation is time limited to one minute.10.These values are time limited to 10 seconds.11.Values above 99°C are time limited to 10 minutes.

Powerplant



Operating Condition

SHP Torque(ft-lbs)1


N1 RPM N1 % PropRPM

N2

Oil Press(PSI)2

Oil Temp°C2,3

Starting ___ ___ 10005 ___ ___ ___ ___ -40 (min)

Low Idle ___ ___ 7506 22,875 61 (min) 12 60 (Min) -40 to 99

High Idle ___ ___ ___ ___ 7 ___ ___ -40 to 99

Takeoff 6 850 2230 800 38,100 101.5 2000 100 to 135 0 to 99 Max Continuous and Cruise

850 22308 770 38,100 101.5 2000 100 to 135 0 to 99


850 22308 770 38,100 101.5 2000 100 to 135 0 to 99

Max Reverse9 ___ ___ 750 ___ 88 1900 100 to 135 0 to 99

Transient ___ 27508 850 38,50010 102.610 22005 2002 0 to 10411

Table 5K-3: King Air B200 Engine Operating Limits (PT6A-42); S/Ns BB-1439, BB-1463; BL-139 and subsequent; BW-1 and subsequent

1. Torque limit applies within range of 1,600 to 2,000 propeller RPM (N2). Below 1,600 RPM, torque is limited to 1,100 ft-lbs.2. When gas generator speeds are above 27,000 RPM (72% N1) and oil temperatures are between 60 and 71°C,normal oil pressures

are: 100 to 135 PSI below 21,000 ft and 85 to 135 PSI at 21,000 ft and above. During extremely cold starts, oil pressure may reach 200 PSI. Oil pressure between 60 and 85 PSI is undesirable; it should be tolerated only for the completion of the flight, and then only at a reduced power setting not exceeding 1,100 ft-lbs torque. Oil pressure below 60 PSI is unsafe; it requires that either the engine be shut down, or that a landing be made as soon as possible with minimum power to sustain flight. Fluctuations of ±10 PSI are acceptable.

3. A minimum oil temperature of 55°C is recommended for fuel heater operation at takeoff power.4. Oil temperature limits are -40°C and 99°C. However, temperature of up to 104°C are permitted for a maximum time of 10 minutes.5. These values are time limited to 5 seconds.6. High ITT at ground idle may be corrected by reducing accessory load and/or increasing N1 RPM.7. At approximately 70% N1.8. Cruise torque values vary with altitude and temperature.9. This operation is time limited to one minute.10. These values are time limited to 10 seconds.11. Values above 99°C are time limited to 5 minutes.12. 1,000 RPM for McCauley propeller and 1,180 RPM for Hartzell propeller.



Instrument Red LineMinimum Limit

Yellow ArcCaution Range

Green ArcNormal Operating

Red LineMaximum Limit

Interstage Turbine Temperature

— — 400 to 750°C1 400 to 800°C2

750°C1

800°C2

Torquemeter — — 400 to 2230 ft-lbs 2230 ft-lbs Propeller Tachometer — — 1600 to 2000 RPM 2000 RPM Gas Generator Tachometer — — 61 to 101.5%3 101.50%Oil Temperature — — 10 to 99°C 99°C Oil Pressure 60 PSI 60 to 100 PSI3 105 to 135 PSI1

100 to 135 PSI2

85 to 135 PSI3

200 PSI 135 PSI3

Table 5K-4: Powerplant Instrument Markings1. King Air 200.2. King Air B200 S/Ns BB-743 to 1433 with exceptions; BL-37 to 138.3 King Air B200 S/Ns BB-1439, 1444 and subsequent except 1463; BL-139 and subsequent; BW-1 and subsequent


5K-23For Training Purposes Only

Powerplant

Lubrication System

PROP

THRUSTBEARING CHIP

DETECTOR

PROPELLERGOVERNORAND BETACONTROL

PROPELLER SHAFTOIL TRANSFER TUBE

REDUCTIONGEARS

TORQUEMETER OILCONTROL VALVE

BEARINGTORQUE LIMITERAND TORQUEMETERPRESSURE INDICATOR

POWERTURBINESHAFT

COMPRESSORSHAFT

COMPRESSOR

OIL FILTER ANDCHECK VALVEASSEMBLY

OIL FILLERAND DIPSTICK

CENTRIFUGALBREATHER

OIL COOLERBYPASS

THERMOSTATICDIVERTER VALVE(IF FITTED)

OIL-TO-FUELHEATER

EXTERNALSCAVENGEPUMP(DUAL ELEMENT)

INTERNALSCAVENGEPUMP(DUAL ELEMENT)

OIL TANK DRAIN

ACCESSORYGEARBOX DRAIN

TO OIL PRESSUREINDICATOR

TO OIL TEMPERATUREINDICATOR

FILTERBYPASSVALVE

PRESSUREREGULATING

AND RELIEFVALVE

OIL PRESSPUMP

PRESSURE OIL (105 TO .135 PSIG)

SCAVENGE OIL

RESERVOIR


5K-24 For Training Purposes Only


Powerplant



Powerplant SystemsPowerplant systems include: lubrication engine fuel and control ignition power control bleed air.

LubricationThe lubrication system provides oil under pressure to the engine for cooling and lubrication of the bearings and bushings; it also operates the propeller, heats fuel, indicates torque, and cleans the engine.The lubrication has three subsystems: pressure, scavenge, and breather. The pressure system supplies oil under pressure from an integral oil tank to the engine bearings and bushings. The scavenge system draws oil either by gravity or scavenge pumps from the bearings and engine sumps and returns it to the oil tank. The breather system vents the bearing compartments and gearboxes to the atmosphere, oil tank, or other parts of the lubrication system to prevent over pressurization and to release air trapped in the lubrication system.The main components of the lubrication system are: Oil tank Centrifugal breather Chip detector Oil pressure pump Oil pressure relief/pressurizing valve Oil filter Oil cooler Fuel heater Oil scavenge pumps Oil pressure gauge Oil temperature gauge.

Pressure and temperature sensors in the oil line downstream of the oil pressure pump drive the oil pressure and oil temperature gauges.



Fuel Control System

TORQUELIMITER

TORQUE METERPRESSURE

L OR R ENGINEFUEL CONT HEAT

PURGEVALVE

TONACELLETANK

NO

1

OR

2

DUAL

FED

BUS

CONDITION LEVERMICROSWITCH

P AIR3

MINIMUMFLOWSTOP

MINIMUMPRESSURIZING

VALVE

FUELCUT-OFFVALVE

POWERLEVERS

ENGINE DRIVENFUEL PUMP

CONDITIONLEVERS

TRANSFERVALVE

DUMPVALV

FUEL TOPPINGGOVERNOR

N

1

N f

5A

OIL TOFUEL HEAT

EXCHANGER

ENGINE DRIVENFUEL BOOST PUMP

Powerplant



Otl TmnkThe engine oil tank is an integral part of the compressor inlet case and is in front of the accessory gearbox. The tank has a combined cap and dipstick at the 11 o’clock position that attaches to the filler tube extending through the gearbox housing into the tank. Marks on the dipstick indicate amount of oil in U.S. quarts required to fill tank. Total capacity of the tank is 2.3 U.S. gallons (total oil system capacity 3.5 U.S. gallons). Oil should be added when oil system is four U.S. quarts low.

