Super King Air B200GT & B200CGT Pilot's Operating · PDF fileModel B200GT/B200CGT 7-3...

Model B200GT/B200CGT

7-1

February, 2012

SECTION 7SYSTEMS DESCRIPTION

TABLE OF CONTENTSSUBJECT PAGE

Airframe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-7Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-7Seating Arrangements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-7

Flight Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-7Control Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-7Operating Mechanisms . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-7Manual Elevator Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-7Electric Elevator Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-8Rudder Boost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-8Yaw Damp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-9

Instrument Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-9Annunciator System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-9

Warning Panel Illustration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-11 Warning Panel Description. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-11 Caution/Advisory Panel Illustration . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-12 Caution/Advisory Panel Description. . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-12

Typical Illustrations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-14Instrument Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-14Control Wheels (With Clocks) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-15Control Wheels (Without Clocks) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-16Overhead Light Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-17Fuel Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-18Right Circuit Breaker Panel. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-19Pedestal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-20

Ground Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-21Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-21Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-21

Landing Gear Extension and Retraction . . . . . . . . . . . . . . . . . . . . . . . . . . 7-21Landing Gear Warning System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-23Manual Landing Gear Extension. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-23Brake System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-24Tires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-25

Baggage Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-25Seats, Seatbelts, and Shoulder Harnesses . . . . . . . . . . . . . . . . . . . . . . . . 7-26

Seats. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-26Cockpit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-26Lumbar Seats . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-26Cabin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-26Foyer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-27

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7-2

February, 2012



TABLE OF CONTENTS (CONT’D)SUBJECT PAGE

B200GT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-27B200CGT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-27

Aft-Cabin Area. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-27Seatbelts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-27Shoulder Harnesses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-27

Cockpit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-27Cabin. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-28Aft-Cabin Area. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-28

Doors, Windows And Exits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-29Airstair Entrance Door (B200GT). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-29Airstair Entrance Door (B200CGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-31Cargo Door (B200CGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-34Emergency Exit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-36

B200GT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-36B200CGT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-36All Serials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-36

Interior Doors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-36Cabin Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-37

Polarized Type . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-37Sun Visors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-37

Operating Instructions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-37Control Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-38Engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-39

Propulsion System Controls. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-40Power Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-40Propeller Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-40Condition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-40Propeller Ground Fine Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-40Propeller Reversing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-41Friction Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-41

Engine Instrumentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-42Propeller Synchrophaser . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-42Engine Lubrication System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-43

Magnetic Chip Detector. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-43Starting and Ignition System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-43

Auto Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-43Induction Air System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-44

Ice Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-44Engine Air Inlet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-44

Engine Anti-Ice System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-44



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Fuel Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-45Engine Compartment Fire Detection System . . . . . . . . . . . . . . . . . . . . . . 7-46Fire Extinguisher System (If Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-46

Propeller System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-47Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-47Low Pitch Stop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-47Propeller Governors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-48Autofeather System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-48

Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-49Fuel System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-50Fuel Pumps. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-51Auxiliary Tank Fuel Transfer System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-52Use Of Aviation Gasoline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-52Crossfeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-53Firewall Shutoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-53Fuel Routing In Engine Compartment. . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-53Fuel Drains . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-53Fuel Purge System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-54Fuel Gaging System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-54

Electrical System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-54Power Distribution Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-56Avionics/Electrical Equipment Bus Connection. . . . . . . . . . . . . . . . . . . . . 7-57External Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-61

Lighting Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-61Cockpit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-61Cabin. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-61

Emergency Exit Lighting System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-62Exterior . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-63

Environmental System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-63Pressurization System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-63

In-Flow System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-64Cabin Pressure Control System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-65Unpressurized Ventilation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-66B200GT Environmental System Schematic . . . . . . . . . . . . . . . . . . . . . 7-68B200CGT Environmental System Schematic. . . . . . . . . . . . . . . . . . . . 7-69Bleed Air Heating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-70

Electric Heating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-70Radiant Heating (B200CGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-71

Air Conditioning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-71


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Environmental Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-71Automatic Mode Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-71Manual Mode Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-72Electric Heat Mode Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-73Using Blowers For Air Recirculation . . . . . . . . . . . . . . . . . . . . . . . . . . 7-73Fault Diagnosis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-73

Oxygen System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-74Oxygen System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-74

Pitot and Static System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-76Pitot and Static System Schematic . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-78

Engine Bleed Air Pneumatic System . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-79Bleed Air Warning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-79

Stall Warning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-79Ice Protection Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-80

Windshield Heat. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-80Propeller Electric Deice System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-80

Propeller Electric Deice System Schematic . . . . . . . . . . . . . . . . . . . . . . 7-81Pitot Mast Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-82Surface Deice System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-82

Pneumatic Bleed Air System and Surface Deice System Schematic . . 7-83Stall Warning Vane Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-84Fuel Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-84Brake Deice System (If Installed). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-84

Comfort Features . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-85Side Facing Toilet (B200GT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-85Front Facing Toilet (B200CGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-85Relief Tubes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-85

Cabin Features . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-86Cabin AC Power Outlets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-86Fire Extinguishers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-86

Windshield Wipers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-87Cargo Restraint (B200CGT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-87Avionics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-88

Electronic Flight Instrument System (EFIS) . . . . . . . . . . . . . . . . . . . . . . . . 7-88Air Data System (ADS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-89AirCell ST 3100 Satellite Telecommunication System (If Installed) . . . . . . 7-90

Telephone Adapter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-91Basic Audio Panel Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-92



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Using The Telephone Adapter . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-93 To Make A Call In The Call Mode. . . . . . . . . . . . . . . . . . . . . . . . 7-93 To Make A Call In The Recall Mode. . . . . . . . . . . . . . . . . . . . . . 7-93 To Make A Call Using The Redial Key . . . . . . . . . . . . . . . . . . . . 7-94 To Receive A Call. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-94 Edit Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-95

AirCell Cordless Handset . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-97To Make A Call . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-97To Join A Call Made From The Cockpit. . . . . . . . . . . . . . . . . . . . . . 7-97To Receive A Call . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-97

Airspeed Scale . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-98Attitude Heading System (AHS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-100Control Display Unit (CDU) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-100Display Control Panels (DCP). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-101Emergency Locator Transmitter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-101Engine Indicating System (EIS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-102Enhanced Ground Proximity Warning System (EGPWS) (If Installed) . . 7-102Electronic Standby Instrument System (ESIS) . . . . . . . . . . . . . . . . . . . . 7-106Flight Guidance Panel (FGP) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-107Flight Guidance System (FGS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-107Flight Management System (FMS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-108Ground Communications Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-108HF-9000 System (If Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-109

RTU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-111CDU . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-113AVTECH CSD-714 SELCAL Decoder (If Installed) . . . . . . . . . . . . . . 7-115

IFIS-5000 Integrated Flight Information System . . . . . . . . . . . . . . . . . . . 7-115Radio Tuning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-116Terrain Awareness and Warning System Plus (TAWS+) (If Installed) . . 7-118Traffic Alert and Collision Avoidance System (TCAS I) . . . . . . . . . . . . . 7-124Traffic Alert and Collision Avoidance System (TCAS II) (If Installed) . . . 7-126

MFD Displays and Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-127PFD Displays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-129System Characteristics. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-130Voice Messages. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-130

Weather Radar System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-132Audio System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-132

Transmit (XMIT) Select Switch. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-132Transceiver and Receiver Audio Controls . . . . . . . . . . . . . . . . . . . . . 7-132Mic Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-133PA (Public Address) Audio Control . . . . . . . . . . . . . . . . . . . . . . . . . . 7-133


7-6

February, 2012




Auto Comm Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-133Speaker (SPKR) Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-133Interphone (INPH) Audio Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-133Alternate Audio Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-134Voice - Both - Ident Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7-134


Systems DescriptionSection 7


7-7

AIRFRAME STRUCTURE The Model B200GT/B200CGT is an all-metal, low-wing monoplane. It has fullycantilevered wings, and a T-tail empennage.

SEATING ARRANGEMENTS The pilot and copilot seats are mounted in a separate forward compartment. Var-ious configurations of passenger chairs and one-place couch installations may beinstalled on the continuous tracks mounted on the cabin floor. One or two fold-upseats may be installed in the aft cabin area. The toilet is also equipped for use asa seat. Seating for up to 15 persons, including crew, is available.

FLIGHT CONTROLS CONTROL SURFACES The airplane is equipped with conventional ailerons and rudder. It utilizes a T-tailhorizontal stabilizer and elevators, mounted at the extreme top of the vertical sta-bilizer.

OPERATING MECHANISMS The airplane is equipped with conventional dual controls for the pilot and copilot.The ailerons and elevators are operated by conventional control wheels intercon-nected by a T-bar. The rudder pedals are interconnected by linkage below thefloor. These systems are connected to the control surfaces through push-rod andcable-and-bellcrank systems. Rudder, elevator, and aileron trim are adjustablewith controls mounted on the center pedestal. A position indicator for each of thetrim tabs is integrated with its respective control.

MANUAL ELEVATOR TRIM Manual control of the elevator trim is accomplished with a trim control wheel locat-ed on the left side of the pedestal. It is a conventional trim wheel which is rolledforward for nose-down trim, and aft for nose-up trim.

October, 2007

101-590168-1A3_sec07.fm Page 7 Friday, February 10, 2012 1:59 PM

Section 7Systems Description Model B200GT/B200CGT

7-8

ELECTRIC ELEVATOR TRIM The electric elevator-trim system (installed in conjunction with the autopilot sys-tem) is controlled by a dual-element thumb switch on each control wheel, a trim-disconnect switch on each control wheel, and a PITCH TRIM circuit breaker in theFLIGHT group on the right side circuit breaker panel. Both elements of either dual-element thumb switch must be simultaneously moved forward to achieve nose-down trim, aft for nose-up trim; when released, they return to the center (OFF) po-sition. Any activation of the trim system by the copilot’s thumb switch can be over-ridden by the pilot’s thumb switch. No one switch element should activate thesystem; only the simultaneous movement of a pair of switch elements in the samedirection should activate the system.

A bi-level, push-button, momentary-on, trim-disconnect switch is located inboardof the dual-element thumb switch on the outboard grip of each control wheel. Theelectric elevator-trim system can be disconnected by depressing either of theseswitches. Depressing either trim-disconnect switch to the first of the two levels dis-connects the autopilot and yaw damp systems; depressing the switch to the sec-ond level disconnects the electric elevator-trim system. The manual-trim controlwheel can be used to change the trim anytime the autopilot is off, whether or notthe electric trim system is in the operative mode.

RUDDER BOOST A rudder boost system is provided to aid the pilot in maintaining directional controlin the event of an engine failure or a large variation of power between the engines.Incorporated into the rudder cable system are two pneumatic rudder-boosting ser-vos that actuate the cables to provide rudder pressure to help compensate forasymmetrical thrust.

During operation, a differential pressure valve accepts bleed air pressure fromeach engine. When the pressure varies between the bleed air systems, the shuttlein the differential pressure valve moves toward the low pressure side. As the pres-sure difference reaches a preset tolerance, a switch on the low pressure side clos-es, activating the rudder boost system. The system is designed only to helpcompensate for asymmetrical thrust. Appropriate trimming is to be accomplishedby the pilot. Moving either or both of the bleed air valve switches on the copilot’ssubpanel to the INSTR & ENVIR OFF position will disengage the rudder boostsystem.

The system is controlled by a toggle switch, placarded RUDDER BOOST - OFF,located on the pedestal. The switch is to be turned ON before flight. A preflightcheck of the system can be performed during the run-up by retarding the poweron one engine to idle and advancing power on the opposite engine until the powerdifference between the engines is great enough to close the switch that activatesthe rudder boost system. Movement of the appropriate rudder pedal (left engineidling, right rudder pedal moves forward) will be noted when the switch closes, in-dicating the system is functioning properly for low engine power on that side. Re-peat the check with opposite power settings to check for movement of theopposite rudder pedal.

February, 2012




7-9

YAW DAMPA yaw damp system is part of the autopilot and is provided to aid the pilot in main-taining directional control, and to increase ride comfort. The system may be usedat any altitude, and is required for flight above 17,000 feet. It should be deactivat-ed for takeoff and landing.

The yaw damper is automatically engaged anytime the autopilot is engaged, ex-cept when rudder boost is activated. To disengage the yaw damper, slide the YD/AP DISC switch down on the FGP panel, press the YD push-button, or press theDISC TRIM/AP YD switch on the pilot’s or copilot’s control wheel to the first detent.

INSTRUMENT PANEL The operation and use of the instruments, lights, switches, and controls locatedon the instrument panel are explained under the systems descriptions relating tothe subject items.

ANNUNCIATOR SYSTEM The annunciator system consists of a warning annunciator panel (red) centrally lo-cated in the glareshield, and a caution/advisory annunciator panel (caution - am-ber; advisory - green) located on the center subpanel. Two red MASTERWARNING flashers located in the glareshield (one in front of the pilot and one infront of the copilot) are a part of the system, as are two amber MASTER CAUTIONflashers (located just inboard of the MASTER WARNING flashers), and a PRESSTO TEST button located immediately to the right of the warning annunciator panel.

The annunciators are of the word-readout type. Whenever a condition monitoredby the annunciator system occurs, a signal is generated and the appropriate an-nunciator is illuminated.

If the fault requires the immediate attention and reaction of the pilot, the appropri-ate red warning annunciator in the warning annunciator panel illuminates and bothMASTER WARNING flashers begin flashing. Any illuminated lens in the warningannunciator panel will remain illuminated until the fault is corrected. However, theMASTER WARNING flashers can be extinguished by depressing the face of ei-ther MASTER WARNING flasher, even if the fault is not corrected. In such a case,the MASTER WARNING flashers will again be activated if an additional warningannunciator illuminates. When a warning fault is corrected, the affected warningannunciator will extinguish, but the MASTER WARNING flashers will continueflashing until one of them is depressed.

October, 2007

sec07.fm Page 9 Tuesday, August 7, 2007 9:01 AM


7-10

Whenever an annunciator-covered fault occurs that requires the pilot’s attentionbut not his immediate reaction, the appropriate amber caution annunciator in thecaution/advisory panel illuminates, and both MASTER CAUTION flashers beginflashing. The MASTER CAUTION annunciators can be extinguished by pressingthe face of either of the flashing annunciators to reset the circuit. Subsequently,when any caution annunciator illuminates, the MASTER CAUTION flashers willagain be activated. An illuminated caution annunciator on the caution/advisory an-nunciator panel will remain illuminated until the fault condition is corrected, atwhich time it will extinguish. The MASTER CAUTION flashers will continue flash-ing until one of them is depressed.

The caution/advisory annunciator panel also contains the green advisory annun-ciators. There are no master flashers associated with these annunciators, sincethey are only advisory in nature, indicating functional situations which do not de-mand the immediate attention or reaction of the pilot. An advisory annunciator canbe extinguished only by correcting the condition indicated on the illuminated lens.

The warning annunciators, caution annunciators, advisory annunciators, redMASTER WARNING and amber MASTER CAUTION flashers feature both a“bright” and a “dim” mode of illumination intensity. The “dim” mode will be selectedautomatically whenever all of the following conditions are met: a generator is onthe line; the OVERHEAD FLOOD LIGHTS are OFF; the PILOT INSTR PNL lightsare ON; and the ambient light level in the cockpit (as sensed by a photoelectriccell located in the overhead light control panel) is below a preset value. Unless allof these conditions are met, the “bright” mode will be selected automatically.

The lamps in the annunciator system should be tested before every flight, andanytime the integrity of a lamp is in question. Depressing the PRESS TO TESTbutton, located to the right of the warning annunciator panel in the glareshield, il-luminates all the annunciator lights, MASTER WARNING flashers, and MASTERCAUTION flashers. Any lamp that fails to illuminate when tested should be re-placed (refer to LAMP REPLACEMENT GUIDE in Section 8, HANDLING, SER-VICING AND MAINTENANCE).

October, 2007




7-11

WARNING PANEL ILLUSTRATION

WARNING PANEL DESCRIPTION

L ENG FIRE - - - DOOR UNLOCKED ALT WARN R ENG FIRE

L FUEL PRESS - - - - - - - - - R FUEL PRESS

L OIL PRESS - - - - - - - - - R OIL PRESS

- - - L BL AIR FAIL - - - R BL AIR FAIL - - -

NOMENCLATURE COLOR CAUSE FOR ILLUMINATION

L ENG FIRE Red Fire in left engine compartment.

DOOR UNLOCKED Red Cabin/cargo door open or not secure.

ALT WARN Red Cabin altitude exceeds 12,500 feet.

R ENG FIRE Red Fire in right engine compartment.

L FUEL PRESS Red Fuel pressure failure on left side.

R FUEL PRESS Red Fuel pressure failure on right side.

L OIL PRESS Red Left oil pressure is low.

R OIL PRESS Red Right oil pressure is low.

L BL AIR FAIL Red Melted or failed plastic left bleed air failurewarning line.

R BL AIR FAIL Red Melted or failed plastic right bleed air failurewarning line.

October, 2007



7-12

CAUTION/ADVISORY PANEL ILLUSTRATION

CAUTION/ADVISORY PANEL DESCRIPTION

L DC GEN - - - HYD FLUID LOW

RVS NOT READY

- - - R DC GEN

L CHIP DETECT

- - - - - - DUCT OVERTEMP

- - - R CHIP DETECT

L ENG ICE FAIL

L PITOT HEAT ELEC HEAT ON

EXT PWR R PITOT HEAT

R ENG ICE FAIL

L AUTO- FEATHER

- - - - - - AIR CNDN1 LOW

- - - R AUTO-FEATHER

L ENGANTI-ICE

*BRAKE DEICE ON

LDG/TAXI LIGHT

PASS OXYON

- - -

R ENGANTI-ICE

L IGNITION ON

L BL AIROFF

- - - FUELCROSSFEED

R BL AIROFF

R IGNITION ON

* Optional Equipment


L DC GEN Amber Left generator off the line.

HYD FLUID LOW Amber Hydraulic fluid in the landing gear system is low.

RVS NOT READY Amber Propeller levers are not in the high rpm, lowpitch position with landing gear extended.

R DC GEN Amber Right generator off the line.

L CHIP DETECT Amber Metal contamination in left engine oil isdetected.

DUCT OVERTEMP Amber Duct air too hot.

R CHIP DETECT Amber Metal contamination in right engine oil isdetected.

L ENG ICE FAIL Amber Left engine selected anti-ice system inoperative.

L PITOT HEAT Amber Left Pitot Heat Inoperative or switch is in theOFF position.

ELEC HEAT ON Amber Cabin electric heat on.

EXT PWR Amber External power connector is plugged in.

R PITOT HEAT Amber Right Pitot Heat Inoperative or switch is in theOFF position.

R ENG ICE FAIL Amber Right engine selected anti-ice system inoperative.

L AUTOFEATHER Green Autofeather armed with power levers advancedabove 90% N1.

AIR CND N1 LOW Green Right engine N1 too low for air conditioningload.

October, 2007




7-13

R AUTOFEATHER Green Autofeather armed with power levers advancedabove 90% N1.

L ENG ANTI-ICE Green Left engine anti-ice vanes in position for icing conditions.

*BRAKE DEICE ON Green Brake deice system in operation.

LDG/TAXI LIGHT Green Landing or taxi lights on with landing gear up.

PASS OXY ON Green Passenger oxygen system charged.

R ENG ANTI-ICE Green Right engine anti-ice vanes in position for icingconditions.

L IGNITION ON Green Left ignition and engine start switch is ON or leftauto ignition system is armed and left enginetorque is below 400 ft-lbs.

L BL AIR OFF Green Left environmental bleed air valve is closed.

FUEL CROSSFEED Green Crossfeed is selected.

R BL AIR OFF Green Right environmental bleed air valve is closed.

R IGNITION ON Green Right ignition and engine start switch is ON orright auto ignition system is armed and rightengine torque is below 400 ft-lbs.

* Optional Equipment


October, 2007



TYPICAL ILLUSTRATIONS

INSTRUMENT PANEL

LEFT

RIGHTLEFT

ENVIR

ENVIRONMENTAL

BLEED AIR

NORMAL

LOW

BLEED AIR VALVES

PNEU & ENVIR OFF

ENVIROFF

OFF

MAN COOL

AUTOHEATMAN

EXT

OFF

DETENG FIRE TEST

ELECHEAT

MAN TEMP

DECR

INCR

MODE

TUA

O

INCRINCRAUTOCABIN

INCR INCRAUTOCOCKPIT

TEMPBLOWER

BLOWER TEMP

WARNSTALL

TEST

E

COFF F

FO

E OFF

ONFURN

NO SMOKE& FSBBRIGHT

START/

DIM

OFF

FSBOFFCABIN LIGHTS

RIGHTOPEN

BARO

BB07C101388AA.AI

7-14 February, 2012



7-15

CONTROL WHEELS (WITH CLOCKS)

PILOT

COPILOT

BB07C101389AA.AI

February, 2012



7-16

CONTROL WHEELS (WITHOUT CLOCKS)

PILOT

COPILOT

BB07C101551AA.AI

February, 2012


Section 7R

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Model B200GT/B200CGT Systems Description

OVERHEAD LIGHT CONTROL PANEL

BB07C 070138AA.AI

00

10020

2030

108040 60

DC VOLTS

% LOAD 0

+60-60 BATT AMPS PROP AMPS

STALL WARNING IS INOPERATIVE WHEN MASTER SWITCH IS OFF. STANDBY COMPASS IS ERRATIC WHEN WINDSHIELD ANTI-ICE AND /OR AIR CONDITIONER

AND/OR HIGH VENT BLOWER AND/OR LANDING LIGHTS ARE ON.

OPERATION LIMITATIONSOFF

MASTER PANEL LIGHTS

ON

PILOT INSTR PNL

OFF

BRT

PILOT DISPLAYS

BRT

PARK SLOW

DO NOT OPERATE ON DRY GLASS

WINDSHIELD WIPER OFF

BRT BRT

FAST

OVERHEAD

FLOOD OFF

INSTRUMENT INDIRECT

OFF ANNUN PUSH-BRT

BRT BRT BRT BRT

COPILOT INSTR PNL

OFFCOPILOT DISPLAYS

SIDE PANEL

OFF

OVHD PED & SUBPANEL

OFF

DIM

THIS AIRPLANE MUST BE OPERATED AS A NORMAL CATEGORY AIRPLANE IN COMPLIANCE WITH THE OPERATING LIMITATIONS STATED IN THE FORM OF PLACARDS, MARKINGS AND MANUALS.

NO ACROBATIC MANEUVERS INCLUDING SPINS ARE APPROVED. THIS AIRPLANE APPROVED FOR VFR, IFR DAY & NIGHT OPERATION & IN ICING CONDITIONS.