Figure 5K-11: Oil Tank

Chnartfugml BrhmachrA centrifugal breather in the accessory gearbox separates oil and air. The breather, driven by the starter/generator gearshaft in the accessory gearbox, uses a two-pass impeller to throw oil outward where it drains into the gearbox. Air separated from the oil travels through the hollow rear section of the gearshaft and into a passage in the accessory diaphragm. The air then passes through a tube to a breather boss on the accessory gearbox housing where it vents to the atmosphere.

Cctp DhahSaorA magnetic chip detector in the bottom of each engine’s reduction gearbox activates the red L/R CHIP DETECT annunciators. As ferrous particles accumulate on the detector in sufficient quantity, they bridge two contacts completing the circuit to illuminate an annunciator. Illumination of the annunciator indicates oil contamination and possible engine failure.If a CHIP DETECT annunciator illuminates in flight, follow the emergency procedure for your aircraft. Have maintenance inspect the engine for signs of damage or wear.

Prhssurh PuepThe pressure pump consists of two gears in a cast housing submerged in the oil tank. The accessory gearbox shaft drives the pressure pump and the internal scavenge pump.



Ignition System

IGNITION ON

ON

OFF

STARTERONLY

IGNITIONSTART

SWITCH

ARM

OFF

AUTOIGNITIONSWITCH

400 FT-LBTORQUESWITCH

OIL

IGNPOWER

5A

N O.3ORN

O.4DUALFEDBUS

IGNITIONEXCITER

STARTCONTROL

5A

PURGEVALVE

Powerplant



Otl Prhssurh Rhlthf/Prhssurtztng VmlvhA pressure relief/pressurizing valve secured to a boss on top of the oil pressure pump regulates engine oil pressure. Oil in excess of normal operating pressure returns directly to the oil tank.

Otl FtlahrThe oil filter, downstream of the oil pump, contains a replaceable cartridge-type filter, a bypass valve, and a check valve. If the filter begins clogging due to contaminants or cold oil, the bypass valve opens to route oil around the filter. The check valve keeps oil in the tank when the engine is not operating.

Otl CoolhrAn externally mounted oil cooler below the accessory gearbox uses air flowing through the engine inlet duct to cool hot oil. The oil cooler has a thermostat-controlled bypass valve that routes oil around the cooler until the oil reaches a set temperature. The cooler also contains a pressure relief valve that allows oil to bypass the cooler if it clogs.To provide additional oil cooling, the system uses a thermostat-controlled bypass louver. As oil temperature reaches 71°C (160°F), the louver begins opening to increase airflow; at 82°C (180°F) the louver opens fully.

Fuhl HhmahrAn oil-to-fuel heat exchanger uses hot engine oil to warm fuel. The heat exchanger on the bottom of the accessory gearbox uses hot oil provided by the No.2 scavenge pump.

Smvhngh PuepsFour scavenge pumps draw oil from the engine bearings and gearbox sumps into the oil tank. Two scavenge pumps are inside the accessory gearbox and two are on the outside of the accessory gearbox.



Otl Prhssurh GmughA pressure-transmitter downstream of the oil pressure pump on the rear accessory case senses oil pressure. The transmitter provides an electrical driving signal to the oil pressure gauge in the cockpit.The oil pressure gauge indicates oil pressure from 0 to 200 PSI. A green colored arc and two red redial lines indicate the normal, minimum, and maximum operating pressures. On aircraft with PT6A-41 engines, the first radial line (minimum) is at 60 PSI and the second (maximum) is at 200 PSI; the green arc (normal operating) covers 105 to 135 PSI. On aircraft with PT6A-42 engines, the first radial line is at 60 PSI and the second is at 200 PSI; the green arc covers 100 to 135 PSI.Some indicators have a yellow/green arc that covers the 85 to 105 PSI (PT6A-41) or 85 to 100 PSI (PT6A- 42) range. This band indicates the normal oil pressure for operation at 21,000 ft or above.

Otl Thephrmaurh GmughA temperature sensor downstream of the oil pressure pump on the rear accessory case senses oil temperature. The sensor drives the oil temperature gauge in the cockpit.The oil temperature gauge indicates oil temperature from 0 to 160°C. A green colored arc and a red radial line indicate the normal and maximum oil temperatures. The green arc (normal operating) covers 10 to 99°C and the red radial line at 99°C indicates the maximum operating temperature.

Figure 5K-12: Oil Temperature Gauge

DtsartbuatonUnder pressure, oil flows from the oil pump to the oil filter. From the filter outlet, the oil flow separates into several paths. Transfer tubes and cored passages then carry the oil to the reduction gearbox, front accessories, the No. 1, 2, 3, and 4 bearings, and the propeller. Strategically located strainers remove particles from the oil before the flow reaches nozzles that provide a fine spray of lubricating oil to the frontand rear faces of the bearings.

Oil Temperature Limitations

Starting . . . . . . . . . . .-40°CTakeoff (-41) . . . .10 - 99°CTakeoff (-42) . . . . .0 - 99°CMax Continuous (-41) . . . . . . . . . .10 - 99°CMax Continuous (-42) . . . . . . . . . . .0 - 99°CTransient . . . . . . 0 - 104°CSee Operating Limitations in the Turboprop section of this chapter for a complete listing.

Powerplant



After lubricating the various components of the engine and propeller, the oil gravity drains from the bearings and gearboxes where collection tubes carry the oil under suction from the scavenge pumps to the oil tank.

Engine Fuel and ControlThe engine fuel and control system supplies and controls the flow of fuel from the aircraft fuel system to the engine.Engine fuel and control consists of: boost pump oil-to-fuel heater fuel pump Fuel Control Unit (FCU) torque limiter flow divider and dump valve fuel manifold and nozzles two fuel drain valves fuel flow gauge FUEL PRESS annunciators.

Boosa PuepThe engine-driven low pressure boost pump provides fuel from the aircraft fuel system to the oil-to-fuel heater. The pump operates when the gas generator shaft (N1) turns and provides sufficient flow pressure for lubrication and cooling of the high pressure pump.

Otl-ao-Fuhl HhmahrFrom the aircraft fuel system, fuel flows through the oil-to-fuel heater where hot engine oil heats the cold fuel. The oil-to-fuel heater, below the fuel pump, consists of a two-pass oil circuit and a two-pass fuel circuit. It heats fuel before it reaches the fuel pump.The heater has a temperature-sensing oil bypass valve that senses the temperature of fuel flowing from the unit. As fuel leaving the heater reaches 21°C (70°F), the valve begins closing; less oil travels through the heater. Once the fuel temperature reaches 32°C (90°F), the valve closes completely to bypass oil back to the oil tank.

Fuhl PuepThe PT6A uses a gear-type, enginedriven positive displacement high pressure fuel pump. Fuel from the oil-to-fuel heater enters the pump through a 74 micron strainer. If the strainer blocks, fuel pressure overcomes a spring to allow the passage of fuel past the strainer. From the strainer, the pump gears pressurize the fuel before it reaches a second filter in the pump outlet. The filter has a bypass valve that bypasses fuel around a clogged filter. After passing through the filter, the fuel flows to the fuel control unit at approximately 800 PSI.

Oil Pressure Limitations

Low Idle (minimum) . . . . . . . . 60 PSI NormalBelow 21,000 ft. . . . 100 - 135 PSIAbove 21,000 ft. . . . . 85 - 135 PSISee Operating Limitations in theTurboprop section of this chapterfor a complete listing.



Fuhl Conarol UntaThe Fuel Control Unit (FCU) on the rear flange of the fuel pump determines the fuel schedule for the engine in response to power lever position. The FCU provides the proper fuel flow for all engine operating speeds.A splined fuel pump/FCU coupling shaft connects to the high pressure fuel pump. This shaft provides a signal to the governing section of the FCU proportional to compressor turbine speed (N1). In turn, the FCU regulates the flow of fuel to control engine power by governing compressor turbine speed (N1).The FCU has three sections: the metering section, pneumatic section (computing), and governor section. Metering SectionThe metering section contains: metering valve bypass valve spill valve high pressure relief valve minimum pressuring valve cut-off valve.

The metering section uses unmetered fuel (P1), metered fuel (P2), governing air (Py) enrichment air pressure (Px), and compressor discharge air (P3) to vary the fuel flow to the flow divider. Compressor discharge air comes through an external line from the diffuser section of the gas generator case. The metering section receives fuel at pump pressure (P1) and provides metered fuel (P2) to the fuel flow divider. A metering valve and bypass valve system establish the fuel flow to the engine. The metering valve consists of a contoured needle moving in a sleeve that regulates fuel flow; the effective orifice area changes to meet engine requirements.The bypass valve works to maintain a constant pressure differential (P1 to P2) across the metering valve. The bypass valve has a chamber separated by a diaphragm; one side receives un-metered fuel (P1), and other receives metered fuel (P2) from the metering valve. A diaphragm and a spring control a valve that returns P1 fuel to the high pressure pump. As the pressure differential (P1 to P2) between the inlet and outlet sides of the metering valve varies, the valve opens and closes to maintain a constant differential pressure.The metering section also has a spill valve that bypasses excess metered fuel back to the fuel pump to avoid hot starts. The valve works similarly to the bypass valve. It uses enrichment air pressure (Px) from the pneumatic section of the FCU to operate a diaphragm that opens and closes a valve. During engine start there is no Px pressure acting against the diaphragm; fuel flows back to the fuel pump. As compressor turbine speed increases, compressed discharge air (P3) pressure increases. An increase in P3 pressure increases Px pressure. The increasing Px pressure forces the diaphragm to move and results in the closing of the valve.The high pressure relief valve prevents excessive unmetered fuel pressure (P1) in the FCU. Excessive P1 pressure overcomes a spring in the valve to allow the flow of fuel back to the fuel pump. Once fuel pressure stabilizes at an acceptable level, the valve closes.

Powerplant



The minimum pressurizing valve in the output line to the fuel divider maintains sufficient pressure within the FCU to maintain correct fuel metering.The cut-off valve controls the flow of metered fuel from the metering valve to the fuel divider. During normal engine operation, the cut-off valve does not restrict the flow of fuel. When operated by the FCU speed set lever, the valve closes to cut fuel flow.

PnhuematS hSatonThe pneumatic (computing) section of the FCU consists of an evacuated (acceleration) bellows and a governing bellows; the bellows connect through a common rod. The bellows use air pressure for operation. Governing air pressure (Py) acts against the outside of the governor bellows; enrichment air pressure (Px) works against the inside of the governor bellows and the outside of the acceleration bellows. As the bellows react to the changing pressures, a torque tube on the bellows’ connecting rod moves. A torque-tube assembly transmits movement of the torque tube, which in turn, moves the metering valve in the FCU metering section to vary fuel flow.Px and Py change with varying engine operating conditions. During engine acceleration, both pressures increase simultaneously. The bellows react to the changing pressure by opening the metering valve to increase compressor turbine speed (N1). As the engine reaches the desired speed, Py decreases, and the bellows react to begin to close the metering valve. During deceleration, both pressures decrease; the bellows move to close the metering valve to its minimum flow stop.

Govhrntng hSatonThe governing section of the FCU uses a mechanical governor with flyweights to regulate governing air pressure (Py) to the governing section governor bellows. The governing section, in turn, controls the metering valve in the metering sectionThe governor section consists of spinning flyweights on a drive shaft connected to the high pressure fuel pump drive coupling. The section also has a spring-bias lever and a governor lever that pivot on the same point. A roller on the end of the spring bias lever contacts the governor spool. Once the spring bias moves due to governor spool movement, it contacts the governor lever; both levers then move simultaneously. Movement of both levers operates the governor valve that regulates governing air pressure (Py) and ambient air pressure (Pa).As compressor turbine speed (N1) increases, the flyweights spin faster and begin moving outward due to centrifugal force. The flyweights then begin pressing against the governor spool. As the force against the governor spool increases, the spool moves forward against the roller on the end of the speed bias lever.The speed bias lever moves until it contacts the governor lever; they begin moving simultaneously. As the two levers move against spring pressure, they operate the governor valve. As the governor valve opens, it allows governing air pressure (Py) to escape from the governor bellows in the pneumatic section. The flow of governing air pressure acts on the bellows and moves the metering valve to control fuel flow.



Torquh LtetahrIn addition to the FCU that controls engine speed through fuel flow metering, the torque limiter monitors torque pressure oil and varies governing air pressure (Py) to the fuel control unit.The torque limiter on the torquemeter pressure transmitter boss contains a sealed bellows connected to the torquemeter oil pressure output, a pivot and spring assembly (pivoted platform), a Py bleed orifice, and a torque-limit adjusting spring.During normal operation, the pressure signal cannot overcome spring pressure; the restrictor assembly remains closed. As torque increases,the pressure signal to the bellows increases until it is strong enough to overcome spring pressure. The pivot and spring assembly move to open the bleed orifice.Opening of the bleed orifice vents Py pressure from the metering section of the fuel control unit to reduce fuel flow. Once engine compressor turbine speed (N1) decreases, torque decreases. The spring overcomes the pressure signal, and the bleed orifice closes.

Flow Dtvtdhr mnd Duep VmlvhAfter leaving the FCU, metered fuel flows to the fuel divider and dump valve on the gas generator case. The divider controls fuel to the primary and secondary fuel manifolds that supply the primary and secondary fuel nozzles. During engine start, the flow divider supplies the primary manifolds. As the engine accelerates, the divider begins supplying the secondary manifolds until both he primary and secondary manifolds receive equal amounts of fuel.During engine operation, fuel pressure overcomes a spring in the flow divider and dump valve to allow fuel to the primary and secondary manifolds. Once the cut-off valve in the FCU cuts fuel flow to the divider, spring pressure moves a piston to cut fuel flow to the manifolds and connect the primary and secondary manifolds. Once the dump valve opens, fuel flows into an EPA collector tank (BB- 2 through BB-665), or pressure in a purge tank (BB-665 and subsequent) forces fuel from the manifolds into the combustion chamber where it is burned. During this process, there may be a transient surge in N1.