CAUTION

00

10020

3010

8040 60

DC VOLTS

% LOAD AIRSPEED (IAS)

MAX GEAR EXTENSION MAX GEAR RETRACT MAX GEAR EXTENDED MAX APPROACH FLAP MAX FULL DOWN FLAP MAX MANEUVERING

181 KNOTS 163 KNOTS 181 KNOTS 200 KNOTS 157 KNOTS 181 KNOTS

7-17February, 2012


FUEL CONTROL PANEL

7-18 February, 2012

Section 7R

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RIGHT CIRCUIT BREAKER PANEL

BL07C 071321AA.AI

CMU

XM WX

FSU

CCP FSU

FAN

RADIO

SELCAL

DIALERTEL

COM ANT

PILOT

RIGHT

LEFT

OILPRESSWARN

HF

COPILOT

COPILOT

PILOT

DCPPFD

DCPPFD

AUDIOAUDIOCONT

AUDIOAUDIOCONT

NO.2

NO.1

AHCSEC

ATC

NO.2

NO.1

NAV

NO.2

NO.1

PFDHEATER

NO.3

AVIONICS

RTU

TCASRADAR

ALTM

RADIOMFD

IEC

INSTCOOLING

FLIGHT

EQPTCOOLING

NOSE

HTR

MFDTAWS

DBU

IAPS

RIGHT

LEFT

COPLT

PILOT

DME

NO.2

NO.1

CDU

NO.2

NO.1

GPS

NO.2

NO.1

AHC

NO.2

NO.1

ADC

NO.2

NO.1

COM

NO.2

NO.1

NO.2

NO.1

DCCONV

ESISHDG

SNSR

DISPBUS

CONT

LIGHTS

CDU

ESIS

PNL

INSTCONT

COPILOT

CONT

PED

INSTCONT

PILOT

MFD &

RTU

PFD &DCP

COPILOT

FGP PILOT

PFD &DCP

CHIPDETR

RIGHT

LEFT

RIGHT RIGHT

LEFTLEFTLEFT VOICE

RCDR

AUDIO

CABIN

AURAL

WARN

TRIM

PITCH

BRAKE

DEICE

AVIONICS

ANN

ENGINST

POWER

FEATHER

AUTO

NO.2

NO SMK

MN ENGANTI-ICE

RIGHT

STBY ENGANTI-ICE

DETR

FIRE

OUTSIDE

AIRTEMP

WEATHER

SYNC

PROP AVIONICS

MASTER

LIGHTER

CIGARMASTER

POWER

BUSFEEDERS

GENCONTROL

RIGHT

LEFT NO.1

NO.2

FGCSERVO

BLEEDAIR

CONTROLTEMP

CONTROLCONTROL

NO.1

RUDDER

BOOST

RIGHTCONTROL

LEFTPRESSOXYGEN

NO.1

NO.2

AVIONICS

AVIONICS

SURF

DEICE

WSHLD

WIPER

FUELVENT

RIGHT

LEFT

BLEEDAIR

WARNLANDING

GEARANN

IND

POWER

IND

WARN

RIGHT

LEFT

WARN

STALL

SUB PNLOVHD &INSTR

SIDE PNL

INSTR

READING

FSB &CABIN

CONSOLEINDIRECT

PLT FLT

COPLT FLT

AVIONICSENGINES

WARNINGS

FURNISHINGS

ELECTRICALFLIGHT

LIGHTS

ENVIRONMENTALCOM

NO.3

10

510

5

5

55

5 5

5

3

5

5

2

5

5

5

3

5

5

2

10 10

5

5 25 5 1

1

10

1 2

1 2

25

15

15

30

10

3

303 1

23 3

30

3

2

2

3

10

10

233

5

1

2

2

2

2

1

1

3

5

2

3

3

3

3

3

5 5

3 10 50 50

50 5010

5 5 5 5 5 5

53 5

5 5 5 5 5 5 10 55

5 55555

* OPTIONAL/IF INSTALLED

*

* *

*

***

***

*

* *

7-19February, 2012


BY07C 084360AA.AI

PEDESTAL DBU-5000

PEDESTAL DBU-4100

7-20 February, 2012



7-21

GROUND CONTROLDirect linkage from the rudder pedals allows for nose wheel steering. When therudder control is augmented by a main wheel brake, the nose wheel deflection canbe considerably increased.

The minimum wing-tip turning radius, using partial braking action and differentialengine power, is 39 feet 10 inches.

FLAPSThe flaps are operated by a sliding switch handle on the pedestal just below thecondition levers. Flap travel is registered on an electric indicator on top of the ped-estal. Three detents provide for quick selection of UP, APPROACH (or takeoff),and DN positions. The flaps cannot be stopped in an intermediate position. A safe-ty mechanism is provided to disconnect power to the electric flap motor in theevent of a malfunction which would cause any flap to be three to six degrees outof phase with the other flaps.

The flap-motor power circuit is protected by a 20-ampere flap-motor circuit break-er placarded FLAP MOTOR, located on the left circuit breaker panel below the fuelcontrol panel. A 5-ampere circuit breaker for the control circuit (placarded FLAPCONTROL) is also located on this panel.

LANDING GEARThe retractable tricycle landing gear is electrically controlled and hydraulically ac-tuated. The system utilizes folding braces, called drag legs, that lock in placewhen the gear is fully extended. The nose gear actuator incorporates an internalmechanical down-lock to hold the gear in the fully extended position. The maingear incorporates mechanical locks on the drag leg and no locks on the actuators.The landing gear is held in the up-lock position by hydraulic pressure.

Hydraulic pressure to the system is supplied by a hydraulic power pack. A hydrau-lic reservoir located in the left center wing section provides hydraulic fluid to thepower pack. The reservoir incorporates a dip stick to provide a visual check of fluidlevel.

Electrically actuated control valves control the flow of hydraulic fluid to the individ-ual gear actuators. The control valves receive electrical power through the landinggear control switch.

Accidental retraction of the landing gear is prevented through safety switches lo-cated on the main landing gears.

LANDING GEAR EXTENSION AND RETRACTION The nose and main landing gear assemblies are extended and retracted by a hy-draulic power pack in conjunction with hydraulic actuators. The hydraulic powerpack is located in the left center section, just forward of the main spar. One hy-draulic actuator is located at each landing gear. The power pack consists of: a hy-draulic pump, a 28-vdc motor, a two section fluid reservoir, filter screens, gear

February, 2012



7-22

selector valve, two solenoids, a fluid level sensor, and an up-lock pressure switch.For manual extension the system has a hand-lever-operated pump located on thefloor between the crew seats. Hydraulic lines, one for normal extension, and onefor retraction, routed from the power pack, and one for emergency extension fromthe hand pump, are routed to the nose and main gear actuators. The normal ex-tension lines and the manual extension lines are connected to the upper end ofeach hydraulic actuator. The hydraulic lines for retraction are fitted to the lowerends of the actuators. Hydraulic fluid under pressure generated by the power packpump and contained in the accumulator acts on the piston faces of the actuatorswhich are attached to folding drag braces resulting in the extension or retractionof the landing gear.

An internal mechanical lock in the nose gear actuator and the over-center actionof the nose gear drag leg assembly lock the nose gear in the down position.Notched hook, lock link and lock link guide attachments fitted to each main gearupper drag leg provide positive down-lock action for the main gear.

Electrical overload to the system is prevented through the use of a 60 ampere cir-cuit breaker located below the flooring.

The landing gear hydraulic power pack motor is controlled by the use of the land-ing gear control handle placarded LDG GEAR CONTROL - UP - DN located onthe pilot’s subpanel. The LDG GEAR CONTROL must be pulled out of a detentbefore it can be moved from either the UP or DN position.

Safety switches, called squat switches, on the main gear torque knees open thecontrol circuit when the strut is compressed. The squat switches must close to ac-tuate a solenoid which moves a down-lock hook on the LDG GEAR CONTROL tothe released position. This mechanism prevents the LDG GEAR CONTROL frombeing placed in the UP position when the airplane is on the ground. The hook au-tomatically disengages when the airplane leaves the ground, and can be overrid-den by pressing down on the red down-lock release button located to the left ofthe LDG GEAR CONTROL.

In flight, as the landing gear moves to the full down position, the down lock switch-es are actuated and interrupt current to the pump motor. When the red gear in-transit light in the LDG GEAR CONTROL extinguishes and the three green GEARDOWN indicators illuminate, the landing gear is in the fully extended position.

Two gear select solenoids located on the valve body of the pump are energizedthrough positioning of the LDG GEAR CONTROL either to the UP or DN position.Once energized, the gear select valve is actuated, allowing hydraulic fluid to flowto the actuators.

Hydraulic system pressure performs the up-lock function, holding the landing gearin the retracted position. When the hydraulic pressure reaches 2775 ± 55 psi, theup-lock pressure switch will cause the landing gear relay to open and interrupt thecurrent to the pump motor. The same pressure switch will cause the pump to ac-tuate, should the hydraulic pressure drop to approximately 2400 psi.

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A caution annunciator, placarded HYD FLUID LOW, in the caution/advisory an-nunciator panel will illuminate (amber) whenever the hydraulic fluid level in the hy-draulic power pack is low. The annunciator is tested by pressing in the HYD FLUIDSENSOR TEST button located on the pilot’s subpanel.

The LDG GEAR CONTROL should never be moved out of the DN detent while theairplane is on the ground. If it is, the landing gear warning horn will sound intermit-tently, and the red gear-in-transit lights in the LDG GEAR CONTROL will illumi-nate (provided the MASTER SWITCH is ON), warning the pilot to return thehandle to the DN position.

Landing gear position is indicated by an assembly of three green annunciators.When illuminated, the annunciators indicate that a particular gear is DN. Absenceof illumination indicates gear UP.

Two red parallel-wired indicator lights, located in the LDG GEAR CONTROL, illu-minate to show that the gear is in transit or unlocked. The red lights in the handlealso illuminate when the landing gear warning horn is actuated.

The red lights may be checked by pressing the HDL LT TEST button located ad-jacent to the LDG GEAR CONTROL.

LANDING GEAR WARNING SYSTEM The landing gear warning system is provided to warn the pilot that the landing gearis not down and locked during specific flight regimes. Various warning modes re-sult, depending upon the position of the flaps.

With the FLAPS in either the UP or APPROACH position and either or both powerlevers retarded below approximately 80% N1, the warning horn will sound inter-mittently and the LDG GEAR CONTROL lights will illuminate. The horn can be si-lenced by pressing the GEAR HORN SILENCE button located on the left powerlever. The lights in the LDG GEAR CONTROL cannot be extinguished. The land-ing gear warning system will be rearmed if the power lever(s) are advanced suffi-ciently.

With the FLAPS beyond APPROACH position, the warning horn and LDG GEARCONTROL lights will be activated regardless of the power settings, and neithercan be cancelled.

MANUAL LANDING GEAR EXTENSION An alternate extension handle, placarded LANDING GEAR ALTERNATE EXTEN-SION, is located on the pilot’s side of the pedestal. To engage the system, pull theLANDING GEAR RELAY circuit breaker, located to the left of the LDG GEARCONTROL on the pilot’s right subpanel, and ensure that the LDG GEAR CON-TROL is in the DN position. Remove the alternate extension handle from the se-curing clip and pump up and down. While pumping, do not lower the handle belowthe level of the securing clip during the down stroke as this will allow accumulatedhydraulic pressure to bleed off. Continue the pumping action until the three greengear-down annunciators are illuminated, and further resistance is felt, then stow

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the handle in the securing clip. If one or more gear-down annunciators do not illu-minate, the alternate extension handle must not be stowed. Instead, leave it at thetop of the up stroke. Continue to pump the handle when conditions permit until thegear is mechanically secured after landing. Refer to LANDING GEAR MANUALEXTENSION in Section 3A, ABNORMAL PROCEDURES. If any of the followingconditions exist, it is likely that an unsafe gear indication is due to an unsafe gearand is not a false indication.

1. The inoperative gear down annunciator illuminates when tested.2. The red lights in the handle are illuminated.3. The gear warning horn sounds when one or both power levers are retarded

below a preset N1.

After a practice manual extension of the landing gear, the gear may be retractedhydraulically. Refer to LANDING GEAR RETRACTION AFTER PRACTICE MAN-UAL EXTENSION in Section 4, NORMAL PROCEDURES.

BRAKE SYSTEM The dual hydraulic brakes are operated by depressing the toe portion of either thepilot’s or copilot’s rudder pedals. The series system plumbing enables braking byeither pilot or copilot.

Dual parking-brake valves are installed adjacent to the rudder pedals between themaster cylinders of the pilot’s rudder pedals and the wheel brakes. A control forthe valves, placarded PARKING BRAKE, is located below the pilot’s left subpanel.After the pilot’s brake pedals have been depressed to build up pressure in thebrake lines, both valves can be closed simultaneously by pulling out the parkingbrake handle. This retains the pressure in the brake lines. The parking brake isreleased by depressing the pedals briefly to equalize the pressure on both sidesof the valve, then pushing in the parking brake handle to open the valve.

The parking brake should be left off and wheel chocks installed if theairplane is to be left unattended. Changes in the ambient tempera-ture can cause the brakes to release or to exert excessive pres-sures.

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TIRES The airplane is normally equipped with dual 18x5.5 Type VII, 8-ply-rated, tubeless,rim-inflated tires on each main gear. For increased service life, 10-ply-rated tiresof the same size may be installed.

Optionally, the airplane may be equipped with dual 22x6.75-10, 8-ply-rated, tube-less tires on each main gear. These tires provide higher flotation, and permit op-eration at approximately 2/3 the inflation pressure required for the standard 18x5.5tires.

The nose gear is equipped with a 22x6.75-10, 8-ply-rated, tubeless tire.

BAGGAGE COMPARTMENT The entire aft-cabin area (which is aft of the foyer) may be utilized as a baggagecompartment. A nylon web is provided for the restraining of loose items. See “Di-mensional and Loading Data” and “Cabin Arrangement Diagrams” in Section 6,WEIGHT AND BALANCE/EQUIPMENT LIST.

Unless authorized by applicable Department of Transportation Reg-ulations, do not carry hazardous material anywhere in the airplane.

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

Baggage and other objects should be secured by webs in order toprevent shifting in turbulent air.

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SEATS, SEATBELTS AND SHOULDER HARNESSES SEATS COCKPIT

The pilot and copilot seats are adjustable fore, aft, and vertically by the use of re-lease levers located beneath the seats. The angle of the seat is adjustable by de-pressing the release lever on the side of the seat. The armrests incorporate bothangular adjustment and vertical stowing. To stow the armrest, lift the adjustmentlever and rotate the armrest to the vertical position.

LUMBAR SEATS

The firmness of the lower seat back may be controlled by utilizing a button locatedon the lower inboard side of the seat back. After adjusting the seat back to a com-fortable position, move forward on the seat to remove all the weight from the seatback. Hold the button in until the support fully inflates, release the button and leanback in the seat. If the support is too firm, hold the button in until the desired de-gree of firmness is obtained.

CABIN

Various configurations of passenger chairs and a one-place couch may be in-stalled on the continuous tracks which are mounted on the cabin floor. All passen-ger chairs are placarded either FRONT FACING ONLY or FRONT OR AFTFACING on the horizontal leg cross brace. Only chairs placarded FRONT OR AFTFACING may be installed facing aft. All aft-facing chairs (and all forward-facingchairs that are equipped with shoulder harnesses) are equipped with adjustableheadrests.

Before takeoff and landing, lateral tracking seats should be in theoutboard position, all seat backs positioned upright, and all head-rests fully extended.

Some passenger chairs can be moved fore and aft by lifting a release lever underthe front of adjustable seats. (“Front” is the direction opposite the seat back, re-gardless of whether the chair faces fore or aft.)

The seat backs can be adjusted to any angle from fully upright to fully reclining, bydepressing the release lever located on the side of the seat at the front inboardcorner. When the lever is depressed and the passenger leans against the seatback, the seat back will slowly recline until the lever is released, or until the fullyreclining position is attained. When no weight is placed against the seat back andthe lever is depressed, the seat back will rise until the lever is released, or until thefully upright position is reached. The seat backs of all occupied seats must be up-right for takeoff and landing.

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The optional lateral-tracking passenger chairs incorporate a release lever under-neath the front inboard corner of the seats. When this lever is lifted, the chairs canbe adjusted laterally and/or fore and aft. The seat back adjustments are the sameas those on the standard passenger chairs. These seats must be in the outboardposition (i.e., against the cabin wall) for takeoff and landing.

Inboard armrests on passenger chairs can be folded flush with the top of the seatcushions to facilitate entry to and egress from the seat. The armrests can be low-ered by lifting the flat, rectangular release plate located under the front end of thearmrest, then moving the armrest toward the front of the seat and downward. Thearmrest can be raised by pulling the armrest upward and toward the seat back un-til it locks into place.

The couch is not adjustable.

FOYER

B200GT

A hinged seat-cushion mounted on top of the toilet forms an extra passenger seatwhen the toilet is not is use.

B200CGT

A forward facing cabin chair is available for an extra passenger seat.

AFT-CABIN AREA

One or two optional folding seats may be installed in the aft-cabin area. They aremounted on the cabin sidewall and swing inboard when unfolded. A latch mecha-nism on the leg locks the seats in place when they are unfolded. When this seatingis not needed, the seat(s) may be folded against the cabin sidewall and held inplace with retaining straps.

A hinged seat-cushion mounted on top of the toilet forms an extra passenger seatwhen the toilet is not in use.

SEATBELTS Every seat occupiable for takeoff and landing is equipped with a seatbelt. All oc-cupants must wear seatbelts during takeoff and landing. The fore and aft facingseats also have shoulder harnesses.

SHOULDER HARNESSES COCKPIT

The shoulder harness installation for the pilot and copilot seats consists of a Y-strap mounted to an inertia reel located in the lower seat back. One strap is wornover each shoulder and terminates with a fitting which attaches to the lap buckle.When the seat is equipped with a rotary buckle, rotation of the release 1/8 of a turnclockwise will simultaneously release the shoulder harness and lap belt.

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The shoulder harness straps proceed from inertia reels built into the crew chairs.Spring loading at the inertia reels keeps the shoulder harnesses snug, but allowsthe pilot and copilot all the freedom of movement normally required in flight. How-ever, the inertia reels incorporate a locking device that will secure the harnessstraps in the event of sudden forward movement.

CABIN

The fore and aft facing seats are equipped with inertia reel shoulder harnesses.

The shoulder harness is worn diagonally and runs from the shoulder to the hiparea where it is secured by hooking the metal fastener around the securing studon the male half of the seatbelt buckle.

AFT-CABIN AREA

The shoulder harness for the aft-cabin area fold-up chairs is of a double-strap con-figuration. The middle portion of the strap is secured by a metal slip ring which isanchored to the aft pressure bulkhead. The two ends (which actually function astwo separate straps) extend downward toward the seatbelt-buckle area. One endof the shoulder harness strap terminates in a slotted bayonet-blade fastener. Theother end is attached to the upper edge of the shoulder harness adjuster. A shortadjusting strap, which is also equipped with a slotted bayonet blade fastener, ex-tends upward from the area of the seatbelt buckle and slides through the lowerportion of the shoulder harness adjuster. A small, flexible adjusting tab is also at-tached to the lower edge of the adjuster.

One shoulder harness strap is worn over each shoulder. When the two bayonetblades are placed together, the shoulder harness straps can be secured by slidingthe male half of the seatbelt buckle through the bayonet slots and into the femalehalf of the seatbelt buckle. The shoulder harness strap can be lengthened bygrasping the tab on the adjuster and pulling upward. The strap can be tightenedby grasping the loose end of the adjusting strap and pulling it through the adjusteruntil the shoulder harness is snug.

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DOORS, WINDOWS AND EXITS AIRSTAIR ENTRANCE DOOR (B200GT) The airstair entrance door (cabin door) is hinged at the bottom. It swings outwardand downward when opened. A stairway is built onto the inboard side of the door.Two of the stairsteps fold flat against the door when the door is closed. A hydraulicdamper ensures that the door will swing down slowly when it opens. While thedoor is open, it is supported by plastic-encased cables, which also serve as hand-rails. Additionally, these cables are utilized when closing the door from inside theairplane. An inflatable rubber seal is installed around the perimeter of the door,and seats against the door frame as the door is closed. When weight is off thelanding gear, engine bleed air supplies pressure to inflate the door seal, whichprovides a positive pressure-vessel seal around the door. The outside door handlecan be locked with a key, for security of the airplane on the ground.

AIRSTAIR ENTRANCE DOOR (B200GT)

Only one person should be on the airstair door stairway at any onetime.

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The door locking mechanism is operated by rotating either the outside or the in-side door handle. The handles are linked together, so they move together. Twolatch bolts at each side of the door, and two latch hooks at the top of the door, lockinto the door frame to secure the airstair door.

Whether unlocking the door from the outside or the inside, the release button ad-jacent to the door handle must be held depressed before the handle can be rotat-ed (counterclockwise from inside the airplane, clockwise from outside) to unlockthe door. Consequently, unlocking the door is a two-hand operation requiring de-liberate action. The release button acts as a safety device to help prevent acciden-tal opening of the door. As an additional safety measure, a differential-pressure-sensitive diaphragm is incorporated into the release-button mechanism. The out-board side of the diaphragm is open to atmospheric pressure, the inboard side tocabin air pressure. As the cabin-to-atmospheric pressure differential increases, itbecomes increasingly difficult to depress the release button, because the dia-phragm moves inboard when either the outboard or inside release button is de-pressed.

Never attempt to unlock or even check the security of the door inflight.

If the DOOR UNLOCKED annunciator illuminates in flight, or if thepilot has any reason to suspect that the door may not be securelylocked, the cabin differential pressure should be reduced to the low-est practical value, and all occupants instructed to remain seatedwith their seatbelts fastened. After the airplane has made a full-stoplanding and the cabin has been depressurized, a crew membershould check the security of the cabin door.

To close the door from outside the airplane, lift up the free end of the airstair doorand push it up against the door frame as far as possible. Then grasp the handlewith one hand and rotate it clockwise as far as it will go. The door will then moveinto the closed position. Then rotate the handle counterclockwise as far as it willgo. The release button should pop out, and the handle should be pointing aft.Check the security of the airstair door by attempting to rotate the handle clockwisewithout depressing the release button; the handle should not move.

To close the door from inside the airplane, grasp the handrail cable and pull theairstair door up against the door frame. Then grasp the handle with one hand androtate it counterclockwise as far as it will go, continuing to pull inward on the door.The door will then move into the closed position. Then turn the handle clockwiseas far as it will go. The handle should be pointing down. Check the security of thedoor by attempting to rotate the handle counterclockwise without depressing therelease button; the handle should not move. Next, lift the folded stairstep that isjust below the door handle. A placard adjacent to the round observation windowadvises the observer that the safety lock arm should be in position around the di-aphragm shaft (plunger) when the handle is in the locked position. The placard

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also presents a diagram showing how the arm and shaft should be positioned. Ared push-button switch near the window turns on a lamp inside the door, which il-luminates the area observable through the window. If the arm is properly posi-tioned around the shaft, proceed to check the indication in each of the visualinspection ports, one of which is located near each corner of the door. The greenstripe painted on the latch bolt should be aligned with the black pointer in the visualinspection port. If any condition specified in this door-locking procedure is not met,DO NOT TAKE OFF.

AIRSTAIR ENTRANCE DOOR (B200CGT) The airstair door is built into the cargo door. It is hinged at the bottom, and swingsdownward when opened. It has a stairway built onto the inboard side. Two of thestairsteps fold flat against the door when the door is closed. When the door isopened, a self-storing platform automatically folds down over the door sill to pro-tect the rubber door seal. A hydraulic damper ensures that the door will swingdown slowly when it opens. While the door is open, it is supported by a plastic-encased cable, which also serves as a handrail. Additionally, this cable is utilizedwhen closing the door from inside the airplane. An inflatable rubber seal is in-stalled around the perimeter of the door, and seats against the door frame as thedoor is closed. When the cabin is pressurized, air seeps into the rubber sealthrough small holes in the outboard side of the seal. The higher the cabin differ-ential pressure, the more the seal inflates. This is a passive-seal system with nomechanical connection to a bleed air source. The outside door handle can belocked with a key, for security of the airplane on the ground.