Fuhl Mmntfold mnd NozzlhsFrom the flow divider, dual fuel manifolds distribute fuel to the primary and secondary fuel nozzles. The manifold consists of 14 adapter assemblies connected by fuel transfer tubes; each transfer tube carries two flows of fuel. Each adapter assemblies consists of a simplex fuel nozzle with a swirl tip and a sheath that extends through the gas generator case. The sheath has holes that allow air flowing between the gas compressor case and combustion chamber liner to cool the tip of the nozzle and assist in fuel atomization. The swirl tip delivers a fine spray of fuel into the combustion chamber.Fuel PurgeThe fuel purge system is designed to ensure that any residual fuel in the fuel manifolds is consumed during engine shutdown. During engine operation, compressor discharge air (P3 air) is routed through a filter and check valve,

Powerplant



pressurizing a small air tank On engine shutdown, the pressure differential between the air tank and fuel manifolds causes air to be discharged from the air tank through a check valve and into the fuel manifold system. The air forces all residual fuel remaining in the fuel manifolds out through the nozzles and into the combustion chamber. The fuel forced into the combustion chamber is consumed.

Fuhl Drmtn VmlvhsTwo drain valves carry fuel from the gas generator case. Compressor discharge air pressure (P3) keeps the valves closed during engine operation.

Fuhl Flow GmughA transmitter between the engine boost pump and engine-driven high pressure pump measures the flow of fuel to the engine. The transmitter drives the fuel flow indicator in the cockpit. The indicator, graduated from 0 to 6, indicates fuel flow in Pounds Per-Hour (PPH) from 0 to 600 PPH.Depending on the aircraft, the fuel flow indicating system operates on DC or AC power (BB-225 and subsequent).

Figure 5K-13: Fuel Flow Gauge

FUEL PRE AnnunStmaorsThe red FUEL PRESS annunciators in the warning annunciator panel monitor engine-driven fuel boost pump pressure. The annunciator illuminates if boost pump outlet pressure drops below 10 PSI.Normally, the boost pump provides sufficient head pressure for proper operation of the fuel pressure pump. If the boost pump fails, the pressure pump does not receive sufficient pressure to operate normally; the pump begins cavitating.If a FUEL PRESS annunciator illuminates, turning on a standby pump provides sufficient pressure for proper pressure pump operation. Once fuel pressure exceeds 10 PSI, the FUEL PRESS annunciator extinguishes.



IgntatonThe PT6A ignition system on the King Air 200/B200 consists of: Ignition exciter Ignition leads Two igniters Ignition switch Auto-ignition system.

The PT6A also has an auto-ignition system that automatically activates to prevent engine flameout at low engine torque.

Igntaton ExStahrThe ignition exciter is a solid-state capacitance-discharge unit containing circuitry, transformers, and diodes that converts a DC input into a high-voltage output. Although designed to operate on 28 VDC, the ignition system can function on 9 to 30V DC power.The left and right ignition exciters receive power through separate ignition power relays. The left exciter receives power from the No. 3 Dual Fed bus and the right exciter receives power from the No. 4 Dual Fed bus.When operating, the ignition exciter supplies a high-voltage charge (approximately 3,500V) through high tension cables to the igniters in the combustion chamber. The design of the ignition system has a dividing and set-up transformer network that supplies separate high-voltage electrical charges to each igniter. If an igniter fails (i.e., shorts or opens), the other igniter still functions.

Igntahr PlugsTwo igniter plugs on each engine, at the 4 and 9 o’clock positions on the gas generator case, extend into the combustion chamber. When supplied with a high voltage charge, they create a spark that ignites the fuel/air mixture. Depending on input voltage to the ignition exciter, each igniter sparks at a rate from 0.8 to 1.0 times per second (9V) to 1.4 to 4.0 times per second (30V).

Powerplant



Igntaton wtaSchsThe three-position (ON/OFF/ STARTER ONLY) IGNITION AND ENGINE START switches on the pilot subpanel control power to the ignition and engine starting systems. In ON, each switch provides power from the respective DC electrical system to the respective -ignition power relay. The relay closes to provide power to the ignition exciter. During ignition system operation, the respective green IGNITION ON annunciator illuminates. The STARTER ONLY position provides power to the engine starter only; the ignition system is inoperative.

Figure 5K-14: Ignition Switches

Auao-Igntaton ysaheThe automatic ignition system automatically energizes the ignition system to keep an engine operating during low torque power conditions. The system uses a pressure switch next to the torquemeter to continuously monitor torquemeter oil pressure. With the ENG AUTO IGNITION switch on the pilot subpanel in ARM, a drop in engine torque below 400 ft-lbs closes the pressure switch. Once the pressure switch closes, it supplies power to the ignition power relay; the relay closes to power the ignition exciter and opens the purge valve. During auto-ignition operation, the respective IGNITION ON annunciator illuminates. Once torque exceeds 400 ft-lbs, the system deactivates the igniters and returns to the armed mode.

Figure 5K-15: Auto-Ignition System



Engine Power Control Engine power control for the PT6A on the Super King Air consists of propeller levers, power levers, and condition levers. See Propeller section of this chapter for propeller lever discussion.

Powhr Lhvhr The power lever controls engine power from maximum takeoff thrust to full reverse thrust.Each power lever in the cockpit connects through a system of bellcranks and pushrods to a controlex (flexible) cable. The cable routed under the cockpit floorboards and through the leading edge of the wing center section connects to a cam assembly in front of the fuel control unit. Each end of the cable has threaded rods that allow control rigging at the power lever and engine ends. The FCU control rod connects the cam assembly to the FCU control arm; the propeller reversing linkage connects the cam assembly to the propeller reversing lever. The linkage is a flexible push-pull cable.In response to power lever movement above idle, the FCU control arm moves to reposition the speed scheduling cam in the governor section of the FCU. Movement of the cam moves the cam follower lever to increase governor spring force. Spring force overcomes governor flyweight force to close the governor valve. The metering valve opens in response to the increase in governing pressure (Py) and enrichment pressure (Px) to admit more fuel to the engine; the engine accelerates.The governor senses any variation in engine speed from the selected speed. The flyweights then increase or decrease speed to vary the pressure on the governor spool and open or close the governor valve. The opening and closing of the governor valve varies the pressures acting on the FCU pneumatic section bellows. Bellow movement varies the metering valve position to either increase or decrease fuel flow; the engine re-establishes the selected speed.Moving the power lever toward idle repositions the speed scheduling cam in the FCU governor section to a lower point on the cam. This reduces spring pressure to allow the governor valve to open; Py pressure drops and the metering valve closes; fuel flow then drops and the engine decelerates.Moving the power lever up and past a detent to the REVERSE position integrates propeller pitch and the FCU. The propeller goes into reverse pitch and the engine accelerates.

Condtaton LhvhrThe condition levers control the cutoff function of the fuel control unit. The FUEL CUT OFF position closes the cut-off valve in the metering section of the FCU. LOW IDLE limits engine N1 idle to approximately 52% to 56% (depending on propeller installation) and HIGH IDLE limits N1 to 70%. The HIGH IDLE position allows quicker engine acceleration to maximum thrust.