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AIRSTAIR ENTRANCE DOOR (B200CGT)

Only one person should be on the airstair door stairway at any onetime.

The door locking mechanism is operated by rotating either the outside or the in-side door handle, both of which move simultaneously. Three hollow, crescentlatches on each side of the door rotate to capture or release latch posts mountedin the cargo door to secure the airstair door. When latched, the airstair door be-comes an integral part of the cargo door.

Whether unlocking the door from the outside or the inside, the release button ad-jacent to the door handle must be held depressed before the handle can be rotat-ed to unlock the door (counterclockwise from inside the airplane, clockwise fromoutside). Consequently, unlocking the door is a two-hand operation requiring de-liberate action. The release button acts as a safety device to help prevent acciden-tal opening of the door. As an additional safety measure, a differential-pressure-sensitive diaphragm is incorporated into the release-button mechanism. The out-board side of the diaphragm is open to atmospheric pressure, the inboard side to

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cabin air pressure. As the cabin-to-atmospheric pressure differential increases, itbecomes increasingly difficult to depress the release button, because the dia-phragm moves inboard when either the outboard or inside release button is de-pressed.

Never attempt to unlock or even check the security of the door inflight.

If the DOOR UNLOCKED annunciator illuminates in flight, or if thepilot has any reason whatever to suspect that the door may not besecurely locked, the cabin should be depressurized (after first con-sidering altitude), and all occupants instructed to remain seated withtheir seatbelts fastened. After the airplane has made a full-stop land-ing and the cabin has been depressurized, only a crew membershould check the security of the airstair door and the cargo door.

To close the door from outside the airplane, lift up the free end of the airstair doorand push it up against the door frame as far as possible. Then grasp the handlewith one hand and rotate it clockwise as far as it will go. The door will then moveinto the closed position. Then rotate the handle counterclockwise as far as it willgo. The release button should pop out, and the handle should be pointing aft.Check the security of the door by attempting to rotate the handle clockwise withoutdepressing the release button; the handle should not move.

To close the door from inside the airplane, grasp the handrail cable and pull theairstair door up against the door frame. Then grasp the handle with one hand androtate it counterclockwise as far as it will go, continuing to pull inward on the door.The door will then move into the closed position. Then turn the handle clockwiseas far as it will go. The release button should pop out, and the handle should bepointing down. Check the security of the door by attempting to rotate the handlecounterclockwise without depressing the release button; the handle should notmove. Next, lift the second folded stair step below the door handle. A placard ad-jacent to the round observation window advises the observer that the safety lockarm should be in position around the diaphragm shaft (plunger) when the handleis in the locked position. The placard also presents a diagram showing how thearm and shaft should be positioned. A red push-button switch near the windowturns on a lamp inside the door, which illuminates the area observable through thewindow. If the arm is properly positioned around the shaft, proceed to check theorange stripe on each of the six rotary latches (three on each side of the airstairdoor). Ensure that each is aligned with the notch in the plate on the door frame.Finally, turn the battery switch ON and check the warning annunciator panel in thecockpit; ensure that the red DOOR UNLOCKED annunciator is extinguished. Withthe battery switch ON, it will be illuminated if the airstair door is open. With the bat-tery switch OFF, it will be illuminated only if the airstair door is closed but not se-curely latched. Perform the “Cabin Door Annunciator Circuitry Check” in Section4, NORMAL PROCEDURES, prior to the first flight of the day. If any conditionspecified in this door-latching procedures is not met, DO NOT TAKE OFF.

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CARGO DOOR (B200CGT) A large, swing-up cargo door, hinged at the top, provides access for the loadingof large items. The cargo door latch system is operated by two handles: one in theupper aft area of the door, and the other in the lower forward area of the door. Twoseparate access covers must be opened in order to operate the two handles. Inorder to move the upper aft handle out of the latched position, depress the blackrelease button in the handle and rotate the yellow handle upward as far as it willgo. This movement is transmitted via cables to two hollow, crescent latches on theforward side and two on the aft side of the cargo door. The latches rotate to re-lease latch posts mounted in the cargo door frame.

CARGO DOOR (B200CGT)

In order to move the lower latch handle out of the CLOSED position (forward), liftthe orange lock hook from the stud on the yellow latch handle, and rotate the han-dle aft as far as it will go. This movement is transmitted via linkage to four latchpins on the bottom of the cargo door. The pins move aft to disengage latch lugsmounted at the bottom of the cargo door frame.

To open the cargo door after it is unlatched, push out on the bottom of the door.After the cargo door is manually opened a few feet, gas springs take over andraise the door to the fully open position.

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To close the cargo door, pull it down and inboard. The gas springs will resist theclosing effort until the door is only open a few feet. Then, as the springs move overcenter, they begin applying a closing force to the door.

An inflatable rubber seal is installed around the perimeter of the cargo door, andseats against the door frame when closed. When the cabin is pressurized, airseeps into the rubber seal through small holes in the outboard side of the seal.The higher the cabin differential pressure, the more the seal inflates. This is a pas-sive-seal system and has no mechanical connection to a bleed air source.

There are no latch handles on the outside of the cargo door, so it can be openedand closed from inside the airplane only.

To latch the cargo door after it is closed, rotate the lower forward latch handle for-ward until the orange lock hook engages the stud on the handle. Check the secu-rity of this handle by attempting to move it aft without raising the lock hook; itshould not move. Close the access cover. Next, check the observation window atthe lower aft corner of the cargo door. Ensure that the orange stripe on the latchpin linkage is aligned with the orange pointer in the observation window.

Next, rotate the upper aft latch handle down until the black release button popsup. Check the security of this handle by attempting to pull it out and up without de-pressing the release button; it should not move. Close the access cover. Then, en-sure that the orange stripe on each of the four rotary latches (two on each side ofthe cargo door) is aligned with the notch in the plate on the door frame. Finally,check the caution annunciator panel in the cockpit and ensure that the amberDOOR UNLOCKED annunciator is extinguished. With the battery switch ON orOFF, it will be illuminated if either the airstair door or the cargo door is open. Per-form the “Cabin/Cargo Door Annunciator Circuitry Check” in Section 4, NORMALPROCEDURES, prior to the first flight of the day. If any condition specified in thisdoor-latching procedures is not met, DO NOT TAKE OFF.

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EMERGENCY EXIT The emergency exit door is located on the right side of the fuselage at the forwardend of the passenger compartment. From the inside, the door is released with apull-down handle, placarded EXIT-PULL. From the outside, the door is releasedwith a flush-mounted, pull-out handle. The nonhinged, plug-type door removescompletely from the frame into the cabin when the latches are released.

B200GTThe door can be locked so that it cannot be removed or opened from the outsideusing the flush-mounted pull-out handle. The door is locked when the lock-lever(inside) is in the down or locked position. Locking the door is for security when theairplane is parked. The lock-lever should be in the up or unlocked position prior toflight, 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 using the EXIT-PULLhandle, since this handle is not locked by the lock-lever. An exit lock placard isplaced on the lock-lever so that it can be read when the lever is in the locked po-sition.

B200CGT

The door can be locked with a key from the inside, to prevent opening from theoutside. The inside handle will unlatch the door, whether or not it is locked, byoverriding the locking mechanism. The key lock should be unlocked prior to flight,to allow removal of the door from the outside in the event of an emergency. Thekeyhole is in the horizontal position when the door is locked. The key cannot beremoved in this position.

ALL SERIALS

A wiper-type disconnect for the air duct that supplies air to the eyeball outlet in theemergency exit door is located on the upper-aft edge of the door. As the door isremoved, the duct is disconnected, since it is an integral part of the door.

Located on the lower-forward edge of the door is an electrical disconnect for thewiring that goes to the lights in the emergency exit door. It will unplug as the dooris being removed. Upon reinstalling the door, the electrical disconnect should bereconnected before moving the door into the closed position.

INTERIOR DOORS Sliding doors are provided between the cockpit and cabin, and between the cabinand foyer. These doors provide privacy, and prevent the spilling of light from onecompartment into another. The doors are closed by sliding the two partition-typedoor panels to the center of the aisle, where they are held together by a magneticstrip in the edge of each door.

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CABIN WINDOWS Each cabin window pane, which is composed of a sheet of polyvinyl butyral (PVB)laminated between two sheets of clear acrylic plastic, is stressed to withstand thecabin-to-ambient air pressure differential. It is then sealed into a window openingin the fuselage, and forms an integral part of the pressure vessel.

POLARIZED TYPE

Two dust panes are mounted inboard of the cabin window pane in each windowframe. Each of these dust panes is composed of a film of polarizing material lam-inated between two sheets of acrylic plastic. The inboard dust pane rotates freelyin the window frame and has a protruding thumb knob near the edge. Rotating thepane through an arc of 90 permits complete light regulation as desired. Rotationchanges the relative alignment between the polarizing films, thus providing anydegree of light transmission from full intensity to almost none.

Do not look directly at the sun, even through polarized windows, be-cause eye damage could result.

When the airplane is to be parked in areas exposed to intense sun-light, the polarized windows should be rotated to the clear positionto prevent deterioration of the polarization coating. Sufficient ultravi-olet protection is provided to prevent fading of the upholstery.

SUN VISORSOPERATING INSTRUCTIONS To Operate From Stowed Position:

Push straight back and pull down. Move along track to desired place and pivot outnear windshield (or window), rotate knob clockwise to lock.

To Change Position:

Rotate knob counterclockwise to unlock, move to desired location and position,then relock knob by turning clockwise.

To Stow:

Rotate knob counterclockwise to unlock, move along track to aft end, pivot upagainst headliner to allow catch to retain sun visor assembly.

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CONTROL LOCKS Install the control locks in the following sequence:

1. Position the U-clamp around the engine control levers. 2. Move the control column as necessary to align the holes, then insert the L-

shaped pin that is attached to the middle of the chain (approx.). The controlwheel position should be full forward and rotated approximately 15 to theleft.

3. Insert the L-shaped pin (attached to the end of the chain) through the holeprovided in the floor aft of the rudder pedals. The rudder pedals must becentered to align the hole in the rudder bellcrank with the hole in the floor.The pin is then inserted until the flange is resting against the floor. This willprevent any rudder movement.

Before starting engines, remove the control locks, reversing theabove procedure.

Remove the control locks before towing the airplane. If towed with atug while the rudder lock is installed, serious damage to the steeringlinkage can result.

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ENGINES

PT6A-52 ENGINE

The Model B200GT/B200CGT is powered by two Pratt & Whitney Canada Corp.PT6A-52 turbo-propeller engines, each rated at 850 SHP. Each engine has athree-stage axial flow, single-stage centrifugal-flow compressor, which is drivenby a single-stage reaction turbine. The power turbine - a two-stage reaction tur-bine counter-rotating with the compressor turbine - drives the output shaft. Boththe compressor turbine and the power turbine are located in the approximate cen-ter of the engine, with their shafts extending in opposite directions. Being a re-verse flow engine, the ram air supply enters the lower portion of the nacelle andis drawn in through the aft protective screens. The air is then routed into the com-pressor. After it is compressed, it is forced into the annular combustion chamber,and mixed with fuel that is sprayed in through 14 nozzles mounted around the gasgenerator case. A capacitance discharge ignition unit and two spark igniter plugsare used to start combustion. After combustion, the exhaust passes through thecompressor turbine and two stages of power turbine and is routed through two ex-haust ports near the front of the engine. A pneumatic fuel control system sched-ules fuel flow to maintain the power set by the gas generator power lever.Propeller speed within the governing range remains constant at any selected pro-peller control lever position through the action of a propeller governor, except inthe beta range where the maximum propeller speed is controlled by the pneumaticsection of the propeller governor.

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The accessory drive at the aft end of the engine provides power to drive the fuelpumps, fuel control, the oil pumps, the refrigerant compressor (right engine), thestarter/generator, and the tachometer generator. At this point, the speed of thedrive (N1) is the true speed of the compressor side of the engine, 37,468 rpm(which corresponds to 100% N1). Maximum continuous speed of the engine is39,000 rpm, which equals 104% N1.

The reduction gearbox forward of the power turbine provides gearing for the pro-peller and drives the propeller tachometer generator, the propeller overspeed gov-ernor, and the propeller governor. The turbine speed on the power side of theengine is 30,145 rpm. After reduction gearing the propeller rpm is 2000.

PROPULSION SYSTEM CONTROLS The propulsion system is operated by three sets of controls; the power levers, pro-peller levers, and condition levers. The power levers serve to control engine pow-er. The condition levers control the flow of fuel at the fuel control outlet and selectfuel cutoff, low idle and high idle functions. The propeller levers are operated con-ventionally and control the constant speed propellers through the primary gover-nor.

POWER LEVERS

The power levers provide control of engine power from idle through take-off powerby operation of the gas generator (N1) governor in the fuel control unit. IncreasingN1 rpm results in increased engine power.

PROPELLER LEVERS

Each propeller lever operates a speeder spring inside the primary governor to re-position the pilot valve, which results in an increase or decrease of propeller rpm.For propeller feathering, each propeller lever lifts the pilot valve to a position whichcauses complete dumping of high pressure oil, allowing the counterweights andfeathering spring to change the pitch. Detents at the rear of lever travel preventinadvertent movement into the feathering range. Operating range is 1600 to 2000rpm.

CONDITION LEVERS

The condition levers have three positions; FUEL CUT OFF, LOW IDLE and HIGHIDLE. Each lever controls the fuel cutoff function of the fuel control unit and limitsidle speed at approximately 61% N1 or as necessary to maintain 1180 propellerRPM for low idle, and 70% N1 for high idle.

PROPELLER GROUND FINE OPERATION

Propeller ground fine is used to provide deceleration on the ground during landingwithout propeller reversing by taking advantage of the maximum available propel-ler drag.

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Ground fine operation is accomplished by a gate position for the power levers onthe quadrant. The power levers must be retarded below the IDLE gate by raisingthem over the gate and retarding the levers to the GROUND FINE gate.

Power levers should not be moved to the GROUND FINE positionwhen the engines are not running as this will cause damage to thesystem.

PROPELLER REVERSING

When the power levers are lifted over the IDLE gate, they control engine powerthrough the GROUND FINE and REVERSE ranges. To operate in the REVERSErange, the power levers must be raised over the GROUND FINE gate and movedaft to the REVERSE position.

Propeller reversing on unimproved surfaces should be accom-plished carefully to prevent propeller erosion from reversed airflowand, in dusty or snowy conditions, to prevent obscuring the opera-tor’s vision.

Condition levers, when set at HIGH IDLE, keep the engines operating at 70% N1high idle speed for maximum reversing performance.

Power levers should not be moved into the reversing position whenthe engines are not running because the reversing system will bedamaged.

FRICTION LOCKS

Four friction locks are located on the power quadrant of the pedestal. When theyare rotated counterclockwise, the propulsion system control levers can be movedfreely. As the friction locks are rotated clockwise, the control levers progressivelybecome more resistant to movement, so that they will not creep out of position.

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ENGINE INSTRUMENTATIONEngine instrument design has incorporated currently approved green, yellow andred operating ranges and monitor time limits of each range to give the pilot a visualindication of engine parameters as they change.

Engine instruments are provided on the MFD at the top of the display. In the eventof an MFD failure, the selection of the reversionary switch to PFD will place theengine instruments at the top of the pilot and copilot PFD. • The ITT indicators• The torquemeters • The N2 (propeller) tachometers • The N1 (gas generator) tachometers • The fuel flow indicators • The oil temperature/oil pressure indicators

A propeller synchroscope is located in the upper right corner of the MFD, directlybelow the oil temperature digits. The synchroscope will translate to the right if theright engine is running faster and like wise to the left if the left engine is runningfaster. A series of non-moving white squares should be visible if the propellers arein sync.

PROPELLER SYNCHROPHASER The propeller synchrophaser system is an electronic system certified for all oper-ations including takeoff and landing. The system automatically matches the RPMof both propellers and positions the propellers at a preset phase relationship in or-der to reduce cabin noise.

Before engaging the system, manually set the RPM of each engine to within 10RPM of each other. When the prop sync switch is turned on, engagement will au-tomatically occur when the relative phase angle of the propellers is within 30 ofthe preset angle. When the system engages, both propeller speeds are increasedby one-half the holding range of the system. To maintain synchronization, the sys-tem increases the RPM of the slower propeller and simultaneously reduces theRPM of the faster propeller. The system will never reduce RPM below that select-ed by the propeller control lever.

To change RPM with the system ON, adjust both propeller controls by the sameamount. If the synchrophaser is ON but does not maintain synchronization, thesystem has reached the end of its range. Increasing the setting of the slow propel-ler, or reducing the setting of the fast propeller, will bring the speeds within the lim-ited synchrophaser range. If preferred, the synchrophaser switch may be turnedoff, the propellers re-synchronized manually, and the synchrophaser turned backon.

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ENGINE LUBRICATION SYSTEM Engine oil, contained in an integral tank between the engine air intake and the ac-cessory case, cools as well as lubricates the engine. An oil cooler located insidethe lower nacelle, keeps the engine oil temperature within the operating limits. Athermal element is used to regulate a bypass door which controls the volume ofcooling air through the cooler. Engine oil also operates the propeller pitch changemechanism and the engine torquemeter system.

The lubrication system capacity per engine is 3.8 U.S. gallons (15.25 U.S. quarts).The oil tank capacity is 2.5 gallons (10 U.S. quarts) with 5 quarts measured on thedipstick for adding purposes. See Pratt & Whitney Service Bulletin Number 13001for approved engine oils.

MAGNETIC CHIP DETECTOR

A magnetic chip detector is installed in the bottom of each reduction gearbox. Thisdetector will activate an amber annunciator, L CHIP DETECT or R CHIP DETECT,to alert the pilot of possible metal contamination in the engine oil supply. Illumina-tion of a CHIP DETECT annunciator is not in itself cause for an engine to be shutdown. Engine instruments should be monitored for abnormal indications. If param-eters are abnormal, a precautionary shutdown may be made at the pilot’s discre-tion. After illumination of a CHIP DETECT annunciator, cause of the malfunctionshould be determined and corrected prior to the next flight.

STARTING AND IGNITION SYSTEM Each engine is started by a three-position switch located on the pilot’s left subpan-el placarded, IGNITION AND ENGINE START - LEFT - RIGHT - ON - OFF -STARTER ONLY. Each switch may be moved downward to the STARTER ONLYposition to motor the engine for the purpose of clearing it of fuel without the ignitioncircuit on. The switch is spring loaded and will return to the center position whenreleased. Moving the switch upward to the ON position activates both the starterand ignition, and the appropriate green IGNITION ON annunciator will illuminate.The starter drive action is stopped by placing the switch in the center OFF posi-tion.

AUTO IGNITION

The auto ignition system provides automatic ignition to prevent engine loss due tocombustion failure. This system is provided to ensure ignition during takeoff, land-ing, turbulence, and penetration of icing or precipitation conditions. To arm thesystem, move the required ENG AUTO IGNITION switches, located on the pilot’ssubpanel, from OFF to ARM. If for any reason the engine torque falls below ap-proximately 460 foot-pounds, the igniter will automatically energize and the IGNI-TION ON annunciator on the caution/advisory annunciator panel will illuminate.For extended ground operation, the system should be turned off to prolong the lifeof the igniter units.

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INDUCTION AIR SYSTEM The PT6A-52 is a reverse-airflow engine. The compressor wheels draw ambientair into the engine through the induction air inlet at the lower front of the enginenacelle. As airspeed increases, ram air pressure rises, compressing the air insidethe induction air duct. The air then flows into an annular inlet-air chamber locatedat the aft end of the engine compartment. It then passes through a protectivescreen and into the primary compressor impeller, where it is further compressed.Then the air is forced through a stator ring and successively through the secondand third axial-flow compressor stages. It is finally compressed in the centrifugal-flow compressor stage, then discharged into the turbine plenum assembly. Airfrom the plenum enters the annular combustion chamber through a series of holesin the aft end of the combustion chamber, and mixes with fuel that is sprayed intothe combustion chamber through 14 nozzles mounted around the gas generatorcase. The air-fuel mixture burns inside the combustion chamber, then the hot gas-ses expand forward out of the chamber and pass through the compressor turbinestage, both stages of the power turbine, and out to the atmosphere through twoexhaust ports located at the side of each nacelle, near the front.

ICE PROTECTION

ENGINE AIR INLET

Engine exhaust heat is utilized for heating the engine air inlet lips. A small tube isinstalled through the bottom of the left engine exhaust stack. The tube is insertedfacing into the exhaust flow which forces a portion of exhaust gases through a flexhose to the engine air inlet anti-ice lip. The hot gases are circulated through theengine air inlet anti-ice lip (hot lip) to prevent the formation of ice. The gases arethen exhausted through a flex hose to the right exhaust stack. No shutoff or tem-perature indicator is necessary for this system.

ENGINE ANTI-ICE SYSTEM

An inertial separation system is built into each engine air inlet to prevent moistureparticles from entering the engine inlet plenum under icing conditions. The systemincludes dual actuators and controls. The system is monitored by L and R ENGANTI-ICE (green) and L and R ENG ANTI-ICE (amber) annunciators. Illuminationof the L and R ENG ANTI-ICE (green) annunciators indicates that the system isactuated. Illumination of the L and R ENG ANTI-ICE (amber) annunciators indi-cates that the system did not operate to the desired position. Immediate illumina-tion of the L or R ENG ANTI-ICE (amber) annunciator indicates loss of electricalpower, whereas delayed illumination indicates an inoperative actuator. In eitherevent, the STANDBY actuator should be selected.

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FUEL CONTROL The engine fuel system consists of an engine-driven fuel pump, a fuel control unit(FCU), a flow divider and dump valve, dual fuel manifold, fourteen fuel nozzles,and two fuel drain valves. The fuel pump/fuel control unit assembly is mounted onthe engine accessory case and is shaft-driven at a speed proportional to that ofthe compressor turbine. The fuel pump delivers fuel to the FCU. Engine poweroutput is established by power lever position. The power lever is linked to the gov-ernor in the FCU which regulates fuel flow to the combustion section and therebycontrols N1 and power output. Increasing N1 rpm results in increased engine pow-er. System function depends upon the interaction of the fuel control unit governorand the propeller governor. The position of the fuel control unit metering valve isdetermined by differential pressures that vary proportionately with power required(as sensed by the fuel control unit) and propeller rpm.

The flow divider directs fuel from the metering valve to the primary and secondaryfuel manifolds (or primary manifold only, depending on engine power require-ments) and thence to the fuel nozzles. The flow divider also incorporates a dumpvalve that automatically drains residual fuel from both manifolds at engine shut-down. The fuel drain valves drain fuel from the combustion chamber at engineshutdown and at engine false starts. A fuel purge system has been installed in thiscase and on engine shutdown when fuel manifold pressure subsides, allows thedump poppet valve to open, the purge tank pressure forces fuel out of the enginefuel manifold lines through the nozzles and into the combustion chamber wherethe fuel is consumed. Constant fuel pressure is maintained by a fuel filter bypassvalve and a pressure relief valve.

The oil-to-fuel heater mounted below the fuel pump on the accessory case is aheat exchanger that transfers heat from the engine lubricating oil to preheat thefuel. A fuel temperature-sensing oil bypass valve regulates the fuel temperatureby either allowing oil to flow through the heater circuit, or bypass it to the engineoil tank.

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ENGINE COMPARTMENT FIRE DETECTION SYSTEM The fire detection system is designed to provide an immediate warning in theevent of fire in either engine compartment. The system consists of a temperaturesensing cable for each engine; two red warning annunciators, L ENG FIRE and RENG FIRE (located in the warning annunciator panel); two test switches on the co-pilot’s left subpanel, and a circuit breaker labeled FIRE DETR located on the rightcircuit breaker panel. When the fire has been extinguished (if the integrity of thedetector system has not been destroyed), the system will reset itself.

The test switches, on the copilot’s left subpanel, are placarded ENG FIRE TEST:DET, OFF, EXT and LEFT and RIGHT. When either of the switches is placed inthe DET position, the corresponding ENG FIRE annunciator along with flashingMASTER WARNING annunciator will illuminate and a red annunciation of FIRE isvisible in the ITT/TORQUE indicator for either engine on the MFD. The systemmay be tested anytime, either on the ground or in flight.

FIRE EXTINGUISHER SYSTEM (IF INSTALLED) The optional Engine Fire Extinguisher system consists of the following items foreach engine; a fire extinguisher cylinder, spray bars to disperse the extinguishingagent, a pyrotechnic cartridge, a control switch incorporating system status an-nunciators, and a test switch.

The fire extinguisher control switches used to activate the left and right extinguish-er systems are located on the glareshield at each end of the warning annunciatorpanel. Their power is derived from the hot battery bus. Each push-to-activateswitch incorporates three annunciators.

ENG FIRE-PUSH TO EXT (Red) - Illuminates in conjunction with the red L orR ENG FIRE annunciator located on the warning panel and the red FIRE an-nunciator located on each ITT/Torque display to indicate an engine fire hasbeen detected. The annunciator also illuminates when the respective DETTest Switch has been selected.DISCH (Amber) - Illuminates when the respective fire extinguisher cylinderhas been discharged and when the respective EXT Test Switch has been se-lected.OK (Green) - Illuminates when the respective EXT Test Switch has been se-lected if the circuit of the fire extinguisher pyrotechnic cartridge is complete.

To discharge a fire extinguisher cylinder, raise the safety-wired clear plastic coveron the control switch and press the face of the annunciator. The pyrotechnic car-tridge will fire, pressurizing the fire extinguisher cylinder. The extinguishing agentwill then be dispersed to the spray bars in the engine compartment. When the cyl-inder is empty, the amber DISCH annunciator will illuminate. The DISCH annun-ciator will remain illuminated, even with power removed from the airplane, until thepyrotechnic cartridge has been replaced.

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The fire extinguisher system is tested prior to the first flight of the day using theTEST SWITCH - ENG FIRE DET SYS; DET - R - L; EXT - R - L. When EXT R orL is selected, the respective amber DISCH and green OK annunciators will illumi-nate on the Fire Extinguisher Control switch.

The fire extinguisher cylinder for each engine is located in the respective wheelwell. A gage, calibrated in psi, is located on each cylinder for determining the levelof charge. A table showing the appropriate charge versus ambient temperature isfound in the Preflight Inspection section of the Hawker Beechcraft Corporation/Flight Safety abbreviated checklist.

PROPELLER SYSTEM DESCRIPTION Each engine is equipped with a conventional four blade, full-feathering, constant-speed, counter-weighted, reversing, variable-pitch propeller mounted on the out-put shaft of the reduction gearbox. The propeller pitch and speed are controlledby engine oil pressure, through single-action, engine-driven propeller governors.Centrifugal counterweights, assisted by a feathering spring, move the blades to-ward the low rpm (high pitch) position and into the feathered position. Governorboosted engine oil pressure moves the propeller to the high rpm (low pitch) hy-draulic stop and reverse position. The propellers have no low rpm (high pitch)stops; this allows the blades to feather after engine shutdown.

LOW PITCH STOP Low pitch propeller position is determined by the Low Pitch Stop which is a me-chanically actuated hydraulic stop. This mechanism allows the blades to rotate be-yond the low pitch position into ground fine and reverse when selected duringground operation. Beta and reverse blade angles are controlled through the powerlevers in the Beta and reverse range by displacing the governor beta valves.

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PROPELLER GOVERNORS Two governors, a constant speed governor and an overspeed governor, controlthe propeller rpm. The constant speed governor, mounted on top of the reductiongearbox, controls the propeller through its entire range. The propeller control leveroperates the propeller by means of this governor. If the constant speed governorshould malfunction by requesting more than 2000 rpm, the overspeed governorcuts in at 2120 rpm and dumps oil from the propeller to keep the rpm from exceed-ing approximately 2120 rpm. A solenoid actuated by the PROP GOV - TESTswitch located on the pilot’s subpanel, is provided for resetting the overspeed gov-ernor to approximately 1830 to 1910 rpm for test purposes.

If the propeller sticks or moves too slowly during a transient condition causing thepropeller governor to act too slowly to prevent an overspeed condition, the powerturbine governor, contained within the constant speed governor housing, acts asa fuel topping governor. When the propeller reaches 2120 rpm, the fuel toppinggovernor limits the fuel flow to the gas generator, reducing N1 rpm, which in turnprevents the propeller rpm from exceeding approximately 2200 rpm. During oper-ation in the reverse range, the fuel topping governor is reset to approximately 95%propeller rpm before the propeller reaches a negative pitch angle. This ensuresthat the engine power is limited to maintain a propeller rpm somewhat less thanthat of the constant speed governor setting. The constant speed governor there-fore will always sense an underspeed condition and direct oil pressure to the pro-peller servo piston to permit operation in Beta and reverse ranges.

AUTOFEATHER SYSTEMThe automatic feathering system provides a means of immediately dumping oilfrom the propeller servo to enable the feathering spring and counterweights to rap-idly feather the propeller in the event of an engine failure. The system is armedusing a switch on the pilot’s subpanel placarded AUTOFEATHER - ARM - OFF -TEST. With the switch in the ARM position and both power levers above approx-imately 90% N1, a green AFX is displayed in the ITT/TORQUE indicator on theMFD, and a green annunciator, placarded (L) and (R) AUTOFEATHER on theCaution/Advisory annunciator panel will illuminate, indicating the system is armed.If either power lever is not above approximately 90% N1, the system will be dis-armed and neither annunciator will be illuminated. When the system is armed andthe torque on a failing engine drops below approximately 410 ft-lbs, the autofeath-er system of the operative engine is disarmed causing its annunciators to extin-guish. When the torque on the failing engine drops below approximately 260 ft-lbs,the oil is dumped from the servo, the feathering spring and counterweights featherthe propeller, and the annunciators for the failed engine extinguish.

The system may be tested on the ground using the spring-loaded TEST positionof the switch. With the switch in the TEST position, the 90% N1 switches are dis-abled and the system will arm with the power levers set at approximately 500 ft-lbs of torque. Retarding a single power lever will then simulate an engine failureand the resulting action of the autofeather system can be checked as described

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in Section 4, NORMAL PROCEDURES. Since an engine is not actually shut downduring a test, the AUTOFEATHER annunciator for the engine being tested will cy-cle on and off as the torque oscillates above and below the 260 ft-lb setting.

FUEL SYSTEM The fuel system consists of two separate systems connected by a valve-controlledcrossfeed line. The fuel system for each engine is further divided into a main andauxiliary fuel system. The main system consists of a nacelle tank, two wing lead-ing edge tanks, two box section bladder tanks, and an integral (wet cell) tank, allinterconnected to flow into the nacelle tank by gravity. This system of tanks is filledfrom the filler located near the wing tip.

The auxiliary fuel system consists of a center section tank with its own filler open-ing, and an automatic fuel transfer system to transfer the fuel into the main fuelsystem.

When the auxiliary tanks are filled, they will be used first. During transfer of auxil-iary fuel, which is automatically controlled, the nacelle tanks are maintained full. Aswing check valve in the gravity feed line from the outboard wing prevents reversefuel flow. Upon exhaustion of the auxiliary fuel, normal gravity transfer of the mainwing fuel into the nacelle tanks will begin.

An anti-siphon valve is installed in each filler port which prevents loss of fuel orcollapse of a fuel cell bladder in the event of improper securing or loss of the fillercap.

The two systems are vented through a recessed ram vent coupled to a protrudingheated ram vent on the underside of the wing adjacent to the nacelle. One vent isrecessed to prevent icing and the protruding vent, added as a backup, is heatedto prevent icing.

All fuel is filtered with a firewall-mounted 20-micron filter. These filters incorporatean internal bypass which opens to permit uninterrupted fuel supply to the enginein the event of filter icing or blockage. In addition, a screen strainer is located ateach tank outlet before the fuel reaches the boost and transfer pumps. The mainengine driven fuel pump has an integral strainer to protect the pump.

A “differential pressure” fuel purge system is provided and is located in the aftcompartment of each nacelle. The system purges the fuel that is left in the fuelmanifolds at engine shutdown by forcing the fuel into the nozzles so that it is con-sumed in the combustion chamber.

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BB

07C005795A

B.A

I

AVIATION FUEL FILLERPROBESSUCTION RELIEF VALVECHECK VALVEFUEL FLOW INDICATORFUEL PRESSURE ANNUNCIATOR

FUEL AT STRAINER OR FILTERFUEL UNDER PUMP PRESSUREFUEL CROSSFEEDFUEL RETURNFUEL PURGEFUEL VENT

LEGEND

F

L

L

ENGINE FUEL MANIFOLD

FUEL CONTROL UNIT

PRESSURE TANKFIREWALL FUEL FILTER

ENGINE-DRIVENBOOST PUMPDRAIN VALVE

GRAVITY FEEDLINE DRAIN

FIREWALLSHUTOFF VALVE

STANDBY BOOST PUMP

VENT FLOAT VALVENACELLE TANKCROSSFEED VALVE

STRAINER, DRAIN& FUEL SWITCH

AUXILIARYF

TRANSFER JET PUMP

FUEL SYSTEM SCHEMATIC

RECESSED VENTHEATED RAM VENT

FLAME ARRESTORDRAIN

AIR INLET

VENT FLOAT VALVE

F

W.S. 290.92

INTEGRAL (WET CELL)

WING LEADING EDGEWING LEADING EDGE

BOX SECTIONBOX

SECTION

DRAINVALVE

AIR FILTER

FUEL HEATERENGINE-DRIVEN FUEL PUMP

P BLEED AIR LINEFUEL FLOW TRANSMITTER AND INDICATOR

LEFT FUEL PRESSURE ANNUNCIATORPRESSURE SWITCH

GRAVITY FLOW CHECK VALVEFUEL CONTROL PURGE VALVE

TRANSFER CONTROL MOTIVE FLOW VALVE

STRAINER AND DEFUELING DRAIN VALVE

PRESSURE SWITCH FOR LEFT NO FUELTRANSFER LIGHT ON FUEL PANEL

3

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FUEL PUMPS The engine-driven fuel pump (high pressure) is mounted on the accessory casein conjunction with the fuel control unit. Failure of this pump results in an immedi-ate flameout. The primary boost pump (low pressure) is also engine driven and ismounted on a drive pad on the aft accessory section of the engine. This pump op-erates when the gas generator (N1) is turning and provides sufficient fuel for start,takeoff, all flight conditions (except operation with hot aviation gasoline above20,000 feet altitude) and operation with crossfeed.

In the event of a primary boost pump failure, the respective red FUEL PRESS an-nunciator will illuminate. This annunciator illuminates when pressure decreasesbelow 10 ± 1 psi. The annunciator will extinguish by switching on the standby fuelpump on that side, thus increasing pressure above 11 ± 2 psi.

Engine operation with a fuel pressure annunciator on is limited to 10hours between overhaul, or replacement, of the engine driven fuelpump.

When using aviation gasoline during climbs above 20,000 feet, the first indicationof insufficient fuel pressure will be an intermittent flicker of the FUEL PRESS an-nunciators. A wide fluctuation of the fuel flow indicator may also be noted. Theseconditions can be eliminated by turning on a standby pump.

An electrically driven standby boost pump (low pressure), located in the bottom ofeach nacelle tank, performs three functions; it is a backup pump for use in theevent of a primary fuel boost pump failure, it is for use with hot aviation gasolineabove 20,000 feet, and it is used during crossfeed operations. In the event of aninoperative standby pump, crossfeed can only be accomplished from the side ofthe operative pump.

Electrical power to operate the standby boost pumps is controlled by lever locktoggle switches, placarded STANDBY PUMP - ON - OFF, located on the fuel con-trol panel. It is supplied power from the number 3 or number 4 feeder bus, and isprotected by a 10-ampere circuit breaker located on the fuel control panel. Thispower is only available when the master switch is turned on. These circuits areprotected by diodes to prevent the failure of one circuit from disabling the othercircuit.

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AUXILIARY FUEL TRANSFER SYSTEM The auxiliary tank fuel transfer system automatically transfers the fuel from theauxiliary tank to the nacelle tank without pilot action. Motive flow to a jet pumpmounted in the auxiliary tank sump is obtained from the engine fuel plumbing sys-tem downstream from the engine-driven boost pump and routed through the trans-fer control motive flow valve. The motive flow valve is energized to the openposition by the control system to transfer auxiliary fuel to the nacelle tank to beconsumed by the engine during the initial portion of the flight. When an engine isstarted, pressure at the engine driven boost pump closes a pressure switch which,after a 30 to 50 second time delay to avoid depletion of fuel pressure during start-ing, energizes the motive flow valve. When the auxiliary fuel is depleted, a low lev-el float switch de-energizes the motive flow valve after a 30 to 60 second timedelay provided to prevent cycling of the motive flow valve due to sloshing fuel.

In the event of a failure of the motive flow valve or the associated control circuitry,the loss of motive flow pressure when there is still fuel remaining in the auxiliaryfuel tank is sensed by a pressure switch and float switch, respectively, which illu-minates a light placarded NO TRANSFER on the fuel control panel. During enginestart, the pilot should note that the NO TRANSFER lights extinguish 30 to 50 sec-onds after engine start. A manual override is incorporated as a backup for the au-tomatic transfer system. This is initiated by placing the AUX TRANSFER switch,located in the fuel control panel to the OVERRIDE position.

USE OF AVIATION GASOLINE If aviation gasoline must be used as an emergency fuel, it will be necessary to de-termine how many hours the engines are operated on gasoline. Since the gasolineis being mixed with the regular fuel, it is expedient to record the number of gallonsof gasoline taken aboard for each engine. Each engine is permitted 150 hours ofoperation on aviation gasoline between overhauls. This means that if one enginehas an average fuel consumption of 50 gallons per hour, for example, it is allowed7500 gallons of aviation gasoline between overhauls. (Two engines; 15,000 gal-lons between overhauls.)

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CROSSFEED During emergency single-engine operation, it may become necessary to supplyfuel to the operative engine from the fuel system on the opposite side. The simpli-fied crossfeed system is placarded for fuel selection with a diagram on the upperfuel control panel. Place the standby fuel pump switches in the OFF position whencrossfeeding. A lever-lock switch, placarded CROSSFEED FLOW, is moved fromthe center OFF position to the left or to the right, depending on direction of fuelflow. This opens the crossfeed valve, energizing the standby pump on the sidefrom which crossfeed is desired, and de-energizes the motive flow valve in the fuelsystem on the side being fed. When the crossfeed mode is energized, a greenFUEL CROSSFEED annunciator on the caution/advisory panel will illuminate.

FIREWALL SHUTOFF The system incorporates two firewall shutoff valves controlled by two switches,one on each side of the fuel system circuit breaker panel, located on the fuel con-trol panel. These switches, respectively LEFT and RIGHT, are placarded FIRE-WALL SHUTOFF VALVE - OPEN - CLOSED. A red guard over each switch is anaid in preventing inadvertent operation. The firewall shutoff valves receive electri-cal power from the main buses and also from the hot battery bus which is connect-ed directly to the battery.

FUEL ROUTING IN ENGINE COMPARTMENT Just forward of the firewall shutoff valve is the primary engine-driven boost pump.From the primary boost pump, the fuel is routed to the main fuel filter, the fuel flowindicator transmitter, through a fuel heater that utilizes heat from the engine oil towarm the fuel, through the engine driven fuel pump, then to the fuel control unit.From there it is directed through the dual fuel manifold to the fuel outlet nozzlesand into the annular combustion chamber. Fuel is also taken from just down-stream of the main fuel filter to supply the jet transfer pump motive flow.

FUEL DRAINS During each preflight, the fuel drains on the tanks, pumps and filters should bedrained to check for fuel contamination. There are five sump drains and one filterdrain in each wing. They are located as follows:

DRAINS LOCATION

Leading Edge Tank Outboard of nacelle underside of wingIntegral Tank Underside of wing forward of aileronFirewall Fuel Filter Underside of cowling forward of firewallSump Strainer Bottom center of nacelle forward of wheel wellGravity Feed Line Aft of wheel wellAuxiliary Tank At wing root just forward of the flap

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FUEL PURGE SYSTEM Engine compressor discharge air (P3 air) pressurizes a small purge tank. Duringengine shutdown, fuel manifold pressure subsides, thus allowing the engine fuelmanifold poppet valve to open. The purge tank pressure forces fuel out of the en-gine fuel manifold lines, through the nozzles, and into the combustion chamber.As the fuel is burned, a momentary surge in (N1) gas generator rpm should be ob-served. The entire operation is automatic and requires no input from the crew.

During engine starting, fuel manifold pressure closes the fuel manifold poppetvalve, allowing P3 air to pressurize the purge tank.

FUEL GAGING SYSTEM The airplane is equipped with a capacitance type fuel quantity indication system.A maximum indication error of 3% full scale may be encountered in the system.The system is designed for the use of Jet A, Jet A1, JP-5 and JP-8 aviation kero-sene, and compensates for changes in fuel density due to temperature changes.If other fuels are used, the system will not indicate correctly. See OTHER PRO-CEDURES in Section 4 for instructions when using Jet B, JP-4, or aviation gaso-line.

The LEFT fuel quantity indicator on the fuel control panel indicates the amount offuel remaining in the left-side main fuel system tanks when the fuel QUANTITYSELECT switch is in the MAIN (upper) position, and the amount of fuel remainingin the left-side auxiliary fuel tank when the fuel QUANTITY SELECT switch is inthe AUXILIARY (lower) position. The RIGHT fuel quantity indicator indicates thesame information for the right-side fuel systems, depending upon the position ofthe FUEL QUANTITY switch. The gages are marked in pounds.

ELECTRICAL SYSTEM The airplane electrical system is a 28-vdc (nominal) system with the negative leadof each power source grounded to the main airplane structure. DC electrical pow-er is provided by one 42-ampere-hour, sealed lead acid battery, and two 250-am-pere starter/generators connected in parallel. The system is capable of supplyingpower to all subsystems that are necessary for normal operation of the airplane.A hot battery bus is provided for emergency operation of certain essential equip-ment and the cabin entry threshold light circuit. Power to the main battery bus fromthe battery is routed through the battery relay which is controlled by a switch plac-arded BAT - ON - OFF, located on the pilot’s subpanel. Power to the bus systemfrom the generators is routed through reverse-current-protection circuitry. Re-verse current protection prevents the generators from absorbing power from thebus when the generator voltage is less than the bus voltage. The generators arecontrolled by switches, placarded GEN 1 and GEN 2, located on the pilot’s sub-panel.

Starter power to each individual starter/generator is provided from the main bat-tery bus through a starter relay. The start cycle is controlled by a three-positionswitch for each engine, placarded IGNITION AND ENGINE START - ON - OFF -STARTER ONLY, on the pilot’s subpanel. The starter/generator drives the com-

February, 2012




7-55

pressor section of the engine through the accessory gearing. The starter/genera-tor initially draws approximately 1100 amperes, then drops rapidly to about 300amperes as the engine reaches 20% N1.

Power is supplied from three sources: the battery, the right generator, and the leftgenerator. The generator buses are interconnected by two 325-ampere currentlimiters. The entire bus system operates as a single bus, with power being sup-plied by the battery and both generators. There are four dual-fed sub-buses. Eachsub-bus is supplied power from either generator main bus through a 60-amp lim-iter, a 70-amp diode, and a 50-amp circuit breaker. Electrical loads are dividedamong the buses as noted on the accompanying Power Distribution Schematic.The equipment on the buses is arranged so that all items with duplicate functions(such as right and left landing lights) are connected to different buses.

Individual control switches are provided in the pilot’s subpanel and are placardedGEN 1 and GEN 2 - OFF/ON/GEN RESET. The generators are self-excited anddo not require battery power for operation. To bring a generator on line, thegenerator switch should be momentarily placed in the GEN RESET position, thenreleased to ON. In the GEN RESET position, the generator voltage builds up to28 volts and the line contactor is open. When the generator switch is released toON, the line contactor is allowed to close.

Generator control units provide voltage regulation, differential voltage, reversecurrent protection, paralleling, cross-start current limiting, and overvoltage protec-tion control for the generators. The voltage regulation circuit controls the generatorto maintain a constant 28-volt output. The differential voltage circuit compares thegenerator output and the center bus voltages then closes the line contactor if thegenerator is within acceptable limits of the center bus voltage.

Reverse current protection circuitry opens the line contactor and disconnects thegenerator if a reverse current condition occurs. If the condition corrects itself, theline contactor will re-close automatically.

The paralleling circuit provides load equalization between both generators. Thecross-start current limiting circuit limits the generator output during engine cross-start operation.

The overvoltage protection circuit senses the generator output voltage and de-ex-cites the generator and opens the line contactor if an overvoltage occurs. If thegenerator is disconnected for overvoltage, it will be necessary to select GEN RE-SET, then ON to reset the generator.

L DC GEN and R DC GEN annunciators are provided. Illumination of the L or RDC GEN annunciator indicates that the line contactor is open and the generator isoff line. Loadmeters in the overhead instrument panel indicate the load applied toeach generator as a percent of generator rating.

Battery current can be monitored with the battery ammeter located in the over-head instrument panel. To meet the battery duration time noted in the Dual Gen-erator Failure Emergency procedure, the battery charge current must be 10 ampsor less prior to takeoff. Takeoff with a battery charge current above 10 amps is per-mitted at the discretion of the pilot.

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POWER DISTRIBUTION SCHEMATIC

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AVIONICS/ELECTRICAL EQUIPMENT BUS CONNECTIONThe following table shows the equipment powered by each sub-bus, organized by system. Each sub-bus, in turn, is powered by the Left and/or Right Generator Bus or the battery.Most of the equipment listed below is protected by a circuit breaker located on either the Right circuit Breaker Panel (no notation), the Fuel Control Panel (noted by FP), or a circuitbreaker switch located on the Pilot’s Subpanel (noted by SW). Equipment shown on the Hot Battery Bus is protected by devices located in the right wing (RW) next to the battery, andthe Landing Gear Relay circuit breaker is located on the Pilot’s Right Subpanel. See additional notes which follow the table.