Powerplant



Blhhd AtrThe PT6A uses three separate systems to provide engine bleed air to engine and airframe systems. One system supplies air for engine bearing compartment sealing, the second system provides engine cooling air, and the third system supplies air for compressor bleed valves and airframe systems.Bleed air from the engine provides: Cabin pressurization Pneumatic system Bleed valve operation Fuel Control Unit (FCU) regulation Turbine cooling Combustion chamber cooling Fuel nozzle cooling Engine bearing sealing Optional brake de-ice system.

Coeprhssor Blhhd VmlvhsThe low and high pressure compressor bleed valves regulate interstage compressor pressure air (P2.5) at low and high rotor speeds; excess P2.5 pressure encourages compressor stalling. The low pressure bleed valve is on the gas generator case at the 9 o’clock position; the high pressure bleed valve is at the 3 o’clock position.The low and high pressure compressor bleed valves uses the relationship of interstage compressor air (P2.5) and compressor discharge pressure (P3) to control the opening and closing of the valves.P3 is always greater than P2.5. P3, metered down once in the low pressure compressor bleed valve and twice in the high pressure compressor bleed valve controls the point at which the respective compressor bleed valve opens. A convergent/divergent metering orifice in each valve controls the build-up of P3 within the valve. Normally, the orifice bleeds P3 from the valve and the compressor valve opens to bleed P2.5 overboard. Once P3 exceeds the venting capability of the orifice it builds up to close the valve.

Cooltng mnd hmltngThe engine uses interstage compressor air (P2.5) and compressor discharge air (P3) for cooling of the combustion chamber, fuel nozzles, and turbines and for bearing sealing. Interstage pressure air seals the No.1 bearing labyrinth seal; compressor discharge air seals the No.2 bearing forward and aft labyrinth seal and the No.3 bearing aft labyrinth seal. Air from these seals eventually flows into the accessory gearbox and the engine.Compressor discharge air flowing past the combustion chamber cools the fuel nozzles and assists in fuel atomization. After cooling the combustion chamber, the air flows through the compressor turbine vane ring and into the vanes for cooling. The air then flows from the vane trailing roots and onto the face of the compressor turbine disk.



Compressor discharge air bled from the engine diffuser exit zone cools the compressor, and first and second-stage power turbine disks.

Atrfrmeh Blhhd AtrA pad on the left side of the gas generator case supplies compressor discharge air pressure (P3) from the diffuser zone of the gas generator case. Engine bleed air supplies the cabin pressurization system, surface de-ice system, and through a bleed air injector supplies vacuum for flight instruments.

Powerplant



Propeller SystemsDepending on model, serial number or modifications, the aircraft uses three or four-bladed constant speed, full-feathering, reversible Hartzell or Mc-Cauley propellers. Supplemental Type Certificates (STCs) exist to replace three bladed propellers with four-bladed units that provide increased performance and reduced noise.Each propeller consists of a hub with three or four blades connected to the output shaft of the power turbine. Centrifugal counterweights on each blade, assisted by a feathering spring, move the propeller blades toward the feathered position (approximately 86 to 90°, depending on manufacturer) with loss of oil pressure to the propeller control system.See Table 5K-5, following page, for a listing of propellers normally found on the aircraft. Propeller systems include: Propeller de-icing Low pitch stop Primary governor Beta and reverse control Overspeed governor Fuel topping governor Autofeather Synchrophaser Indication and control.

Figure 5K-16: Propeller



Propeller Systems

SPEEDSETTINGSCREW

SPEEDERSPRING

FLYWEIGHT

OVERSPEEDGOVERNOR

PROP GOVERNORTEST VALVE N.C.

AUTOFEATHERDUMPVALVE N.C.

TOCASE

POWERLEVER

PROPLEVER

FROM ENGINEOIL SYSTEM

PRIMARYGOVERNOR

PILOTVALVERETURN

LINEBETAVALVE

OFF

TEST

PROPGOVERNORTEST SWITCH

NO 4 DUAL-FED BUS

SERVOPISTON

FEATHERRETURNSPRING

PISTONSEAL COUNTER

WEIGHT

FOLLOW UPCOLLAR

TRANSFERGLAND

LOW PITCHSTOP NUT

SPRING

RETURN

GOV PUMP

STATIC

5A

PROP GOV

Powerplant



Autofeather System

L AUTOFEATHER

ARM

OFF

TEST

TORQUESWITCH

200 FT-LBS

DUMPVALVE

DUMPVALVE

ARMINGRELAY

TORQUESWITCH400 FT-LBS

POWER LEVER(MICROSWITCHCLOSES AT90% N1)

ARMINGRELAY

R AUTOFEATHER

N.C. N.C.

MECHANICAL CONNECTION

POWER LEVER(MICROSWITCHCLOSES AT90% N1)

5AAUTO FEATHER

NO 2 DUAL FED BUS

TORQUESWITCH400 FT-LBS

AUTOIGNITION

AUTOIGNITION



Type I and Type II Sychrophaser

LH PROP(MASTER)

RH PRIMARYGOVENOR

RH PROP(SLAVE)

RHOVERSPEEDGOVENOR

SYNCHROPHASERACTUATOR

LHOVERSPEEDGOVENOR

CONTROLBOX

RPM

PHASE

RPM

PHASE

ON

OFF

PROPSYNCH

5A

UP

DN

RH GEARUP-LOCK SWITCH

LH PROPRH PROP

RH PRIMARYGOVERNOR

LH PRIMARYGOVERNOR

CONTROLBOX

RPM & PHASE RPM & PHASE

ON

OFF

PROP SYNC

5A

TYPE I SYSTEM

TYPE II SYSTEM

TO #1DUALFEDBUS

TO #1DUALFEDBUS

TOANNUN.LITE

Powerplant



Propeller De-IceEach propeller blade has an electrically-powered single or dual element de-icer boot. Dual element boots have an inner and outer boot that operate independently of each other. The heated boots reduce ice adhesion so that centrifugal action and the air stream remove ice from the propellers. Slip rings and brush assemblies deliver DC power to the de-ice boots. The PROP AUTO switches and PROP INNER-OUTER (PROP MAN-OFF) switches control the system. An ammeter on the copilot subpanel provides an indication of automatic boot operation.See the Ice and Rain chapter for a complete description of the propeller de-ice system.

Low Pitch StopA servo piston in the propeller spider hub connects through links to each blade to control blade angle. From the servo piston, spring-loaded sliding rods connect to a mechanical feedback slip ring behind the propeller. A follow-up arm and carbon block supported by a bracket on the front of the engine rides on the slip ring. The arm transfers movement from the reversing lever through the low pitch valve (beta valve).The mechanical linkage formed by the servo piston, beta valve, slip ring, and rods creates a low pitch stop. The low pitch stop operates by controlling oil flow to the propeller dome. The beta valve maintains blade angle when propeller RPM is lower than that selected by the governor pilot valve position.