System No. 1 Dual-Fed Bus

No. 2 Dual-Fed Bus

No. 3 Dual-Fed Bus

No. 4 Dual-Fed Bus

Avionics Bus No. 1

(L Gen Bus)

Avionics Bus No. 2

(R Gen Bus)

Avionics Bus No. 3

(L Gen Bus)

No. 1 Subpanel Bus(L Gen Bus)

No. 2Subpanel Bus(R Gen Bus)

ESIS Battery Bus(No. 3 Dual-

Fed Bus)

Hot Bat Bus(Battery)

(RW)

Avionics AURAL WARN IAPS, L IAPS, R ADC NO. 1 ADC NO. 2 AVIONICSAVIONICS MASTER

DBU AHC NO. 1 (1) AHC NO. 2 (1) GND COMM

AHC SEC NO. 2 (Secondary

Power)

AHC SEC NO. 1 (Secondary

Power)

CDU NO. 2 (opt)

CDU NO. 1

ESIS BAT CHG (FP)

PILOT DCP COPILOT DCP

MFD PILOT PFD COPILOT PFD RADIO ALTM RTU IEC PILOT PFD

HEATERMFD HEATER COPILOT PFD

HEATER

DC CONV NO. 2 (2)

DC CONV NO. 1 (2)

FLIGHT INST COOLING

NOSE EQPT COOLING

VOICE RCDR TAWS NAV NO. 1 NAV NO. 2 TCAS COM NO. 1 COM NO. 2 RADAR

DME NO. 2 (opt)

DME NO. 1

ATC NO. 1 ATC NO. 2 GPS NO. 2

(opt)GPS NO. 1

PILOT AUDIO CABIN AUDIO COPILOT AUDIO

SELCAL (opt)

PILOT AUDIO CONT

COPILOT AUDIO CONT

HF ANT(opt)

CCP DIALER (opt) HF COM (opt)TEL (opt) FSU

COM NO. 3 (opt)

FSU FAN

CMU (opt)XM WX (opt)

7-57February, 2012

Systems Description Model B200GT/B200CGTSection 7

Left Fold Under


No. 2 Dual-Fed Bus

No. 3 Dual-Fed Bus

No. 4 Dual-Fed Bus

Avionics Bus No. 1

(L Gen Bus)

Avionics Bus No. 2

(R Gen Bus)

Avionics Bus No. 3

(L Gen Bus)




Fed Bus)


(RW)

ESIS ESIS BUS CONT

ESIS HDG SNSR

ESIS DISPESIS PNL

Electrical GEN CONTROL, L

GEN CONTROL, R

R CB PANEL NO. 3 (FP) (2)

R CB PANEL NO. 4 (FP) (2)

BATTERY RELAY

BUS FEEDERS NO. 1 (2 ea.)



(FP)


(FP)Engines

(including Engine

Instruments and

Propellers)

CHIP DTR, L CHIP DTR, R START CONTROL, L

(FP)

START CONTROL, R

(FP)

ENG FIRE EXTINGUISH

ER, L (opt)DCU SEC, NO.

1 (FP)DCU SEC, NO.

2 (FP)EDC NO. 1 (FP) EDC NO. 2 (FP) ENG FIRE

EXTINGUISHER, R (opt)

ENG INST POWER, L

ENG INST POWER, R

DCU NO. 1 (FP) DCU NO. 2 (FP)

FIRE DETR IGNITER POWER, L (FP)

IGNITER POWER, R (FP)

MN ENG ANTI ICE, L

MN ENG ANTI ICE, R

OIL PRESS, L (FP)

OIL PRESS, R (FP)

OIL PRESS WARN, L

OIL PRESS WARN, R

PROP SYNC AUTOFEATHER PROP GOV (FP)

STBY ENG ANTI ICE, L

STBY ENG ANTI ICE, R

TORQUE, L (FP)

TORQUE, R (FP)

Environmental BLEED AIR CONTROL, L

BLEED AIR CONTROL, R

OXYGEN CONTROL

TEMP CONTROL

PRESS CONTROL

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No. 2 Dual-Fed Bus

No. 3 Dual-Fed Bus

No. 4 Dual-Fed Bus

Avionics Bus No. 1

(L Gen Bus)

Avionics Bus No. 2

(R Gen Bus)

Avionics Bus No. 3

(L Gen Bus)




Fed Bus)


(RW)

Flight PITCH TRIM RUDDER BOOST

FLAP MOTOR (FP)

OUTSIDE AIR TEMP

FLAP CONTROL (incl. Flap Indicator)

(FP)FGC SERVO

NO. 1FGC SERVO

NO. 2Furnishings CIGAR

LIGHTERMASTER POWER

Fuel AUX TRANSFER, L (FP) (Incl. Aux Fuel Qty Warn)

AUX TRANSFER, R (FP) (Incl. Aux Fuel Qty Warn)

FIREWALL SHUTOFF VALVE, L

FIREWALL VALVE, L (FP)

FIREWALL VALVE, R (FP)

FIREWALL SHUTOFF VALVE, R

PRESS WARNING, L

(FP)

PRESS WARNING, R

(FP)QTY IND, L (FP) QTY IND, R

(FP)CROSSFEED

(FP)STANDBY

PUMP, L (FP)STANDBY

PUMP, R (FP)Landing Gear LANDING

GEAR RELAYLights AVIONICS ANN INSTR.

INDIRECTFGP CDU LIGHTS BEACON (SW) RECOG. (SW) ENTRY

LIGHTNO SMK FSB &

CABINREADING MFD & RTU LANDING, L

(SW)LANDING, R

(SW)PILOT FLT SIDE PNL

COPLT FLT INSTR

PILOT INST CONT (2)

COPILOT INST CONT

STROBE (SW) NAV (SW)

ICE (SW) PILOT PFD & DCP

COPILOT PFD & DCP

TAIL FLOOD (SW)

SUB PNL, OVHD &

CONSOLE

PED CONT (2) TAXI (SW)

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Systems Description Model B200GT/B200CGTSection 7

Left Fold Under

(1) If either AHC circuit breaker opens, power will be lost to the respective AHC after approximately 10 minutes.(2) There are some cases where loss of one circuit breaker can affect the power delivered to other circuit breakers. The following table shows these effects.

In addition to the equipment shown in the above table, the following equipment is powered directly from the Left or Right Generator Buses through individual circuit protection devices.These devices (breakers or current limiters) are located under the floor boards and are not accessible to the pilot.


No. 2 Dual-Fed Bus

No. 3 Dual-Fed Bus

No. 4 Dual-Fed Bus

Avionics Bus No. 1

(L Gen Bus)

Avionics Bus No. 2

(R Gen Bus)

Avionics Bus No. 3

(L Gen Bus)




Fed Bus)


(RW)

Warnings ANN POWER ANN INDBLEED AIR WARN, L

BLEED AIR WARN, R

LANDING GEAR WARN

LANDING GEAR IND

STALL WARN

Weather FUEL VENT, L FUEL VENT, R PROP DEICE, L (manual) (FP)

PROP DEICE, R (manual) (FP)

PITOT, L (SW) PITOT, R (SW)

SURF DEICE WSHLD WIPER PROP DEICE CONTROL, (FP)

PROP AUTO (SW)

STALL WARN (SW)

BRAKE DEICE (opt)

Tripped Circuit Breaker EffectDC CONV NO. 1 Power will be lost to the Pilot’s PFD and the AHC SEC NO. 1 Circuit Breakers.DC CONV NO. 2 Power will be lost to the MFD and AHC SEC NO. 2 Circuit Breakers.PILOT INST CONT Power will be lost to the FGP, MFD & RTU, & PILOT PFD & DCP Circuit Breakers.PED CONT Power will be lost to the CDU LIGHTS Circuit Breaker.R CB PANEL NO. 3 Power will be lost to the PILOT INST CONT, FGC SERVO NO. 1, DBU, LEFT IAPS, and PILOT PFD

HEATER Circuit Breakers.R CB PANEL NO. 4 Power will be lost to the PED CONT, COPILOT INST CONT, FGC SERVO NO. 2, EGPWS, RIGHT IAPS,

and MFD HEATER Circuit Breakers.

Miscellaneous Items on the Left Generator Bus Miscellaneous Items on the Right Generator Bus Miscellaneous Items on the Isolation BusBlower, Cabin Fwd Air Conditioner Clutch Electric HeatCondenser Blower Blower, Cabin AftPilot Windshield Heat Blower, CockpitRadiant Heat, (Cargo Door Only) Copilot Windshield Heat

DC Test JackInverter for Cabin OutletsLanding Gear MotorRefreshment BarToilet

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EXTERNAL POWER For ground operation, an external power socket, located under the right wing out-board of the nacelle, is provided for connecting an auxiliary power unit. A relay inthe external power circuit will close only if the external source polarity is correct.The BATTERY MASTER SWITCH must be on before the external power relay willclose and allow external power to power the airplane electrical system. The bat-tery will also tend to absorb voltage transients when operating avionic equipmentand during engine starts. Otherwise, the transients might damage the many solidstate components in the airplane.

When an external power source is used, the Ground Power Unit (GPU) shall becapable of producing 1000 amperes for 5 seconds, 500 amperes for two minutesand 300 amperes continuously. A caution annunciator on the caution/advisory an-nunciator panel, EXT PWR, is provided to alert the operator when an external DCpower plug is connected to the airplane.

LIGHTING SYSTEMS COCKPIT An overhead light control panel, easily accessible to both pilot and copilot, incor-porates a functional arrangement of all lighting systems in the cockpit. Each lightgroup has its own rheostat switch placarded BRT - OFF. The MASTER PANELLIGHTS - ON - OFF switch controls the pilot instrument panel lights, copilot instru-ment panel lights, overhead panel & overhead instrument lights, pedestal lights,sidepanel lights, and subpanel lights. The instrument indirect lights in theglareshield and overhead map lights are individually controlled by separate rheo-stat switches.

CABIN A three-position switch on the copilot’s subpanel, placarded CABIN LIGHTS -START/BRIGHT - DIM - OFF, controls the cold cathode cabin lights. The switchto the right of the interior light switch activates the cabin NO SMOKING/FASTENseatbelt signs and accompanying chimes. This three-position switch is placardedCABIN LIGHTS - NO SMOKE & FSB - FSB - OFF.

The baggage-area light is controlled by a two-position switch just inside the airstairdoor aft of the door frame and is connected to the hot battery bus.

A threshold light is located forward of the airstair door at floor level, and an aislelight is located at floor level aft of the spar cover. A switch adjacent to the thresholdlight turns both these lights on and off. The switch also turns the exterior entry lighton and off. When the airstair door is closed, all the lights controlled by the thresh-old light switch will extinguish.

When the master switch is on, the individual reading lights along the top of thecabin may be turned on or off by the passengers with a push-button switch adja-cent to each light.

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EMERGENCY EXIT LIGHTING SYSTEM

B200GTInstalled by Kit P/N 101-5301-1:

The emergency exit lighting system is self-activating and designed to provide illu-mination of the cabin during an emergency evacuation. The system may also beused to provide courtesy lighting during normal operations. Illumination is provid-ed by three lights: one in the emergency exit, one opposite the airstair door andone in the aft baggage compartment. Power for the system is provided by the air-plane battery.

During an emergency, the lights will illuminate during a loss of power from the No.1 dual-fed bus or when the airstair door is unlocked, whichever occurs first. Onceactivated, the lights will remain illuminated for approximately 15 minutes.

During normal operations, the system will operate as follows. Upon entering theairplane the lights will illuminate when the airstair door is unlocked. If the airstairdoor was previously left open for longer than 15 minutes, the lights may be re-ac-tivated by either of two methods:

1. Press and release the door lock switch, which is contained in the lower left door-locking-pin receptacle on the fuselage.

2. Turn the battery switch on, then turn the battery switch off.

If the lights are activated by unlocking the airstair door (or pressing the door lockswitch), they will extinguish when the door is closed and re-locked. If the lights areactivated by cycling the airplane power, the lights will extinguish when the batteryswitch is turned on, such as before an engine start. At the end of the flight, thelights will again illuminate once the battery switch is turned off. If the airstair dooris closed after exiting the airplane, the lights will extinguish approximately 15 min-utes after battery switch was turned off. If the airstair door is left open, the lightswill extinguish approximately 15 minutes after the airstair door was unlocked.

B200CGTInstalled by Kit P/N 101-5301-3:

The emergency exit lighting system is self-activating and designed to provide illu-mination of the cabin during an emergency evacuation. The system may also beused to provide courtesy lighting during normal operations. Illumination is pro-vided by three lights: one in the emergency exit, one above the airstair door andone in the aft baggage compartment. Power for the system is provided by the air-plane battery.

During an emergency, the lights will illuminate during a loss of power from the No.1 dual-fed bus or when the airstair door is opened, whichever occurs first. Onceactivated, the lights will remain illuminated for approximately 15 minutes.

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7-63

During normal operations, the system will operate as follows. Upon entering theairplane, the lights will illuminate when the airstair door is opened. If the airstairdoor was previously left open for longer than 15 minutes, the lights may be re-ac-tivated by either of two methods:

1. Press and release the door lock switch, which is contained in the lower left door-locking-pin receptacle on the fuselage.

2. Turn the battery switch on, then turn the battery switch off.

If the lights are activated by opening the airstair door (or pressing the airstair doorsill switch), they will extinguish when the door is closed. If the lights are activatedby cycling the airplane power, the lights will extinguish when the battery switch isturned on, such as before an engine start. At the end of the flight, the lights willagain illuminate once the battery switch is turned off. If the airstair door is closedafter exiting the airplane, the lights will extinguish approximately 15 minutes afterbattery switch was turned off. If the airstair door is left open, the lights will extin-guish approximately 15 minutes after the airstair door was opened.

EXTERIOR Switches for the landing lights, taxi lights, wing ice lights, navigation lights, recog-nition lights, flashing beacons, and wing-tip and tail strobe lights are located on thepilot’s subpanel. They are appropriately placarded as to their function.

Tail floodlights are incorporated into the horizontal stabilizers and are designed toilluminate both sides of the vertical stabilizer. A switch for these lights, placardedLIGHTS - TAIL FLOOD - OFF, is located on the pilot’s subpanel.

A flush-mounted floodlight forward of the flaps in the bottom of the left wing is in-stalled. This entry light provides illumination of the area around the airstair door,to provide passenger convenience at night. It is controlled by the threshold lightswitch just inside the door on the forward door frame, and will extinguish automat-ically whenever the cabin door is closed.

ENVIRONMENTAL SYSTEM The environmental system consists of the bleed air pressurization, heating andcooling systems, and their associated controls.

PRESSURIZATION SYSTEM The pressurization system is designed to provide a normal working pressure dif-ferential of 6.5 ± .1 psi, which will provide cabin pressure altitudes of approximate-ly; 2800 feet at an airplane altitude of 20,000 feet, 8600 feet at 31,000 feet; and10,400 feet at 35,000 feet.

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7-64

IN-FLOW SYSTEM

The pressurization air in-flow system consists of a bleed air flow control valve foreach engine that is controlled by 2 three-position switches placarded BLEED AIRVALVES - LEFT - RIGHT in the ENVIRONMENTAL controls group on the copilot’sleft subpanel. The three switch positions are placarded OPEN - ENVIR OFF -PNEU & ENVIR OFF. When a switch is in either the ENVIR OFF or the PNEU &ENVIR OFF position, the respective right or left environmental air valve is closed.When a switch is in the OPEN position, the air mixture will flow through the valvetoward the cabin.

Environmental bleed air flow volume is controlled by the switch placarded ENVIRBLEED AIR - NORMAL - AUTO - LOW in the ENVIRONMENTAL controls groupon the copilot’s left subpanel. The LOW position reduces the bleed air extractedfrom the engines for environmental purposes to approximately half the normalamount. This position may be used during operations in ambient temperaturesabove 10°C for improved cooling. The NORMAL position may be used for in-creased heating or if increased pressurization airflow is required. This positionshould be selected during the climb phase of flight to ensure optimum perfor-mance of the pressurization system at higher altitudes. The AUTO position is therecommended setting, and it allows the environmental system controller to auto-matically select the flow setting based upon the heat demanded to maintain cabin/cockpit temperature or cabin pressure requirements. The default flow setting willbe LOW in all environmental cases except when the system is demanding addi-tional heat. In order for the Bleed Air AUTO position to function properly in re-sponse to the heating/cooling requirements as commanded by the controller, theEnvironmental Mode control must be selected to the AUTO position. If the red ALTWARN annunciator is detected as illuminated by the controller and after a five sec-ond time delay, the flow setting will switch to NORMAL until the annunciator is ex-tinguished and a pre-set time-delay has expired. When the controller switchesfrom LOW flow to NORMAL, the controller will monitor bleed duct temperature andmake corrections to the heat exchanger bypass valves to prevent any duct overtemperatures. If the environmental control knob is selected to MAN HEAT, thebleed flow defaults to NORMAL. If the flow is selected to MAN COOL, the bleedflow defaults to LOW.

Always monitor cabin pressurization requirements if in MAN COOL.Manual adjustments to the ENVIR BLEED AIR flow setting may berequired.

For maximum engine performance and/or high altitude take-off requirements, theENVIR BLEED AIR switch should be manually selected to LOW.

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7-65

CABIN PRESSURE CONTROL SYSTEM

The cabin pressure control system consists of an outflow valve and a safety valve,which are vented overboard to preclude moisture build-up in the aft fuselage, anda pressurization controller.

The pressurization controller, mounted in the pedestal, controls modulation of theoutflow valve. The outer scale (CABIN ALT) of the dual-scale indicator dial indi-cates the cabin pressure altitude which the controller is set to maintain. The innerscale (ACFT ALT) indicates the maximum ambient pressure altitude at which theairplane can fly without causing the cabin pressure altitude to exceed the valueselected on the outer scale (CABIN ALT) of the dial.

The rate control selector knob is placarded RATE - MIN - MAX. The rate at whichthe cabin pressure altitude changes is controlled by rotating the rate control selec-tor knob. The rate of change selected may be from a minimum of approximately200 to a maximum of approximately 2000 feet per minute.

The actual cabin pressure altitude is continuously indicated by the cabin altimeter,which is mounted in the right side of the panel. Immediately to the left of the cabinaltimeter is the cabin vertical speed (CABIN CLIMB) indicator, which continuouslyindicates the rate at which the cabin pressure altitude is changing.

The cabin pressure switch, located forward of the pressurization controller on thepedestal, is placarded CABIN PRESS - DUMP - PRESS - TEST. When this switchis in the DUMP (forward) position, the safety valve is held open so that the cabinwill depressurize and/or remain unpressurized. When it is in the PRESS (center)position, the safety valve is normally closed in flight, and the outflow valve is con-trolled by the pressurization controller so that the cabin will pressurize. When theswitch is held in the spring-loaded TEST (aft) position, the safety valve is heldclosed, bypassing the landing-gear safety switch, to facilitate testing of the pres-surization system on the ground.

Prior to takeoff, the cabin altitude selector knob should be adjusted so that theACFT ALT scale on the indicator dial indicates an altitude approximately 1000 feetabove the planned cruise pressure altitude, or the CABIN ALT scale indicates analtitude at least 500 feet above the take-off field pressure altitude. The rate controlselector knob should be adjusted as desired; setting the index mark at the 12-o’clock position will provide the most comfortable cabin rate of climb. The cabinpressure switch should be checked to ensure that it is in the PRESS position. Asthe airplane climbs, the cabin pressure altitude climbs at the selected rate ofchange until the cabin reaches the selected pressure altitude. The system thenmaintains cabin pressure altitude at the selected value. If the airplane climbs to analtitude higher than the value indexed on the ACFT ALT scale of the dial on theface of the controller, the cabin-to-ambient pressure differential will reach thepressure relief setting of the outflow valve and safety valve (6.5 psi cabin-to-am-bient differential). Refer to the High Differential Pressure (Cabin Differential Pres-sure Exceeds 6.6 psi) procedure in Section 3, EMERGENCY PROCEDURES, forcorrective action.

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7-66

At 12,500 feet, the oxygen masks will drop out. During cruise operation, if the flightplan calls for an altitude change of 1000 feet or more, reselect the new altitudeplus 1000 feet on the CABIN ALT dial. During descent and in preparation for land-ing, the cabin altitude selector should be set to indicate a cabin altitude of approx-imately 500 feet above the landing field pressure altitude.

UNPRESSURIZED VENTILATION

Fresh-air ventilation is provided from both engines, which is available during boththe pressurized and the unpressurized mode, is the bleed air heating system. Thisair mixes with recirculated cabin air and enters the cabin through the floor regis-ters.

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7-67

THIS PAGE INTENTIONALLY LEFT BLANK

February, 2012



B200GT ENVIRONMENTAL SCHEMATIC

RECEIVER DRYER

CONDENSER BLOWER

OUTLET AIR

CONDENSER

INLET AIR

REFRIGERANT SERVICE VALVES

COCKPIT HEAT & DEFROST AIR

PRESSURIZATION CONTROLLER

ENGINE BLEED AIR

TEMP BULB & EXPANSION VALVE

FWD EVAPORATOR AND BLOWER

VENTURIR.H. SUBPANEL

DUCT OVERTEMP SENSOR SWITCH

R.H. C.B. PANEL

ELEC GROUND HEAT OUTLET

REFRIGERANT COMPRESSOR

ENGINE BLEED AIR

AMBIENT TEMP SENSOR

AMBIENT AIR INLETFIREWALL SHUTOFF VALVE

BLEED AIR SHUTOFF VALVE

FLOW CONTROL VALVE

FIREWALL

FIREWALL

HEAT EXCHANGER

BLEED AIR BYPASS VALVE

AFT MIXING BOX

FWD MIXING BOX

HIGH / LOW PRESSURE SWITCH

AIR INLET SCOOP

CABIN COOL AIR OUTLETS (6 PLACES)

TEMP CONTROLLER

CHECK VALVES

OVERBOARD DUCTS

OUTFLOW VALVE

SAFETY VALVE

AFT COMPARTMENT COOL AIR OUTLET

AFT COMPARTMENT HEAT OUTLETS

AFT EVAPORATORS AND BLOWERS

CABIN HEAT OUTLETS 6 PLACES

DC POWER DISTRIBUTION PANEL

HEAT EXCHANGER

FLOW CONTROL VALVE


FIREWALL SHUTOFF VALVE

AMBIENT AIR INLET

COCKPIT COOL AIR OUTLETS

CIRCUIT CARD BOX

AIR INLET SCOOP

ENGINE BLEED AIR DUCT

VALVE ASSY AND BLEED AIR BYPASS

LINESREFRIGERANT

ELEC GROUND HEAT ELEMENTS

MUFFLER/DISTRIBUTION BOX

AMBIENT AIR

COOLED AIR

HEATED AIR

BB07C 062003AA.AI

7-68 February, 2012

Section 7R

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B200CGT ENVIRONMENTAL SCHEMATIC

RECEIVER DRYER

CONDENSER BLOWER

OUTLET AIR

CONDENSER

INLET AIR

REFRIGERANT SERVICE VALVES

COCKPIT HEAT & DEFROST AIR

PRESSURIZATION CONTROLLER

ENGINE BLEED AIR

TEMP BULB & EXPANSION VALVE

FWD EVAPORATOR AND BLOWER

VENTURIR.H. SUBPANEL

DUCT OVERTEMP SENSOR SWITCH

R.H. C.B. PANEL

ELEC GROUND HEAT OUTLET

REFRIGERANT COMPRESSOR

ENGINE BLEED AIR

AMBIENT TEMP SENSOR

AMBIENT AIR INLETFIREWALL SHUTOFF VALVE


FLOW CONTROL VALVE

FIREWALL

FIREWALL

HEAT EXCHANGER

BLEED AIR BYPASS VALVE

AFT MIXING BOX

FWD MIXING BOX

HIGH / LOW PRESSURE SWITCH

AIR INLET SCOOP

CABIN COOL AIR OUTLETS (6 PLACES)

TEMP CONTROLLER

CHECK VALVES

OVERBOARD DUCTS

OUTFLOW VALVE

SAFETY VALVE

AFT COMPARTMENT COOL AIR OUTLET

AFT COMPARTMENT HEAT OUTLETS

AFT EVAPORATORS AND BLOWERS

CABIN HEAT OUTLETS 6 PLACES

DC POWER DISTRIBUTION PANEL

HEAT EXCHANGER

FLOW CONTROL VALVE


FIREWALL SHUTOFF VALVE

AMBIENT AIR INLET

COCKPIT COOL AIR OUTLETS

CIRCUIT CARD BOX

AIR INLET SCOOP

ENGINE BLEED AIR DUCT

VALVE ASSY AND BLEED AIR BYPASS

LINESREFRIGERANT

ELEC GROUND HEAT ELEMENTS

MUFFLER/DISTRIBUTION BOX

AMBIENT AIR

COOLED AIR

HEATED AIR

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BLEED AIR HEATING Engine bleed air, through the Bleed Air Valves, is utilized to warm the cockpit andcabin.

Heating air outlets are provided for each pilot under the instrument panel, and inthe floor, outboard of the pilot’s seats. The COCKPIT TEMP control knob, locatedin the ENVIRONMENTAL group on the copilot’s left subpanel, regulates the tem-perature of the air supplied to these outlets while in the AUTO mode. A constantflow of conditioned air is supplied to the glareshield outlets and the windshield de-frost outlets. In AUTO mode, this temperature is regulated to a maximum defaulttemperature of approximately 70°F. In colder applications when more heat is ini-tially demanded, this outlet duct temperature is allowed to reach approximately105°F. In MAN HEAT, the overhead and glareshield duct temperatures are fullycontrollable by the COCKPIT/CABIN TEMP knobs, which regulate the amount ofheat added to the recirculated air exiting these outlets.

If the bleed air temperature in the ducts supplying the floor outlets becomes ex-cessive, the amber DUCT OVERTEMP annunciator will illuminate. Refer to theDUCT OVERTEMPERATURE procedure in Section 3A, ABNORMAL PROCE-DURES, for corrective action.

ELECTRIC HEATING

A supplemental electric heating system is available for cabin conditioning. It is op-erated by rotating the MODE knob in the ENVIRONMENTAL group on the copi-lot’s left subpanel, to the ELEC HEAT position. The supplemental electric heatingsystem can only be used on the ground.

Do not operate the electric heat with the pedestal floor outlet blockedor the cockpit door closed.

This system utilizes a heater assembly containing six heating elements located ina duct aft of the forward evaporator. The cockpit blower is used to distribute airthrough the electric heating duct, and will operate automatically when the ELECHEAT mode is selected. Heated air is directed into the cabin through a single floodoutlet located directly aft of the cockpit pedestal. The blower will operate at maxi-mum speed regardless of the indication of the COCKPIT BLOWER knob. An am-ber ELEC HEAT ON caution annunciator is provided to indicate that the electricheat power relay is closed and applying power to the electric heat power controlrelay for the heating elements. When the ELEC HEAT mode is deselected theELEC HEAT ON annunciator must extinguish to verify that power is removed fromthe heating elements. If it remains illuminated, the system is not operating properlyand maintenance is required prior to flight. In order to maintain airflow across heat-ing elements, ELEC HEAT should be re-selected until airplane is shutdown. Safety devices built into the heater assembly may continue to temporarily powerthe blower at a low speed to cool the heater elements and avoid overheating theduct. In the event that residual heat in the elements causes the duct temperature

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to rise after the blower has initially shutdown, the blower will automatically cycleto cool the elements regardless of Battery switch position. When the airplane isnot powered, this blower is powered through circuit breaker CB14 located in thehot battery bus box, which is just outboard of the battery.

NOTEThe electric heat system will draw approximately 160 amps.

RADIANT HEATING (B200CGT)

On the B200CGT, a radiant heater element is installed in the cargo door. It is con-trolled by the Mode switch and operates in all heating modes. This unit providessupplemental heat to the cabin for additional passenger comfort.

AIR CONDITIONING SYSTEM Cabin/Cockpit air conditioning is provided by a vapor-cycle refrigeration system.The compressor, driven by the right engine, will operate as required in the AUTOor MAN COOL control modes, provided operation is not prohibited by the systemprotection controls. System protection controls will prevent compressor operationif refrigerant pressure is too high or too low, if the bleed air bypass valve hasreached a limit switch (indicating significant heat is being introduced), or if right en-gine is below 62% N1. If operation is prevented due to low N1 speed, the greenAIR COND N1 LOW annunciator will illuminate.

The cockpit blower assembly recirculates cockpit air through the forward evapo-rator and into the 4 cockpit distribution ducts which supply glareshield outlets andwindshield defrost. Two cabin blower assemblies provide cooling to the main cab-in by recirculating cabin air through two evaporators and into ducting which sup-plies the 7 eyeball air outlets in the cabin headliner, and the 2 eyeball air outletsin the cockpit headliner.

ENVIRONMENTAL CONTROLS The ENVIRONMENTAL control section on the copilot’s left subpanel provides forautomatic or manual control of the system. The system is a dual zone design, al-lowing for independent control to the temperature in the cabin and cockpit.

AUTOMATIC MODE CONTROL

When the MODE selector switch is set to the AUTO position, the heating and airconditioning systems operate automatically. The system will automatically adjustblowers speed, bleed air temperature and compressor clutch on/off state to main-tain the temperature setpoints selected via the TEMP knobs. The recommendedtemperature setting is straight up at the 12 o’clock position which equates to ap-proximately 75°F. In addition, the controller will modulate 5 servo-operated airflowvalves in the bleed air heat ducting that direct bleed air into the various sectionsof the airplane on an “on demand” basis in order to help maintain the desired tem-perature setpoints. When there is little or no demand for bleed air heat in the cabinor cockpit, the majority of the conditioned bleed air is directed aft to the baggage

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compartment. Temperature sensing devices in the cockpit and cabin, in conjunc-tion with the TEMP settings, signal the controller to make the necessary adjust-ments to maintain the setpoint temperature. If at any time the operator wishes tomanually select a different blower speed, the BLOWER control can be rotated outof the AUTO detent to the desired speed. User patience should be practiced withthe temperature control setpoint. The controller will maintain setpoint over time.This slow transition is purposefully built-in to eliminate unwanted temperature vari-ations.

MANUAL MODE CONTROL

When the MODE selector is set to the MAN HEAT position, the Cockpit and Cabinfloor heat servos are fully opened and regulation of the cabin and cockpit temper-atures is accomplished by actuating the MAN TEMP switch to either the INCR orDECR position desired. When released, this switch will return to the center (nochange) position. This regulates the temperature of the bleed air entering the air-plane, while the flow rate remains unchanged. Bleed air temperature response isproportional to the length of time the MAN TEMP switch is actuated with approxi-mately 30 seconds required to go from full increase to full decrease or vice versa,Actuations should be 2-3 seconds in duration with approximately 60 seconds inbetween to avoid temperature over/undershoots.

Longer than 2-3 second switch actuations and shorter than 60 sec-onds in between may result in a duct overheat situation. Refer to theDUCT OVERTEMPERATURE procedure in Section 3A, ABNOR-MAL PROCEDURES if this occurs.

The COCKPIT TEMP control knob can then be used to manually select the tem-perature of the glareshield and windshield defrost outlets. The CABIN TEMP con-trol works similarly: When the TEMP control rotated to full counter clockwise(CCW), the air out of the cabin and cockpit overhead outlets is the coolest (recir-culated air without added bleed air heat). The CABIN and COCKPIT BLOWERcontrols can be used in this mode to control the amount of recirculated air exitingfrom the appropriate outlets for air recirculation.

When the MODE selector switch is set to the MAN COOL position, the air condi-tioner system will operate, provided the speed of the right engine is above 62%N1, and the system pressures are within range. To prevent the evaporator coilsfrom freezing, the blowers will default to a preset minimum speed. In this mode,the TEMP setting knobs operate the same as in MAN HEAT and blower speed canbe changed by varying the CABIN and COCKPIT BLOWER speeds.

In either manual modes, it is the responsibility of the pilot to activelymonitor the temperature and flow of the bleed air entering the cabin.

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ELECTRIC HEAT MODE CONTROL

When the MODE selector switch is set to the ELEC HEAT position, air is directedover resistive heater elements and into the cabin as described above under theELECTRIC HEATING section. This mode is only operative on the ground.

USING BLOWERS FOR AIR RECIRCULATION

COCKPIT and CABIN BLOWER speed may be manually adjusted to obtain thedesired amount of flow exiting the outlets for air recirculation in any mode exceptELEC HEAT, where the cockpit blower defaults to a high setting.

When the blower switches are in the AUTO detent (full CCW), and the Environ-mental Mode control is in AUTO, blower speed will be adjusted automatically bythe controller. When predominantly cooling the cockpit/cabin, the further the actu-al cockpit/cabin temperature is away from the selected setpoint, the higher theblower speed will be. As the actual temperature approaches the setpoint temper-ature, the controller will begin to reduce the blower speed until the setpoint tem-perature is achieved. At this point the blower speed will be maintained at a defaultminimum speed. When predominantly heating the cockpit/cabin, the blower willdefault to a minimum speed.

FAULT DIAGNOSIS

If the environmental controller detects an open or erratic reading on one or moreof the temperatures sensors throughout the airplane and the Environmental Con-trol knob and COCKPIT BLOWER are selected to AUTO, the cockpit blower willpulse between high and low setting within a time period of approximately 5 sec-onds. If this occurs, it is recommended not to operate the environmental controlsystem in AUTO until the problem is rectified. The blowers can be operated out ofLOW speed as desired/required if this fault is detected.

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OXYGEN SYSTEM

OXYGEN SYSTEM SCHEMATIC

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The Model B200GT/B200CGT oxygen system is based on an adequate oxygenflow for a pressure altitude of 35,000 feet. The masks and Oxygen Duration Chart(Section 4, NORMAL PROCEDURES) are based on a flow rate of 3.9 liters perminute - normal temperature pressure dry (LPM-NTPD). The only exception is thediluter-demand crew mask. For oxygen duration computation, each diluter-de-mand mask being used is counted as two masks at 3.9 LPM-NTPD. At cabin alti-tudes above 20,000 feet, SELECT 100% MODE.

A push/pull handle (PULL ON - SYStem READY), located on the left side of thepedestal, is used in conjunction with the automatically deployed passenger oxy-gen system. This handle operates a cable which opens and closes the shutoffvalve located at the oxygen supply bottle in the aft, unpressurized area of the fu-selage. When this handle is pushed in, no oxygen supply is available anywhere inthe airplane. It should be pulled out prior to engine starting to ensure that oxygenwill be immediately available anytime it is needed. When this handle is pulled out,the primary oxygen supply line is charged with oxygen, provided the oxygen sup-ply bottle is not empty (check the oxygen supply pressure gage on the right sub-panel and verify that sufficient oxygen is available for the flight). The primaryoxygen supply line delivers oxygen to the two crew oxygen outlets in the cockpit,to the first aid oxygen outlet in the toilet area, and to the passenger oxygen systemshutoff valve.

The crew is provided with diluter-demand, quick-donning oxygen masks which arelocated in the overhead. The crew masks should always be plugged in and stowedso that oxygen will be immediately available when required. This will not cause aloss of oxygen since the diluter demand masks deliver oxygen only upon inhala-tion. To don the mask, grasp the red levers protruding from the stowage compart-ment and pull the mask down. Inflate the mask harness by depressing the redlever on the left side of the regulator and then don the mask and release the lever.Three modes of operation are available to the crew: NORMAL (diluted oxygen),100% and EMERG. NORMAL or 100% can be used at any altitude at the user’sdiscretion and is controlled by a selector lever on the bottom right side of the reg-ulator. The EMERG mode supplies a positive pressure to the face piece andshould be used if smoke or fumes are present in the cabin. To use the mask in theEMERG mode, turn the emergency knob located on the bottom of the regulator.After donning the mask, check the flow indicator in the oxygen supply hose to en-sure oxygen is being supplied to the regulator (RED - no flow; WHITE - flow). Thecrew masks also contain integral microphones.

Anytime the primary oxygen supply line is charged, oxygen can be obtained fromthe first aid oxygen mask located in the toilet area, by manually opening the over-head access door (placarded FIRST AID OXYGEN - PULL) and opening the ON-OFF valve inside the box. A placard (NOTE: CREW SYStem MUST BE ON) re-minds the user that the PULL ON - SYStem READY handle in the cockpit must bepulled out before oxygen will flow from the first aid oxygen mask.

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The passenger oxygen system is of the constant flow type. Anytime the cabinpressure altitude exceeds approximately 12,500 feet, a barometric-pressureswitch automatically energizes a solenoid which opens the passenger oxygensystem shutoff valve. The pilot can open the valve manually anytime by pulling outthe PASSENGER MANUAL OverRIDE handle, located on the right side of thepedestal. Once the passenger oxygen system shutoff valve has been opened (ei-ther automatically or manually), oxygen will flow into the passenger oxygen supplyline, if the primary oxygen system line has been charged (i.e., if the oxygen supplybottle contains oxygen and the PULL ON - SYStem READY handle in the cockpitis pulled out). When oxygen flows into the passenger oxygen system supply line,a pressure-sensitive switch in the line closes a circuit to illuminate the green PASSOXYGEN ON annunciator on the caution/advisory annunciator panel. This switchwill also cause the cabin lights, the foyer light and the center baggage compart-ment light to illuminate in the full bright mode, regardless of the position of the in-terior lights switch placarded CABIN LIGHTS - START/BRIGHT - DIM - OFFlocated on the copilot’s left subpanel.

The pressure of the oxygen in the passenger oxygen system supply line then au-tomatically extends a plunger against each of the passenger oxygen mask dis-penser doors, forcing the doors open. The oxygen masks then drop down about9 inches below the dispensers. The lanyard valve pin at the top of the oxygenmask hose must be pulled out in order for oxygen to flow from the mask. The pinis connected to the oxygen mask via a flexible cord; when the oxygen mask ispulled down for use, the cord pulls the pin out of the lanyard valve. The lanyardvalve pin must be manually reinserted into the valve in order to stop the flow ofoxygen when the mask is no longer needed. The passenger oxygen can be shutoff and the remaining oxygen isolated to the crew and first aid outlets by pullingthe OXYGEN CONTROL circuit breaker in the ENVIRONMENTAL group on theright side panel, providing the PASSENGER MANUAL O’RIDE handle is pushedin to the OFF position.

PITOT AND STATIC SYSTEM Independent pitot and static systems are provided for the pilot’s flight indicatorsand the copilot’s flight indicators.

The pilot’s and copilot’s pitot masts are located on the nose section of the airplane.Each mast provides total air pressure to its respective air data computer (ADC).The copilot’s mast also provides total air pressure to the Electronic Standby Instru-ment System (ESIS).

Dual static air ports are located on each side of the aft fuselage in a vertical ar-rangement. The top port on the left side is connected to the bottom port on theright side and the resulting average pressure is supplied to the pilot’s static airsource valve. In addition, an alternate static air source is provided to the pilot’sstatic air source valve from the aft side of the rear pressure bulkhead. The outputfrom the pilot’s static air source valve is manually selected by the crew and pro-vides either normal static air pressure or alternate static air pressure to the pilot’s

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ADC. The bottom static port on the left side of the fuselage is connected to the topstatic port on the right side of the fuselage and the resulting average pressure issupplied to the copilot’s ADC and the ESIS.

In addition to pitot and static air pressure inputs, the pilot’s and copilot’s ADCs alsoreceive separate inputs of air temperature. With these inputs, each ADC is capa-ble of supplying digital air data to the pilot’s and copilot’s PFDs and the MFD viadata buses for the display of the following information. Uncorrected Pressure AltitudeBaro-Corrected AltitudeVertical SpeedAirspeed (KIAS & KCAS)Indicated Airspeed Trend VectorMach NumberMaximum Airspeed (VMO/MMO)True AirspeedRam Air Temperature (RAT)Static Air Temperature (SAT)ISA Delta Temperature

The pilot’s static air source valve is located on the right side panel in the cockpit.The alternate static air source is selected by lifting a spring-loaded retaining clipand moving the valve handle aft from the NORMAL position to the ALTERNATEposition.

The pilot’s airspeed, altitude and vertical speed indications willchange when the alternate static air source is activated. Refer to theAirspeed Calibration - Alternate System, and the Altimeter Correc-tion - Alternate System graphs in Section 5, PERFORMANCE, foroperations when the alternate static air source is selected.

When the alternate static air source is no longer needed, ensure that the static airsource valve is returned to the NORMAL position.

Three petcocks are provided to drain moisture from the static air lines. They arelocated behind an access cover below the circuit breakers on the right side panel.The drain valves should be opened to release any trapped moisture during eachphase inspection and during any other ground inspection when contamination ofthe static lines is suspected.

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PITOT AND STATIC SYSTEM SCHEMATIC

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ENGINE BLEED AIR PNEUMATIC SYSTEM High-pressure bleed air from each engine compressor, routed through the firewallshutoff valves and regulated at 18 psi, supplies pressure for surface deice systemand vacuum source. Vacuum is derived from a bleed air ejector. One engine cansupply sufficient bleed air for all systems.

During single-engine operation, a check valve in the bleed air line from each en-gine prevents flow back through the line on the side of the inoperative engine. Avacuum gage calibrated in inches of mercury, located on the copilot’s subpanel,indicates vacuum. To the right of the vacuum gage is a pneumatic pressure gage,calibrated in pounds per square inch, which indicates air pressure available to thedeice distributor valve.

Refer to the Pneumatic Bleed Air and Surface Deice System Schematic.

BLEED AIR WARNING SYSTEM The bleed air lines from the engines to the cabin are shielded with insulation toprotect other components from heat. Heat is also dissipated in the air-to-air heatexchanger in the center wing section. The bleed air lines are accompanied inclose proximity by plastic tubing from the engines to the cabin. One end of the tub-ing is plugged off; the other end is connected to a bleed air source in the cabin, tosupply the line with pressure. Excessive heat on the plastic tubing caused by aruptured bleed air line will cause the tubing to fail. Upon release of pressure in thetubing, a normally open switch in the line, located under the copilot’s floor in thefuselage, will close, causing a circuit to be completed to the respective BL AIRFAIL annunciator in the warning annunciator panel. When the indication of bleedair line failure becomes evident, the bleed air for that side should be turned off byplacing the respective lever-lock BLEED AIR VALVE switch on the copilot’s sub-panel in the PNEU & ENVIR - OFF position.

STALL WARNING SYSTEM The stall warning system consists of a transducer, a lift computer, a warning horn,and a test switch. Angle of attack is sensed by aerodynamic pressure on the lifttransducer vane located on the left wing leading edge. When a stall is imminent,the output of the transducer activates a stall warning horn.

The system has preflight capability through the use of a switch placarded STALLWARN - TEST - OFF on the copilot’s left subpanel. Holding this switch in the TESTposition activates the warning horn.

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ICE PROTECTION SYSTEMS WINDSHIELD HEAT Windshield heat switches are located on the pilot’s subpanel (inboard) and areplacarded ICE - WSHLD ANTI-ICE - NORMAL - OFF - HI - PILOT - COPILOT.

Two levels of heat are provided. When the switches are in the NORMAL (up) po-sition, heat is supplied to the major portion of the windshields. When they are inthe HI (down) position, a higher level of heat is supplied to a smaller area of thewindshields. Each switch must be lifted over a detent before it can be moved intothe HI position. This lever-lock feature prevents inadvertent selection of the HI po-sition when moving the switches from NORMAL to the OFF (center) position.

Controllers with temperature-sensing units provide for proper heat at the wind-shield surfaces. Five-ampere circuit breakers, located on a panel on the forwardpressure bulkhead, protect the control circuits. The power circuit of each systemis protected by a 50-ampere circuit breaker located in the power distribution panelunder the floor forward of the main spar.

NOTEErratic operation of the magnetic compass may occur while wind-shield heat is being used.

PROPELLER ELECTRIC DEICE SYSTEM The propeller electric deice system includes: electrically heated deice boots, sliprings and brush block assemblies, a timer for automatic operation, an ammeter,three circuit breakers located on the fuel control panel for left and right propellerand control circuit protection, and two switches located on the pilot’s subpanel forautomatic or manual control of the system.

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PROPELLER ELECTRIC DEICE SYSTEM SCHEMATIC

A circuit breaker switch located on the pilot’s subpanel, placarded PROP - AUTO- OFF, is provided to activate the automatic system. Upon placing the switch to theAUTO position, the timer diverts power through the brush block and slip ring to allheating elements on one propeller. Subsequently, the timer then diverts power toall heating elements on the other propeller for the same length of time. This cyclewill continue as long as the switch is in the AUTO position. The system utilizes ametal foil type single heating element energized by DC voltage. The timer switch-es every 90 seconds, resulting in a complete cycle in approximately 3 minutes.

A manual prop deice system is provided as a backup to the automatic system. Acontrol switch located on the left subpanel, placarded PROP - MANUAL - OFF,controls the manual override relay. Upon placing the switch in the MANUAL posi-tion, the automatic timer is overridden and power is then supplied to the heatingelements of both propellers simultaneously. This switch is of the momentary typeand must be held in position for approximately 90 seconds to dislodge ice from thepropeller surface. Repeat this procedure as required to avoid significant buildupof ice which will result in loss of performance, vibration, and impingement upon thefuselage. The prop deice ammeter will not indicate a load while the propeller deicesystem is being utilized in the manual mode. However, the loadmeters will indicatean approximate 5% increase of load per meter while the manual prop deice sys-tem is operating.

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PITOT MAST HEAT Heating elements are installed in the pitot masts located on the nose. Each heat-ing element is controlled by an individual circuit breaker switch placarded PITOT- LEFT - RIGHT, located on the pilot’s right subpanel. A failure is indicated by illu-mination of the L PITOT HEAT or R PITOT HEAT annunciator, located on the cau-tion/advisory panel. Illumination of these annunciators indicates that pitot mastheat is inoperative. The annunciators will also illuminate anytime the PITOTswitches are in the OFF position. It is not advisable to operate the pitot heat sys-tem on the ground except for testing or for short intervals of time to remove ice orsnow from the masts.

SURFACE DEICE SYSTEM The surface deice system removes ice accumulations from the leading edges ofthe wings and horizontal stabilizers. Ice removal is accomplished by alternately in-flating and deflating the deice boots. Pressure-regulated bleed air from the en-gines supplies pressure to inflate the boots. A venturi ejector, operated by bleedair, creates vacuum to deflate the boots and hold them down while not in use. Toassure operation of the system in the event of failure of one engine, a check valveis incorporated in the bleed air line from each engine to prevent loss of pressurethrough the compressor of the inoperative engine. Inflation and deflation phasesare controlled by a distributor valve.