Figure 5K-17: Propeller De-Icer Boot Figure 5K-18: Propeller Deicing Control Panel



Primary Governor and Beta ValveBefore engine start, the propeller levers are normally full forward (maximum RPM requested). A linkage from the propeller lever to a cam within the primary propeller governor depresses the speeder spring to a 2,000 RPM setting. Since the flyweights are static at this time, the pilot valve remains fully depressed.During the start cycle, the governor oil pump increases engine oil pressure to approximately 450 PSI. This oil flows through the beta valve past the depressed pilot valve of the governor and into the propeller dome via a transfer gland. Acting as a hydraulic cylinder, the dome moves forward. The propeller blades linked to the dome move from the feathered position to a lower pitch angle. The mechanical linkage of the dome to the beta valve now begins to move the valve which regulates the governor pump oil flow to the dome to establish a hydraulic low pitch stop (see Table 5K-5). The propeller governor pilot valve remains fully open offering no restriction to the oil flow.As the power levers are moved and engine speed increases, the beta valve maintains the low pitch stop until engine RPM reaches the requested governor range. Governor flyweight centrifugal force now equals the speeder spring tension. When governor force equals speeder spring tension the system is in equilibrium (no movement). Additional increase engine RPM causes the governor pilot valve to move to a more closed position; this reduces pressure in the dome resulting in a higher blade angle with a propeller speed decrease to the selected setting.

Propeller Diameter Full Feathered MechanicalReverse Pitch

StopHartzell 3 Blade 98.5 90±5° -9±.4° Hartzell 4 Blade 94 86° -10.5°

McCauley 4 Blade 94 87.5±3° -10±.4°

Table 5K-5:King Air 200/B200 PropellersA calibrated leak in the transfer gland on the propeller shaft allows a small quantity of governing oil to flow back into the engine oil system. This provides a continuous circulation of warm oil within the propeller dome.Movement of the propeller dome aft (less oil pressure) adjusts the propeller blade to a higher pitch angle. This presents increased load and maintains requested governor RPM as horsepower increases. As the dome moves aft, the mechanical linkage moves the beta valve to allow flow of oil through it as long as the propeller governor is active.

Movement of the propeller levers aft reduces speeder spring tension to allow the flyweights to find another balancing RPM. The pilot valve adjusts, in turn, regulating oil entry into the propeller dome with its related blade angle.

Powerplant



As engine power increases, decreases, or an airspeed change occurs (loading), the governor adjusts blade angle to maintain the requested RPM. Power can be reduced below a point that the governor cannot control (pilot valve opens to allow maximum oil flow into the dome). In this event, the blade pitch is now equal to the preset low pitch stop and is again controlled by the beta valve. In summary, the beta valve controls flow to the propeller dome except when RPM is in a requested RPM range (1,600 to 2,000 RPM).Moving the propeller levers aft past a detent causes the governor pilot valve to be lifted above the regulating port; this allows oil from the dome to be dumped back through the governor body to the engine case. Without oil pressure in the dome, the propeller goes to feather due to the centrifugal twisting movement created by propeller rotation and blade counterweights. The normal low pitch stop set by the beta valve is +18° blade angle for three bladed propellers and +11° blade angle for four bladed propellers. Further reduction of the blade angle can be made (i.e., +5° equals approximately zero thrust) by relocating the upper linkage of the beta control arm. Moving the upper pivot point aft allows the propeller dome to move further forward before being hydraulically blocked by the beta valve.Lifting and moving the power levers aft past the normal idle stop adjusts the upper pivot point of the beta control arm. As the propeller dome is allowed to move further forward, the blade angle decreases from its original low pitch stop through the zero thrust angle (beta range) to the full reverse blade angle of -9° for three-bladed propellers and -7° for four-bladed propellers (reverse range). Pulling the power levers further aft into the barber pole area on the power quadrant does not increase negative blade angle but adds fuel (power) resulting in increased propeller RPM (maximum reverse thrust). The primary governor includes fuel topping function for the reversing mode (i.e., reduces 2,000 RPM by 95% to limit propeller speed to 1,900 RPM). Fuel restriction by the fuel topping function prevents propeller overspeed. The propeller levers should be in the full forward position for maximum reversing.

Overspeed GovernorIf the primary governor fails, the overspeed governor prevents propeller overspeed. The governor, on the reduction gear housing, regulates the flow of oil from the propeller pitchchange mechanism. The overspeed governor has flyweight, speeder spring, and pilot valve that operate similarly to the primary governor. This system of flyweights and springs regulates propeller speed to a preset value of 104% of maximum N2 (approxi mately 2,080). If the primary governor fails and the propeller overspeeds, the flyweights move outward to lift the pilot valve; the lifting pilot valve decreases the flow of oil from the propeller dome to increase the blade angle and reduce propeller RPM to overspeed limits.



Autofeather SystemThe autofeather system automatically dumps oil from the servo (dome) to feather a propeller if an engine fails. The system consists of power lever position switches, high and low torque-meter oil pressure switches, control relays, solenoid-operated governor dump valves, arming switches, and system annunciators.The system does not arm until the AUTOFEATHER switch is in ARM and the power levers are at the 90% N1 or greater position; the high and low pressure switches sense torque loss. The high pressure switch actuates at 400 ±50 ft-lbs on the 200 and at 410 ±50 ft-lbs on the B200. The low pressure switch actuates at 225 ±65 ft-lbs on the 200 and at 260 ±50 ft-lbs on the B200. Each high pressure switch controls power to the opposite engine’s arming-lightout relay; once torque exceeds approximately 400 ft-lbs, the high pressure switch actuates and energizes the opposite engine’s arminglight-out relay; the opposite engine’s green AUTO-FEATHER annunciator illuminates.

Figure 5K-19: Autofeather System

If an engine fails, torquemeter oil pressure begins to drop. Once the torque reaches 400 ft-lbs, the switch on the failing engine de-energizes the opposite engine’s AUTOFEATHER annunciator.As the engine continues to fail and its torque decreases further to the low pressure switch actuation point. Actuation of the low pressure switch provides a ground for the arming relay and the normally closed dump valve solenoid. The arming relay closes extinguishing the annunciator and provides power to the dump valve. The dump valve reduces dome oil pressure to zero and the counterweights, springs, and centrifugal force feathers the propeller. As a safety feature to prevent both autofeathering of both propellers, autofeathering of one propeller disables the autofeathering system of its opposite; the opposite AUTO-FEATHER annunciator extinguishes.

Powerplant



Propeller SynchrophaserTwo types of propeller synchrophasers are on the aircraft: Type I – BB-2 through BB-934,BB-991, BT-1 through BT-22, BL-1 through

BL-41 and BN-1 Type II – BB-935 through BB-990, BB-992 and subsequent,BT-23 and

subsequent, BL-42and subsequent, and BN-2 andsubsequent.The Type I synchrophaser varies the right propeller (slave) to match left propeller (master). The Type II system does not operate on the slave/master principle. The system adjusts the RPM of a slower propeller to match its opposite. Both systems have a PROP SYN ON/OFF switch and propeller synchroscope on lower left instrument panel.