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

A three-position switch in the ICE group on the pilot’s subpanel identified: DEICECYCLE - SINGLE - OFF - MANUAL, controls the deicing operation. The switch isspring-loaded to return to the OFF position from SINGLE or MANUAL. When theSINGLE position is selected, the distributor valve opens to inflate the wing boots.After an inflation period of approximately 6 seconds, an electronic timer switchesthe distributor to deflate the wing boots, and a 4-second inflation begins in the hor-izontal stabilizer boots. When these boots have inflated and deflated, the cycle iscomplete.

When the switch is held in the MANUAL position, all the boots will inflate simulta-neously and remain inflated until the switch is released. The switch will return tothe OFF position when released. After the cycle, the boots will remain in the vac-uum hold-down condition until again actuated by the switch.

For most effective deicing operation, allow at least 1/2 inch of ice to form beforeattempting ice removal. Very thin ice may crack and cling to the boots instead ofshedding. Subsequent cyclings of the boots will then have a tendency to build upa shell of ice outside the contour of the leading edge, thus making ice removal ef-forts ineffective.

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PNEUMATIC BLEED AIR SYSTEMAND

SURFACE DEICE SYSTEM SCHEMATIC

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STALL WARNING VANE HEAT The lift transducer is equipped with anti-icing capability on both the mounting plateand the vane. The heat is controlled by a switch in the ice group located on thepilot’s right subpanel identified: STALL WARN. The level of heat is minimal forground operation, but is automatically increased for flight operation through theleft landing gear safety switch.

The heating elements protect the lift transducer vane and face platefrom ice. However, a buildup of ice on the wing may change or dis-rupt the airflow and prevent the system from accurately indicating animminent stall. Remember that the stall speed increases wheneverice accumulates on any airplane.

FUEL HEAT An oil-to-fuel heat exchanger, located on the engine accessory case, operatescontinuously and automatically to heat the fuel sufficiently to prevent ice from col-lecting in the fuel control unit when the procedure under “APPROVED FUEL AD-DITIVES” located in Section 2, LIMITATIONS, is complied with.

BRAKE DEICE SYSTEM (IF INSTALLED)High temperature engine compressor bleed air is directed onto the brake assem-blies by a distributor manifold on each main landing gear. This high pressure airis supplied by the standard bleed air pneumatic system which also provides reg-ulated pressure to the surface deice system and vacuum source. High tempera-ture air from the pneumatic system is routed through a solenoid control valve ineach main wheel well through a flexible hose on the main gear strut, and to thedistribution manifold around the brake assembly.

A switch on the pilot’s subpanel, placarded BRAKE DEICE, controls the brake de-ice system. When this switch is activated, both solenoid control valves are openedand the BRAKE DEICE ON annunciator on the caution/advisory panel is illuminat-ed to advise the system is in operation.

The brake deice system may be operated as required on a continuous basis withthe landing gear extended, provided the appropriate limitations are observed. Toavoid excessive wheel well temperatures with the landing gear retracted, a timeris incorporated to automatically terminate system operation approximately tenminutes after the landing gear is retracted. The system annunciator should bemonitored and the control switch selected OFF when the annunciator extinguish-es, or if brake deice operation has not automatically terminated within approxi-mately ten minutes. The landing gear must be extended before the timer is resetand the system can be activated again.

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Brake deice valves will give better service and have longer service life if they arecycled on some regular basis throughout the year. Long periods of valve inactivitycan cause accumulation of contaminants around the valve mechanism, causing itto seize, or preventing complete closure. The brake deice system should be cy-cled as part of the maintenance engine runups, and once during each flight oper-ation. Besides keeping the valves functional, this procedure will give you theassurance that the system will work properly when it is needed.

COMFORT FEATURES SIDE FACING TOILET (B200GT) A side facing electrically flushing toilet, when installed, is located in the foyer andfaces the airstair door. The foyer can be closed off from the cabin by sliding thetwo partition-type door panels to the center of the fuselage, where they are heldclosed by magnetic strips. The bottom cushion must be raised to gain access tothe toilet. A toilet tissue dispenser is contained in the seat shroud.

FRONT FACING TOILET (B200CGT)A front facing toilet, when installed, is located in the aft cargo area and is enclosedby the cargo partition. The toilet may be either the chemical type or the electricallyflushing type. In either case, the bottom cushion must be raised to gain access tothe toilet. A toilet tissue dispenser is provided in the top of the cabinet.

If a Monogram electrically flushing toilet is installed, the sliding knifevalve should be open at all times, except when actually servicing theunit. The cabinet below the toilet must be opened in order to gain ac-cess to the knife valve actuator handle.

RELIEF TUBES A relief tube is provided in the seat shroud of the side facing toilet (B200GT), oron the baggage compartment wall forward of the front facing toilet (B200CGT). Anoptional relief tube may be installed in the cockpit and stowed under the pilot’schair. A valve lever is located on the side of the relief tube horn. This valve levermust be depressed at all times while the relief tube is in use.

NOTEThe relief tubes are for use during flight only.

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CABIN FEATURES CABIN AC POWER OUTLETSThe airplane is equipped with four AC power outlets to provide 115-vac to be usedto supply power for laptop computers. The outlets are located on each side of thecabin in the lower sidewalls above the main spar and beneath the cabin tables.Lift the cover placarded 115 VAC to access the outlets.

One 115-volt, 60-Hz inverter provides the power for the outlets. The inverter is lo-cated in the right center section wing just outboard of the nacelle. Input power (28-vdc) comes from the right generator bus through a circuit breaker placarded IN-VERTER located on an electrical equipment panel under the floor on the copilot’sside of the cockpit. The inverter output is protected with a circuit breaker placarded115 VAC - 5 AMP adjacent to the inverter.

The Furnishings switch in the copilot’s inboard subpanel controls the inverter inputpower. The switch is placarded FURN ON - COFFEE OFF - OFF. The inverter op-erates when the Furnishings switch is in the FURN ON or COFFEE OFF position.

For normal operation, input current to the inverter can vary from approximately 0.5amperes to approximately 20 amperes depending on the load connected to theoutlets. The inverter is capable of providing a continuous output of 4 amperes. Thetotal electrical load connected to the four outlets must not exceed 4 amperes. Ex-cess load may cause the inverter input circuit breaker to open.

The inverter will shutdown for input over voltage, under voltage and high internaltemperature conditions. It will automatically reset when the conditions are correct-ed. The inverter will also shutdown for an output short circuit. Following a short cir-cuit shutdown, the inverter can be manually reset by turning the Furnishingsswitch OFF and ON.

The outlets should only be used when the generators or external power are on-line. Using the outlets without the generators or external power on-line could de-plete the airplane battery. In addition, the higher input current resulting from lowvoltage operation could cause the inverter input circuit breaker to open.

FIRE EXTINGUISHERS An optional portable fire extinguisher may be installed on the floor on the left sideof the airplane forward of the airstair entrance door, just under the rearmost seat.Another one may also be installed underneath the copilot’s seat.

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WINDSHIELD WIPERS The dual windshield wiper installation consists of an electric motor, arm and wiperassemblies, drive shafts, and converters, all located forward of the instrumentpanel. The system also includes a control switch, located in the upper left cornerof the overhead light control panel, and a circuit breaker located on the right circuitbreaker panel. The control knob, placarded WINDSHIELD WIPER - PARK - OFF- SLOW - FAST, controls the wipers. The wipers have two speeds, one for lightand one for heavy precipitation. After the control is turned to PARK to bring thewiper arms to their most inboard position, spring-loading returns the control to theOFF position.

Windshield wipers may be used during either ground or flight operations.

Do not operate windshield wipers on dry glass.

CARGO RESTRAINT (B200CGT) Hawker Beechcraft Corporation offers an FAA approved cargo restraint system asKit No. 101-5040. Any other restraint system used in this airplane must be ap-proved by the FAA. Such approval is the sole responsibility of the owner/operatorof the airplane.

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AVIONICSELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS)The system consists of three 8” x 10” color composite Adaptive Flight Displays(AFD). These AFD’s are provided as two Primary Flight Displays (PFD) and oneMultifunction Display (MFD).

Each PFD displays airplane attitude, heading, airspeed, altitude, vertical speed,flight guidance system annunciations and navigation data on a single integrateddisplay. The PFD also provides engine display information when selected in re-versionary mode.

The upper area of the PFD is used to present the basic “T” instruments, an AttitudeDirector Indicator (ADI), altitude scale, airspeed scale and vertical speed scale.Flight guidance system mode information displays in the area above the ADI.

The lower area of the PFD is used to present a Horizontal Situation Indicator (HSI)with a full or partial compass arc or map format, as selected by the pilot. Weatherradar or EGPWS information can be overlaid on the partial arc format or map for-mat. The space to either side of the HSI format is used to present a lateral navi-gation data field, a weather radar mode field, EGPWS mode field, systemmessages and selected menus.

The area along the bottom of the PFD is used to present radio tuning, time andtemperature displays. Normal control, reversion and warning annunciations arealso presented. In the case of a failed AFD, either the PFD or the MFD can bemanually reverted to a composite MFD/PFD format. This format presentation in-cludes Engine Indicating System (EIS) displays across the top of the format andthe basic “T” information presented below.

The MFD can be used to present a variety of information, including: Present Po-sition Map, TCAS, and FMS based textual data, navigation data, weather radarand EGPWS. Engine data and the electronic checklist are also presented on theMFD.

Line select keys are provided on each side of the displays and are used to controlthe basic display formats. The bezel mounted line select keys along with the Dis-play Control Panel (DCP) and Flight Guidance Panel (FGP) provide primary pilotinterface to control the PFD and MFD. Control of the radar, NAV sources andbearing pointers is through the DCP and the PFD line select keys. Control of thecourse, preselect heading, altitude and speed references is through the FGP.

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AFD display dimming is provided through an External Dimming control located onthe overhead panel, as well as a Trimmer located on each AFD. The following pro-cedure is to be used to set up the minimum display brightness for each AFD.

1. Cover the AFD light sensor located on the lower left corner of the displayand keep covered during this procedure. Wait 20 seconds before proceed-ing.

2. Adjust the External Dimming control to the full dim position.3. Adjust the Trimmer on the AFD until the display is barely discernible.4. Adjust the External Dimming control to achieve the desired level of bright-

ness.

NOTEIn extremely dark or bright ambient conditions, External Dimmingcontrol adjustment may be end limited prior to reaching the desireddisplay intensity. In this event, the Trimmer(s) should be used tomake the final intensity adjustment.

AIR DATA SYSTEM (ADS)Dual ADSs sense and process data obtained from the air mass around the air-plane. The two Air Data Computers (ADC) connect to the pitot and static air inputports and to a temperature sensor. The ADCs process air data and provide outputparameters to the PFD/MFD, AHCs and the IAPS concentrators.

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AIRCELL ST 3100 SATELLITE TELECOMMUNICATION SYS-TEM (IF INSTALLED)The AirCell ST 3100 system provides worldwide voice and data communicationusing the Iridium satellite network of 66 low-earth orbit satellites and operates inthe frequency range of 1616 to 1625.5 MHz. Communications are available on theground, provided line-of-sight is maintained with the satellites, and in the air.

NOTEThis system is not considered a portable system and thus 14 CFRPart 91.21, Portable Electronic Devices, does not apply. In addition,the FCC has issued a limited waiver which exempts the system fromthe requirements of 47 CFR Part 22.925.

The AirCell ST 3100 system includes the following equipment:• An AirCell Satellite Transceiver with RS 232 Modem installed in the aft avi-

onics compartment• An AirCell Cordless Handset mounted on a base station in the cabin• A serial port located next to each table in the cabin to support laptop com-

puters• A Northern Airborne Technology PTA12-100 POTS Telephone Adapter, lo-

cated in the cockpit pedestal• An Iridium Patch antenna, located on top of the fuselage• Two circuit breakers (TEL and DIALER) located on the right circuit breaker

panel

The pilot’s and copilot’s audio panels include a TEL position on the mic selectswitch and a TEL audio ON-OFF switch and volume control. These controls, alongwith the Telephone Adapter, allow the pilot and copilot to make and receive callsin the cockpit using headsets or speakers, and boom mics or hand mics. In addi-tion, the crew can join a telephone conversation that was initially received in thecabin.

Power to the transceiver and base station for the cordless phone is provided bythe No. 2 Avionics through the 7 1/2-amp TEL circuit breaker. Power to the Tele-phone Adapter is provided by the same Avionics bus through the 1-amp DIALERcircuit breaker. Power to both units is controlled by the Avionics Master switch.

Lighting to the Telephone Adapter is controlled by the Master Panel Lights switchand the OVHD PED & SUBPANEL rheostat. The brightness of the display is con-trolled by the BRT key on the Telephone Adapter.

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TELEPHONE ADAPTER

The following is a basic description of the controls on the Telephone Adapter.

CONTROL DESCRIPTIONHOOK Key This key is pressed to initiate or receive a call. Each time the

key is pressed the hook switch toggles on and off. When theswitch is “off-hook” the condition is similar to lifting a telephonehandset off the hook, and is indicated by the illumination of thegreen LED above the HOOK key and the message Call Modeshown on the display. The push-to-talk switches on the controlwheels are not required to transmit over the telephone.

HOOK LED (green)

When the Telephone Adapter is inactive, or on-hook, the LEDwill blink to indicate an incoming call. When the HOOK key ispressed to answer a call, the LED will illuminate steady.

HOLD Key The HOLD key is used to mute the mic and audio when theTelephone Adapter is active. Each push of the key cyclesbetween the Mute Mode and the un-muted mode.

HOLD LED (green)

The LED illuminates steady when the HOLD key is pressed tomute the mic and audio during a call.

REDIAL Key This key is used to recall and redial the last number dialed.The last number is stored in non-volatile memory and thus willbe retained when the system is powered down.

FLASH Key This key has no function on this installation

Key

This key increases () and decreases () the output volumeof the phone when the HOOK switch is active or when theRinger Detect circuit is active (i.e. the phone is ringing.)Volume increments can be adjusted from 1 (lowest) to 32(highest). When the HOOK switch is inactive, the key is usedto enter the Recall Mode and cycles through the 16 (01 - 16)available addresses which can be programmed with namesand telephone numbers.

BRT Key This key is used to control the brightness of the display.Brightness levels from 1 (dimmest) to 8 (brightest) areavailable.

ENTR Key Pressing the ENTR key sends a number dialed from thedefault or Recall Mode. It is also used to select the Edit Mode.This mode is used to program up to 16 names and telephonenumbers.

ESC Key In the Default Mode this key will cancel the current key activity.In the Edit Mode pressing the key returns the system to theDefault Mode.

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BASIC AUDIO PANEL OPERATION1. Interphone Control - This control may be either on or off when using the tele-

phone.2. Speaker Control - If the speakers are selected on, the telephone audio will

be heard over the speakers.3. AUTO COMM Switch - This switch interacts with the TEL position of the

Transmit switch and the TEL audio control as follows:a. With the AUTO COMM Switch on, selecting the TEL position of the

Transmit switch will take the telephone Off Hook unless the other side ofthe cockpit has both the TEL mic select switch and AUTO COMM Switchon.

b. Deselecting the TEL position of the Transmit switch or turning the AUTOCOMM Switch off will place the telephone back ON HOOK (disconnectthe phone call) unless the opposite side has selected the TEL transmitswitch and the AUTO COMM Switch.

c. With the pilot and copilot AUTO COMM Switches off, the mic key on thecontrol wheel or the hand mic can be used to transmit over the tele-phone. This mode of operation is not recommended.

• The TEL audio control must be on.• The mic key must be continuously held down.• If the mic key is released, the phone will hang up.• With either AUTO COMM Switch on, neither the pilot’s nor the copilot’s

mic switch will activate the telephone Hook switch.d. Transmissions cannot be made on the phone unless the AUTO COMM

Switch is selected on either the pilot’s panel or the copilot’s panel. Withat least one switch selected on, both pilot and copilot can transmit on thephone. Neither mic switch is required to talk, even on the side with theAUTO COMM Switch off.

4. When the telephone has been taken Off Hook using the TEL transmit switchor AUTO COMM Switch, the telephone can be placed ON Hook by pressingthe HOOK key on the telephone adapter.

5. If either crew member wishes to leave a phone conversation without discon-necting a call, the other crew member should ensure that the TEL mic switchand AUTO COMM Switch are selected.

6. The Interphone switch does not have to be on for both crew members to talkon the phone.

7. The TEL Audio Control and the master volume control (located on the micselect switch) may be used to modulate the output volume set on the tele-phone adapter. The master volume control cannot increase the volumeabove the level set on the Telephone Adapter.

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USING THE TELEPHONE ADAPTER

The following abbreviated procedures can be used to operate the Northern Air-borne Technology PTA12-100 (POTS) Telephone Adapter. For more detailed in-formation refer to the PTA12-100 POTS Telephone Adapter Operator’s Manual,obtained from Northern Airborne Technology Ltd.

TO MAKE A CALL IN THE CALL MODE1. Mic Select Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .TEL2. AUTO COMM Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON

• If the opposite side has the TEL mic switch and the AUTO COMM Switchselected on, the above two steps will not take the phone OFF HOOK.

3. Hook Key (if required) . . . . . . . . . . PRESS TO TAKE PHONE OFF HOOK• The green HOOK LED will illuminate• Call Mode will be displayed• A dial tone will be heard

4. Type in Phone Number• 1 + Area Code + Telephone Number

(US and Canada)• 00 + Country Code + Area Code + Telephone Number

(Outside US and Canada)5. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS

• Phone Number will be dialed• A series of short tones will be heard• The number called will ring

6. Use Boom mic (or hand mic) and headsets (or speakers) to communicate7. HOOK Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .PRESS TO END CALL

TO MAKE A CALL IN THE RECALL MODE1. Mic Select Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .TEL2. AUTO COMM Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON

• If the above steps take the phone OFF HOOK, press the HOOK key on thetelephone adapter to place the phone back ON HOOK.

3. Key in valid address for a stored number (01 - 16) or press the or keysto scroll through the list.

4. HOOK Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS• The green HOOK LED will illuminate• The phone number will be displayed• A dial tone will be heard

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5. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS• Phone Number will be dialed• A series of short tones will be heard• The number called will ring

6. Use Boom mic (or hand mic) and headsets (or speakers) to communicate7. HOOK Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS TO END CALL

TO MAKE A CALL USING THE REDIAL KEY1. Mic Select Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TEL2. AUTO COMM Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON

• If the above steps take the phone OFF HOOK, press the HOOK key on thetelephone adapter to place the phone back ON HOOK.

3. REDIAL Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS• Last number called will be displayed

4. HOOK Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS• The green HOOK LED will illuminate• The phone number will be displayed• A dial tone will be heard

5. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS• Phone Number will be dialed• A series of short tones will be heard• The number called will ring


TO RECEIVE A CALL

The green HOOK LED will blink and the ringer will sound in the headset if the TELaudio switch is pulled on. If the speaker(s) are on, the ringer will sound throughthe speaker(s).

1. Mic Select Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . TEL2. AUTO COMM Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .ON3. HOOK Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .PRESS (if required)

• The green HOOK LED will steadily illuminate• Call Mode will be displayed


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EDIT MODE

NOTEPressing the ESC key during any portion of the Edit Mode will returnthe Telephone Adapter to the Default Mode indicated by the displayPTA12-100 X.XX.X.

Storing A New Name And Phone Number1. ESC Key (if required) . . . . . . . PRESS TO RETURN TO DEFAULT MODE2. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS

Edit Mode message appears for 2 - 3 sec followed by or Address message

3. Select empty address by:a. Pressing the or keys to select address (01 - 16) or,b. Keying in address directly (01 - 16)c. Empty! message indicates empty address

4. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS Enter Name message is displayed

5. Press the appropriate key repeatedly until the desired letter or number ap-pears. For example, pressing the “3 DEF” key once = D, twice = E, threetimes = F and 4 times = the number 3.

6. To correct an entry, use the or keys. This will activate a cursor whichcan be placed over the character that requires a change.

7. To enter a space, press the “0 Space” key.8. If two successive letters and/or numbers are on the same key, the key will

need to be pressed to advance to the next space before entering the nextletter/number.

9. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS TO STORE NAME Enter Number message is displayed

10. Press the appropriate keys to enter the phone number using the above pro-cedures. (Alpha characters are not active in this mode. Press each key onlyonce to enter a number. Press the “0 Space” key twice to enter a space.)

11. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS TO STORE NUMBER Stored! message is displayed

12. ESC Key. . . . . . . . . . . . . . . . . PRESS TO RETURN TO DEFAULT MODE

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Editing An Occupied Address1. ESC Key (if required) . . . . . . . PRESS TO RETURN TO DEFAULT MODE2. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS


3. Select desired address by:a. Pressing the or keys to select address (01 - 16) or,b. Keying in address directly (01 - 16)

4. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS Del or Edit message is displayed

5. Press the key to edit the name in a selected address. The address nameis displayed.

6. Press the key to place the cursor over the character to be edited, thenpress the desired key to change the character.

7. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS TO STORE NAME The telephone number will be displayed

8. Press the key to place the cursor over the number to be edited, then pressthe desired key to change the number.

9. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS TO STORE NUMBER Stored! message is displayed

10. ESC Key . . . . . . . . . . . . . . . . . PRESS TO RETURN TO DEFAULT MODE

Deleting An Occupied Address1. ESC Key (if required) . . . . . . . PRESS TO RETURN TO DEFAULT MODE2. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS


3. Select desired address by:a. Pressing the or keys to select address (01 - 16) or,b. Keying in address directly (01 - 16)

4. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS Del or Edit message is displayed

5. Press the key to delete the name and number in the selected address. ENTR to Delete message is displayed

6. ENTR Key . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS Deleted! message is displayed

7. ESC Key . . . . . . . . . . . . . . . . . PRESS TO RETURN TO DEFAULT MODE

For more detailed information, refer to the PTA12-100 POTS Telephone AdapterOperator’s Manual, obtained from Northern Airborne Technology Ltd.

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AIRCELL CORDLESS HANDSET

The following are abbreviated procedures for using the AirCell Cordless Handset.Refer to the AirCell ST 3100 Satellite Telecommunication System Users Manual,P/N 810-10680, Rev. F or later revision, for information concerning the cordlesshandset installed in the cabin.

TO MAKE A CALL1. Remove handset from cradle.2. Verify Iridium appears in the display.3. Verify the message No Serv does not appear in the display.4. Dial Telephone Number

• 1 + Area Code + Telephone Number(US and Canada)

• 00 + Country Code + Area Code + Telephone Number(Outside US and Canada)

5. Press the SND key.6. In Use message appears and Iridium disappears in the display.7. Volume can be adjusted using the or key.8. Press the END key to end the call.

TO JOIN A CALL MADE FROM THE COCKPIT1. Remove handset from cradle.2. Verify Iridium appears in the display.3. Mute message appears in the display.4. Press the FCN key followed by the 7 key to unmute the handset.5. Press the SND key.

TO RECEIVE A CALL1. Remove handset from cradle.2. Press the SND key to answer the phone.3. In Use message appears in the display.4. Volume can be adjusted using the or key.5. Press the END key to end the call.