Typh I ynScropcmshr A magnetic pickup in each propeller overspeed governor and a phase pickup on each propeller bulkhead senses propeller rotation speed. The pickups send electrical pulses to a control box. The control box converts the electrical pulses into correction commands for the synchrophaser actuator on the right engine. A flexible shaft and trimmer assembly adjust (trims) the right propeller governor linkage to match the right RPM with the left (master) RPM. The trimmer between the governor control arm and the control cable, screws in and out to adjust the governor; propeller control lever position does not change.The synchrophaser has a ±30 RPM authority range from the governor setting. This prevents the synchrophaser from trying to match the slave propeller with a failing master. This possible 30 RPM reduction necessitates a restriction against using the Type I synchrophaser during takeoff and landing. Before turning the synchrophaser ON, synchronize both propellers manually. Turning the synchrophaser OFF automatically adjusts the trimmer to its center position.

Typh II ynScropcmshrThe Type II synchrophaser uses ferrous metal targets on the propeller bulkheads that induce an alternating current as they rotate past a stationary magnetic speed pickup. As propeller speed varies the waveform of the alternating current varies. The system control box receives the signals from both propeller pickups and sends a correction signal to slower propeller’s primary governor. The governor increases the speed of the slower propeller to match the other.The system has limited control over the propellers (approximately 25 ±2 RPM). It cannot reduce the RPM of either propeller below the speed set by the propeller control lever. This allows the use of the Type II synchrophaser during takeoff and landing as the system does not decrease propeller speed to match the other.Synchronize the propellers manually then turn the synchrophaser ON. This keeps the synchrophaser within its authority range.



Propeller Synchroscope

CATTON Operation of the three-bladed propeller during approach in the 1,750 to 1,850 RPM range should be avoided as it may cause ILS interference.

The propeller synchroscope on the lower left instrument panel provides a visual indication of propeller synchronization. A black and white cross pattern spins in the direction of the faster propeller. If the right propeller is operating at a higher RPM than the left; the cross spins right (clockwise). With the synchroscope the pilot can manually adjust propeller RPM until the indicator stops.

Figure 5K-20: Propeller Synchroscope

Powerplant



Prop TachometerThe reduction gearbox drives a tachometer-generator that electrically drives a propeller tachometer in the cockpit. The indicator uses two pointers to indicate propeller RPM; the inner (shorter) pointer indicates RPM in thousands and the outer (longer) needle indicates RPM in hundreds. A green colored arc from 1,600 to 2,000 RPM covers the normal operating speed of the propeller; a red radial line at 2,000 RPM indicates the maximum operating speed.

Figure 5K-21: Propeller Tachometer

For aircraft S/N BB-1439, BB-1444 and subsequent, BN-5 and subsequent, and BL-139 and subsequent, the N1 percent of RPM indicator is a single needle with LED readout. The normal (green arc) is from 62% to 101.5%. The red arc reading is 101.5%.

Figure 5K-22: RPM indicator

Propeller Limitations

Takeoff . . . . . .2,000 RPMMax Continuous. . . . 2,000 RPMMax Cruise . . . 2,000 RPMCruise Climb . . 2,000 RPMMax Reverse . . 1,900 RPMTransient . . . . . 2,200 RPM



Power LeversThe power levers mechanically connect through a cam box to the fuel control unit, beta valve, and the fuel topping governor. A detent at the IDLE position prevents inadvertent movement of the power lever into BETA/REVERSE. In the forward thrust range, the power levers control gas generator RPM. In beta range (top of the reverse range marks), lifting and moving the power levers aft reduces blade angle resulting in less residual propeller thrust. In the reverse range, the power levers select a negative blade angle proportionate to position of the lever, control fuel flow to maintain reverse power, and reset the power turbine governor from its normal 100% to 95%.

Figure 5K-23: Power Levers

Propeller Control LeversThe propeller control levers control propeller speed within the range of the primary governor (1,600 to 2,000) RPM. Full forward sets the primary governor at 2,000 RPM; full aft (at feathering detent) sets the governor at 1,600 RPM. Pulling the propeller control lever past feathering detent lifts the pilot valve in the primary governor and dumps oil from the system through the governor into the engine case to allow propellers to feather.

Figure 5K-24: Propeller Control Levers

Powerplant




PreflightDuring the exterior preflight inspection, check engine oil level; add oil if required. Secure oil tank cap. Open the engine cowlings to check lines, linkages, hoses, and accessories for abnormal or excessive wear, security, fraying, or leakage. Check the engine air inlet screen for damage. Check that all cowl latches and camlocks are secure. Check the condition of the engine intake; the ice vanes should be retracted and the bypass door under the engine flush with the cowling.Examine exhaust pipes and scuppers for cracks in the welds. Check condition of the scupper inside the stack. heck the propellers for cracks, nicks,corrosion, or other damage. The blades should not twist. When rotating the propeller, there should not be any unusual noise or binding. Check the propeller deice boots for security and condition. Examine the spinner and blade roots for signs of oil leaks.

Emergency ProceduresEmergency procedures for the powerplant and powerplant systems include: emergency engine shutdown engine fire on ground engine failure during ground roll engine failure after lift-off engine failure in flight below VMCA

engine flameout (2nd engine) low oil pressure airstart/starter assist windmilling engine and propeller Airstart propeller overspeed (up to 2,080 RPM) propeller overspeed (above 2,080 RPM).

EehrghnSy Engtnh cuadownPerform an emergency engine shutdown if there is: an unscheduled engine torque increase and the engine fails to respond to

power lever movement an engine fire in flight an engine failure in flight an illumination of a magnetic chip detector annunciator (ENG CHIP DETECT) low oil pressure oil or fuel leak.

Usually, an unscheduled increase in torque is accompanied by an increase in fuel flow, ITT and N1; this indicates a possible fuel control unit failure.



Begin the engine shutdown by moving the condition lever to FUEL CUT OFF to cut the flow of fuel to the engine; the engine begins winding down. Move the affected engine propeller lever to FEATHER to reduce propeller induced drag. Move the fuel firewall valve to CLOSED. If there is an engine fire, actuate the engine fire extinguisher.Turn the affected engine ENG AUTO IGNITION and GEN switches OFF. Turn the propeller synchrophaser OFF. Monitor the electrical load to prevent overloading the operating generator.

Engtnh Ftrh On GroundIf an engine fire occurs on the ground,move the affected engine condition lever to CUT OFF to cut fuel to the engine. CLOSE the firewall shutoff valve. Move the starter switch to STARTER ONLY to motor the engine; this helps in blowing an internal/exhaust fire out, clearing fuel from the engine, and cooling the engine. If necessary, use the fire extinguisher (external fire in the nacelle).

Engtnh Fmtlurh Durtng Ground RollA decrease in ITT, torque, propeller RPM, turbine RPM, and fuel flow are all cockpit indications of an engine failure.If an engine fails during the ground roll (before V1), bring both power levers to IDLE and apply sufficient braking pressure to stop the aircraft. If required, use maximum reverse thrust on the operating engine. Use extreme care when applying reverse thrust on surfaces with reduced traction to avoid loss of directional control with asymmetrical thrust.If there is insufficient runway left to stop the aircraft, move both condition levers to CUT OFF to cut fuel to the engines. Close both fuel firewall valves. Turn the MASTER SWITCH off (gang bar down) to cut electrical power.

Engtnh Fmtlurh Afahr Ltfa-OffIf an engine fails after lift-off, increase power on the operating engine to the maximum allowable. Maintain airspeed at takeoff speed or above. Retract the landing gear to reduce drag.If the aircraft has an autofeather system, do not retard the power lever until the propeller completely feathers; prematurely retarding the power lever interrupts the autofeather sequence.Feather the inoperative engine propeller. Continue the climb out at V2 until clear of all obstacles then accelerate to best rate of climb speed (VYSE). Retract the flaps.As conditions permit, complete the engine shutdown sequence by moving the condition lever to CUT OFF. Close the fuel firewall valve, turn OFF the affected engine AUTO FEATHER, ENGINE AUTO IGNITION, and GEN switches. Monitor the electrical load to prevent overloading the operating generator.

Powerplant



Fmtlurh tn Fltgca Bhlow Mtnteue VMCAIf an engine fails below the minimum control speed (VMCA), reduce power on the operating engine to maintain control. Lower the nose to accelerate above VMCA (86 KIAS).Once above VMCA, adjust power on the operating engine. Perform an engine shutdown.

Engtnh Flmehoua (hSond Engtnh)If after an engine failure the second engine flames out, immediately bring the affected engine’s power lever to IDLE. Move the condition lever to CUT OFF. DO NOT FEATHER the second engine propeller.A feathered propeller reduces airflow through the engine power turbines and may not allow sufficient N1 to develop. Fuel pressures, therefore, may not be sufficient to overcome the minimum pressurizing valves to allow engine relight.This procedure resets the Fuel Control Unit (FCU) to a “zero fuel condition” and allows the pilot to attempt a Windmilling Airstart. If no possibility exists for an airstart, feather both propellers and conduct the Glide Maneuver.

Low Otl PrhssurhThe normal operating oil pressure is between 105 and 135 PSI (PT6A-41) or 100 and 135 PSI (PT6A-42). Oil pressure between 60 and 85 PSI is undesirable and below 60 PSI is unsafe. If oil pressure falls to 40 PSI, the respective red OIL PRESS annunciator illuminates.If oil pressure remains between 60 and 85 PSI, continue the flight with a maximum power setting of 1,100 ft-lbs of torque on the affected engine. If the pressure falls below 60 PSI, shut down the engine and land as soon as possible.

Atrsamra/amrahr AsstsaBefore attempting any airstart determine the cause of the engine failure. Do not attempt an engine airstart if there are any indications of engine damage, fire, or systems failure (i.e., zero N1 RPM, fuel or oil leaks). To reduce the load on the electrical system, turn the cabin temperature mode, aft blower, and windshield anti-icing switches OFF, select AUTO with the BLOWER switch, and select the radar either OFF or STANDBY.Move the power lever to IDLE, place the condition lever in CUT OFF, and OPEN the fuel firewall valve.Turn the ignition and start switch to ON; check that the respective IGNITION annunciator illuminates. Bring the condition lever to LOW IDLE and observe N1. Once N1 exceeds 50%, turn the ignition and start switch to OFF. As required, adjust the power and propeller levers. Turn the generator switch to ON and ARM the engine auto ignition system. Once the generator comes on-line, turn the desired electrical equipment on.



Wtndetlltng AtrsamraAirflow (N1), (fuel flow), and ignition are necessary for combustion. This procedure is most affective in gaining an immediate airstart after a second engine flameout, however, it may be utilized any time an engine has lost power and an immediate shutdown procedure has not yet been accomplished.Airflow is increased through the engine by the windmilling propeller and power turbines. This airflow (approximately 10% N1) is necessary to gain sufficient fuel pressure (80 PSI) to overcome the FCU minimum pressurizing valve.Ensure the power lever is at IDLE and the propeller control set FULL FORWARD. The Condition Lever should be in CUT-OFF, and (if conditions permit), ensure the engine ice-vanes are retracted to maximize inlet air volume. The firewall valve should be confirmed OPEN. Turn the generator switch OFF to prevent the increasing field strength from working against N1 during startup. For optimum air volume through the compressor, use the airspeed and altitude envelope outlined in the checklist. The Auto Ignition switch should be in ARM and the IGNITION annunciator illuminated.Move the Condition Lever to LOW IDLE. If N1 exceeds approximately 10%, the engine should light-off normally. As engine RPM increases and ITT peaks, adjust power as required. After a current limiter check, turn the second generator ON and reinstate the electrical systems previously turned off.

Prophllhr Ovhrsphhd (Up ao 2,080 RPM)The maximum operating speed of the propeller is 2,000 RPM. If the propeller overspeeds up to 2,080 RPM, there is a failure of the primary governor; the overspeed governor limits propeller RPM to 2,080.Retard the propeller lever to decrease propeller RPM. If the RPM does not decrease by retarding the propeller lever, reduce power to 1,800 ft-lbs or less. This avoids exceeding the engine Shaft Horsepower limit of 850. Adjust the unaffected engine’s propeller to 2,000 RPM to match the engine levels as closely as possible.

Prophllhr Ovhrsphhd (Abovh 2,080 RPM)Maximum propeller overspeed is 2,200 RPM and is time limited to 5 seconds. Sustained propeller overspeeds exceeding 2,080 RPM are unapproved and must be avoided. These overspeeds are indicative of the propeller blades not responding to changes in oil pressure. For these situations, the aircraft is designed with an alternate means of controlling propeller speed with power rather than oil and aerodynamics.The fuel topping capability of the primary governor depends implicitly on its proper operation. The governor must be functioning normally for the Fuel Topping Governor (FTG) to have reference RPM. The FTG begins limiting fuel flow to the FUC when the propeller speed reaches 106% of selected RPM (i.e., 2,000 selected RPM + 6% = 2,120 effective RPM).Propeller speeds in excess of 2,080 must be controlled through a combination of power and airspeed, with many other variables to be considered, including air density, prop blade angle, etc.

Powerplant



Data Summary

Powerplant System

POWER SOURCE Reverse flow, free turbine engines Pratt and Whitney PT6A-41 (200) Pratt and Whitney PT6A-42 (B200)

DISTRIBUTION Air from inlet screen to: Axial-flow compressor Centrifugal-flow compressor section Annular combustion chamber Hot, high-pressure gas from combustion chamber to: Single-stage, axial-flow turbine (to drive compressor and accessory section) Two-stage, axial-flow turbine (to drive power turbine shaft) Power turbine shaft drives Propeller Reduction gearbox

CONTROL Levers Power Propeller Condition (fuel control unit) Switches IGNITION AND ENGINE START (ON/OFF/STARTER ONLY) (L/R) ENG AUTO IGNITION (ARM/OFF) (L/R) ICE VANE (EXTEND/RETRACT) (L/R) PROP GOV (TEST/OFF)



MONITOR Engine Operation ITT Torque Prop RPM N1 RPM Indicators FUEL FLOW OIL TEMP OIL PRESSWarning annunciators ENG FIRE L/R FUEL PRESS L/R OIL PRESS L/R CHIP DETECT L/R Caution/advisory annunciators ICE VANE L/R (amber) ICE VANE EXT L/R (green) AUTOFEATHER L/R IGNITION ON (L/R)

Powerplant System (continued)

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