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AIRSPEED SCALEOverspeed Alerts

The VMO/MMO marker is a solid red bar that is constantly positioned from the topof the airspeed scale to the current value of VMO/MMO. The marker becomes em-phasized and the airspeed digital readout changes color if the airspeed trend vec-tor exceeds VMO/MMO (Overspeed Pre-alert) and/or the airspeed exceeds VMO/MMO (Overspeed Alert). These changes occur as follows:

The overspeed pre-alert will occur when the airspeed trend vector exceeds VMO/MMO by 3 knots for 5 seconds or more. The overspeed pre-alert symbology con-sists of the following indications.

1. The VMO/MMO marker will be emphasized by changing to a wide, red-out-lined, red-and-black checkerboard pattern.

2. The airspeed digital readout will change color to yellow, flash for 5 seconds,then become steady.

3. If the MACH readout is displayed, the readout will change color to yellow,flash for 5 seconds, then become steady.

4. The overspeed pre-alert symbology will cease when the airspeed trend vec-tor is less than or equal to VMO/MMO.

The overspeed alert will occur when the current airspeed exceeds VMO/MMO by 3knots. The overspeed alert symbology and aural alert consists of the following in-dications.

1. The PFD’s overspeed aural alert will sound.2. The VMO/MMO marker will be emphasized by changing to a wide, red-out-

lined, red-and-black checkerboard pattern.

NOTEActivation of the overspeed pre-alert will typically have alreadyplaced the checkerboard pattern into view.

3. The airspeed digital readout will change color to red, flash for 5 seconds,then become steady.

4. If the MACH readout is displayed, the readout will change color to red, flashfor 5 seconds, then become steady.

5. The overspeed alert symbology and the aural alert will cease when the air-speed is less than or equal to VMO/MMO.

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Low Speed Warnings

The Impending Stall Speed (ISS) Low Speed Cue (LSC) Marker is a solid red barthat is constantly positioned from the bottom of the airspeed scale to the currentheavy weight stall speed as defined by the existing flap position. The ISS LSC isadvisory in nature and is not a substitute for the stall warning system. The ISS LSCis not displayed until the weight-on-wheels switch is in the “air” position for at least6 seconds. The following table defines the ISS LSC value for each flap position.

When a new flap position is selected, the ISS LSC value will change to the newISS LSC value at the following rates:

1. When the flap lever is selected to an extend position (i.e. from UP to AP-PROACH, or from APPROACH to DOWN) the rate-of-change will be 1.5knots per second.

2. When the flap lever is selected to a retract position (i.e. from DOWN to AP-PROACH, or from APPROACH to UP) the rate-of-change will be 3.5 knotsper second.

The ISS LSC pre-warning occurs when the airspeed trend vector extends belowthe current ISS LSC value for at least 5 seconds. The ISS LSC pre-warning sym-bology consists of the following indications.

1. The ISS LSC marker will be emphasized by changing to a wide, red-out-lined, red-and-black checkerboard pattern.

2. The airspeed digital readout will change color to yellow, flash for 5 seconds,then become steady.

3. If the MACH readout is displayed, the readout will change color to yellow,flash for 5 seconds, then become steady.

4. The ISS LSC pre-warning symbology will cease when the airspeed trendvector is greater than or equal to the ISS LSC value.

The ISS LSC warning will occur when the current airspeed becomes less than theISS LSC value. The ISS LSC warning symbology consists of the following indica-tions.

1. The ISS LSC marker will be emphasized by changing to a wide, red-out-lined, red-and-black checkerboard pattern.

NOTEActivation of the pre-warning will typically have already placed thecheckerboard pattern into view.

Flap Position ISS LSC (KIAS)UP 99

APPROACH 85DOWN 75

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2. The airspeed digital readout will change color to red, flash for 5 seconds,then become steady.

3. If the MACH readout is displayed, the readout will change color to red, flashfor 5 seconds, then become steady.

4. The ISS LSC warning symbology will cease when the airspeed is greaterthan or equal to the ISS LSC value.

If the flap lever input becomes invalid, a default ISS LSC is displayed. This displayconsists of the checkerboard bar with a yellow bar on top. The junction of thecheckerboard bar with the yellow bar represents the flaps-down ISS LSC value,while the top of the yellow bar represents the flaps-up ISS LSC value. The ISSLSC warning function will continue to work as previously described, except the de-fault ISS LSC marker will already be in view.

ATTITUDE HEADING SYSTEM (AHS)The AHS is a dual reference system consisting of two AHC-3000 Attitude HeadingComputers (AHC), two FDU-3000 Flux Detector Units (FDU) and two ExternalCompensation Units (ECU).

The AHS supplies attitude, stabilized magnetic or free gyro heading and linear ac-celeration data to the Flight Guidance System, Electronic Flight Instrument Sys-tem, Integrated Avionics Processor System and Weather Radar System.

The AHCs are functionally and physically isolated from each other and replace theconventional vertical gyro, directional gyro, three rate gyros, and three linear ac-celerometers. The AHS has two operation modes, slaved and Directional Gyro(DG) mode.

CONTROL DISPLAY UNIT (CDU)A CDU is installed in the pedestal. The CDU is a color LCD-based display unit withan integrated keyboard having 16 keys and a full alphanumeric keypad. In additionto the integrated keyboard, the CDU has six line select keys located in the bezelon each side of the color LCD display.

The pilot’s CDU is used to control the Radio Sensor System (RSS) and provideintegrated control of several combinations of airplane communications (includinga normal means of radio tuning) and navigation radio subsystems.

The CDU also provides the following functions:• Control display from the Flight Management System (FMS).• Back-up display for Global Positioning System (GPS) data.• MFD menus

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DISPLAY CONTROL PANELS (DCP)Two dedicated DCPs control the display and menu functions of the PFDs/MFD.

The DCPs are located adjacent to the respective PFD/MFD they normally control.The DCP, when combined with the line select keys on the PFD, provides controlof the weather radar, NAV sources, bearing pointers, V speeds, BARO minimums(Barometric Altitude-based Minimum Descent Altitude) and RA Minimums (RadioAltitude-based Decision Height Minimums).

EMERGENCY LOCATOR TRANSMITTER

The Artex C406-2 Emergency Locator Transmitter (ELT) System is designed to meetthe requirements of TSO C91a and C126. The system consists of the ELT transmitterand an alert horn located in the aft fuselage area, an antenna mounted on the aft fuse-lage, and a remote switch with a yellow transmit light located on the left cockpit sidewallnext to the OAT gage. The purpose of the alert horn is to notify personnel that the ELThas been activated. The remote switch is lever-locked in the ARM and the ON posi-tions. Neither this switch nor the switch on the ELT transmitter can be positioned to pre-vent the automatic activation of the ELT transmitter. The system is independent fromother airplane systems except for the transmit light, which is hot-wired to the airplanebattery, and the edge lit panel which is controlled by a rheostat switch on the overheadpanel placarded SIDE PANEL - OFF - BRT.

Upon activation, the ELT will sound the alert horn and transmit a sweeping tone on121.5 and 243.0 MHz. An additional frequency of 406.025 MHz is also transmitted,which is used by orbiting satellites to assist in determining airplane location. This acti-vation is independent of the remote switch setting or availability of airplane power. Theremote switch is installed to perform the following functions:

• Test the ELT.• Deactivate the ELT, if it has been inadvertently activated by the “G” switch.• Activate the ELT in an in-flight emergency if an off-airport landing is antici-

pated.• Activate the ELT after an off-airport landing, if the impact did not automati-

cally activate it.

The ELT should be tested every twelve months. The test consists of turning theunit on and then resetting it using the following procedures:

• Tests should be conducted between the times of on-the-hour until 5 minutesafter the hour.

• Notify any nearby control towers.

NOTEDo note allow the test to exceed 15 seconds. The satellite systemrecognizes 406.025 MHZ transmissions in excess of 15 seconds tobe valid distress signal.

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• Provide power to an airplane radio and tune it to 121.5 MHz.• Place the ELT remote switch to ON. Wait for at least 3 sweeping tones on

the airplane radio, which will take about 1 second, then return the switch toARM.

• The test is successful if the sweeping tones are heard and the transmit lightnext to the switch illuminates immediately. If there is a delay in the illumina-tion of the transmit light, the system is not working properly.

If the ELT should be inadvertently activated by the “G” switch, the transmit lightnext to the switch will blink. The ELT can be deactivated by momentarily placingthe remote switch ON and then back to ARM.

For test procedures on the 406.025 MHZ frequency, refer to the Artex Installationand Operation Manual.

ENGINE INDICATING SYSTEM (EIS)The EIS digitizes airplane engine data for input to the avionics system while con-verting avionics digital data into outputs that drive airplane annunciators and auralhorns.

The EIS consists of four Data Concentration Units (DCUs). Two units are wired asDCUs and two units are wired as Engine Data Concentrators (EDCs).

The DCUs transfer airplane sensor information (analog fuel flow, strut status, etc.)to the IAPS concentrators while transferring caution/warning advisory informationfrom the Flight Control Computers to the annunciators. The EDUs also provide re-dundant engine data to the displays.

ENHANCED GROUND PROXIMITY WARNING SYSTEM (EGP-WS) (IF INSTALLED)Refer to the following pilot’s guides:

Honeywell EGPWS Mark VI and MK VIII Enhanced Ground Proximity WarningSystem Pilot’s Guide, P/N 060-4314-000, Revision B, dated February 2002, or lat-er revision.

Operator’s Guide, Collins Pro Line 21 Avionics System for the Hawker BeechcraftCorporation King Air, P/N 523-0790065-003117, dated August 27, 2003, or laterrevision.

The following information supplements those pilot’s guides:

The EGPWS is powered from the No. 4 Dual-Fed Bus and is protected by a 3-ampcircuit breaker, placarded TAWS, located on the copilot’s circuit breaker panel.

The following equipment must be operational for the proper functioning of Modes1 through 6 of the GPWS:

1. Enhanced Ground Proximity Warning Computer (EGPWC)2. Radio Altimeter

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3. Vertical Speed from the Air Data Computer4. Airspeed from the Air Data Computer5. Glideslope Deviation6. Landing Gear position7. Flap position8. Roll attitude from Pilot’s Attitude System (for “BANK ANGLE” voice mes-

sage)9. Decision Height System (for “MINIMUMS” voice message)

The following equipment must be operational for the proper functioning of the en-hanced features of the EGPWS.

1. Enhanced Ground Proximity Warning Computer (EGPWC)2. Heading from the No. 1 Compass System3. GPS position from the Flight Management System4. Terrain and Obstacle Database

The following enhanced features are available:1. Terrain (or obstacle) Alerting and Display TAD. The display is configured for

the Peaks Display option and the Pop-Up feature (MFD only) option.2. Envelope modulation of GPWS Modes.3. Terrain Clearance Floor (TCF).4. Runway Field Clearance Floor (RFCF).

The following three switches for the GPWS and the EGPWS are located abovethe pilot’s DCP. When a switch is pressed to activate a function, the annunciatorACTIVE is displayed below the switch placard in the color noted.

Switch Color FunctionFLAPOVRD

ACTIVE AMBER

Pressing the switch disables the GPWS Mode4b alert, TOO LOW FLAPS, and desensitizesthe Mode 1 alert boundaries.

G/SINHIB

ACTIVE AMBER

Pressing the switch disables the Mode 5glideslope alert, GLIDESLOPE, anytime theairplane is below 2000 feet AGL, theglideslope flag is pulled, and the gear is down.(While the airplane is on the ground, thisswitch is used to initiate system self-test.)

TERRINHIB

ACTIVE GREEN

Pressing the switch deselects all enhancedfunctions of the EGPWS system.

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The following annunciators, voice alerts, and voice warnings are provided forModes 1 - 6 of the GPWS. See Section 3, EMERGENCY PROCEDURES, andSection 3A, ABNORMAL PROCEDURES, for the appropriate procedure associ-ated with each alert and warning.

* The “500” callout is configured as a Smart 500 callout.

Mode Function PFD Caution Message(Yellow)

Voice Alert PFD Warning Message

(Red)

Voice Warning

1 Excessive Descent Rate

GND PROX “Sink Rate” PULL UP “Pull Up”

2 Excessive Closure on

Terrain

GND PROX “Terrain, Terrain”

PULL UP “Pull Up”

3 Altitude Loss After Takeoff

GND PROX “Don’t Sink, Don’t Sink”

N/A N/A

4a Unsafe Terrain Clearance

GND PROX “Too Low Gear”

N/A N/A

5 Excessive Glideslope Deviation

GND PROX “Glideslope” N/A N/A

6 Excessive Bank Angle

N/A “Bank Angle” N/A N/A

6 Altitude Callouts N/A “500*, 200, 100, 50, 40, 30, 20, 10”

N/A N/A

6 Minimums Approach Altitude

(RA/BARO MIN)

N/A “Minimums, Minimums”

N/A N/A

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The following annunciators, voice alerts, and voice warning are provided for theEnhanced portion of the EGPWS. See Section 3, EMERGENCY PROCEDURES,and Section 3A, ABNORMAL PROCEDURES, for the appropriate procedure as-sociated with each alert and warning.

Function PFD Caution Message (Yellow)

Voice Alert PFD Warning Message

(Red)

Voice Warning

Terrain Alerting and Display

GND PROX “Caution Terrain, Caution Terrain”

PULL UP “Terrain, Terrain, Pull

Up”

Obstacle Alerting and Display

GND PROX “Caution Obstacle, Caution

Obstacle”

PULL UP “Obstacle, Obstacle, Pull Up”

Terrain ClearanceFloor GND PROX

“Too Low, Terrain”

N/A N/A

Runway Field ClearanceFloor

GND PROX“Too Low, Terrain”

N/A N/A

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ELECTRONIC STANDBY INSTRUMENT SYSTEM (ESIS)The Goodrich Model GH-3100 Electronic Standby Instrument System is locatedin the center of the instrument panel. The ESIS is a full-color, active matrix LCDwhich provides the following information:• Attitude• Altitude• Airspeed• Heading• Barometric Pressure• NAV 1 (NOTE: This information will be available until the airplane battery is de-

pleted. DME 1 will not be available if power is lost to the No. 3 Avionics Bus. Thisbus is manually shed after a dual generator failure.)

The system consists of the display, a standby battery, a magnetometer, and aconfiguration module. The display is protected by a 3-amp circuit breaker, plac-arded, ESIS DISP. A switch, placarded ESIS-ON-OFF-TEST, is located on the pi-lot’s outboard subpanel. Two indicator lights are installed immediately adjacent tothe switch. During normal operation, this switch is in the ON position and bothlights are extinguished. In the event that power is lost from the No. 3 dual-fed bus,the standby battery will provide power to the ESIS system causing an amber lightto illuminate. Illumination of this light indicates the ESIS system is using the stand-by battery. During operations with the amber light illuminated the standby batteryis not being recharged by the airplane system.

The ESIS standby battery can be tested prior to powering up the airplane by mo-mentarily holding the ESIS-ON-OFF-TEST switch in the TEST position. A greenlight will illuminate indicating a successful standby battery test.

The ESIS system does not incorporate an aural warning if the red VMO/MMO mark-er on the airspeed scale is exceeded.

The altitude display is corrected for Static System errors in the same manner asthe pilot and copilot altitude displays.

The altitude display is not approved for use in RVSM airspace.

Refer to the latest version of the Goodrich GH-3100 ESIS Pilot’s Guide P/N TP-560 for further information.

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FLIGHT GUIDANCE PANEL (FGP)The FGP is used to control the Flight Guidance System and is located in the centerof the glareshield.

Command of the Flight Guidance System is accomplished by using the lateral andvertical mode select switches, VS/pitch wheel, autopilot switches, FD switchesand various control knobs of the FGP along with the control wheel mounted syn-chronization (SYNC), autopilot disconnect (AP DISC) switches, and a left powerlever mounted go-around (GA) switch.

Attitude reference, heading reference, airspeed reference, vertical speed refer-ence and VS pitch reference are also controlled from the FGP.

FLIGHT GUIDANCE SYSTEM (FGS)The FGS provides autopilot and dual flight guidance functions by utilizing twoidentical computers, three primary servos, a pitch trim servo and a flight guidancepanel.

The Flight Guidance Computers (FGC) receive Attitude Heading System (AHS)data directly from the AHC to provide independent flight guidance computationwhile operating together to provide 3-axis autopilot, pitch trim and yaw damperfunctions.

The two FGSs apply differential autopilot command drive to each primary servo tomove the airplane elevator, aileron and rudder control surfaces.

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FLIGHT MANAGEMENT SYSTEM (FMS)The Flight Management System provides flight plan management, multisensornavigation, and radio tuning. The FMS consists of a Flight Management Computer(FMC), a CDU used to control the FMS and a Data Base Unit (DBU).

The FMC is a lateral and vertical navigator used by the autopilot to fly a pro-grammed flight plan and processes coupled VNAV, NAV-to-NAV capture, navaiddatabase storage, and several control/planning functions. The FMS will providenavigation solutions if it is receiving suitable navigation information from one ormore of the following:• One VOR/DME.• Two DMEs.• One GPS.

The DBU-4100 or DBU-5000 is a data loader used primarily to load monthly data-base updates to the FMC and load and download maintenance data from theMDC maintenance computer.

GROUND COMMUNICATIONS POWERThe ground communications electric bus provides electric power directly from thebattery when selected by the pilot. Control of the system consists of a push on/push off solenoid-held annunciator switch located on the instrument panel. Circuitprotection is provided by the GND COM circuit breaker. Activation of the systemallows operation of the comm system connected to the GND COM circuit breaker.

Audio is provided in the headphones, with speaker audio selectable. Subsequentactivation of the battery switch will result in automatic disconnection of the groundcommunications bus from the comm system; however, the normal method for de-activation of the system is accomplished by pressing the GND COM switch.

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HF-9000 SYSTEM (IF INSTALLED)The Collins HF-9000 High-Frequency Communications System provides a world-wide communications capability. The system consists of an HF-9031A Receiver-Transmitter and a HF-9041 Antenna Coupler, both located in the aft avionics com-partment. An HF wire antenna is located on the top of the fuselage and extendsback to the vertical stabilizer. All functions of the HF radio can be controlledthrough the pilot’s CDU and/or the Radio Tuning Unit (RTU). The HF radio is pow-ered from the No. 3 Avionics Bus. Two circuit breakers are located on the right cir-cuit breaker panel, one placarded HF COM (25 amps) to protect the receiver-transmitter, and one placarded HF ANT (5 amps) to protect the antenna coupler.Power is applied to the HF radio when the Avionics Master switch is turned on.

The HF radio operates in the HF band of 2.0000 to 29.9999 MHz in 100-Hz steps.Operating Emission modes include upper sideband voice (UV), lower sidebandvoice (LV) and amplitude modulation equivalent (AM). Both Simplex and Half-Du-plex Tuning Modes are available.

NOTEThe AM Emission Mode has a frequency band width of 15 KHz.Thus, radio stations with frequencies separated by 15 KHz or lessmay be received simultaneously.

Four HF Tuning Modes are available; a Frequency Tuning Mode, an EmergencyTuning Mode, a Maritime (ITU) Tuning Mode, and a Preset Tuning Mode. The Fre-quency Tuning Mode allows any valid frequency (either Simplex or Half-Duplex)and its emission mode to be manually set in the Active frequency window or theRecall frequency window of the RTU, or the Active frequency window of the CDU.The Emergency Tuning Mode consists of 6 emergency channels (stored in non-volatile memory) which may be set into the Active or Recall frequency window ofthe RTU, or the Active frequency window of the CDU. The Maritime Tuning Modeconsists of 249 ITU (International Telecommunications Union) public correspon-dence channels that are stored in non-volatile memory. These channels may alsobe set into the Active or Recall frequency windows of the RTU or the Active fre-quency window of the CDU. A listing of the available public correspondence chan-nels may be found in the Collins HF-9000 Operator’s Manual. The Preset TuningMode consists of twenty user-defined channels. The 20 available channels in theRTU are independent from the 20 available channels in the CDU. Thus, a total of40 different preset frequencies are available; 20 in the RTU and 20 in the CDU.The 20 channels in the RTU can only be filled with Simplex or Half-Duplex fre-quencies and their emission modes. Maritime and Emergency channels cannot bestored in this unit. The CDU can be filled with Simplex and Half-Duplex frequen-cies and their emission modes as well as Emergency channels and Maritimechannels.

A TEST mode is available for testing the system on the ground or in the air. Thetest may be conducted from the RTU or the CDU. No HF setting can be changedduring the test.

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Three levels of transmission power are available; LO (approximately 15 Watts av-erage), MED (approximately 30 Watts average), and HI (approximately 50 Wattsaverage.)

Because HF radio transmissions can interfere with ADF signal reception, the ADFreceiver freezes the existing bearing indication whenever the HF transmitter iskeyed.

The RTU contains three pages that can be used for control of the HF radio; theNEXT page of the Top Level Display Page, the HF Main Display page (HF) andthe HF Preset pages (HF).

The CDU also contains three pages that can be used for control of the HF radio;page 2 of the TUNE page, the HF1 Control Page and the HF1 Presets page.

The following table shows which HF features can be set on each page of the RTUand the CDU:

The following procedures are a basic guide to operating the HF radio using theRTU and the CDU. A short hand is used to denote key positions. For example, thethird Line Select Key on the Right side of the display will appear as LSK R3.

Function Top Level

TunePg. 2

Main Display

HF Control

Preset Preset

RTU CDU RTU CDU RTU CDUTune Active f X X X X X

Tune Recall f X X

Change Squelch X X X X

Change Emission Mode (UV, LV, AM)

X X X X X X

Change Transmission Power (LO, MED, HI)

X X

Access HF1 Preset page

X X

Conduct SELF TEST X X

Select Simplex or Duplex

X X X X X

Change Tuning Mode(FREQ, PRESET, EMER, MAR)

X X

Program Preset f’s X X

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RTU1. From the Top Level Display page, select the NEXT PAGE using LSK L4 to

access the Top Level HF Display Page which shows the Active and RecallHF frequencies.

2. The display of HF Active and Recall frequencies will vary depending on theTuning Mode of the HF radio (FREQ, PRESET, EMER, or MAR) and wheth-er or not Simplex or Half-Duplex frequencies are being used.a. FREQ Mode. In this mode frequencies from 2.0 to 29.9999 MHz may be

set in the Active or Recall display. If the Half-Duplex mode has been se-lected, both a receive (R) and a transmit (T) frequency may be set in ei-ther display.

b. PRESET Mode. In this mode a preset channel from 1 to 20 can be se-lected. A preset channel may contain a Simplex frequency, or a Half-Du-plex frequency. Emergency and Maritime channels can not beprogrammed into a preset channel of the RTU.

c. EMER Mode. In the Emergency Mode one of six emergency channelsmay be selected.

d. MAR Mode. In the Maritime Mode, one of 249 ITU public correspon-dence channels may be selected. (0401 - 429, 0601 - 0608, 0801 - 0837,1201 - 1241, 1601 - 1656, 1801 - 1815, 2202 - 2253, and 2501 - 2510.)

3. The Active and Recall HF frequencies and corresponding Emission modesmay be set on the Top Level HF Display page as well as the Squelch setting.For example, to set the active frequency window:a. Push LSK L1 to activate the tuning window.b. Rotate the large knob to position the window at the location where a

change needs to be made.c. Rotate the small knob to change the value.d. Rotate the large knob to the next position and repeat the process.e. The Squelch can be adjusted in increments from 0 to 3 using the small

knob.f. The Emission source can be set at UV, LV or AM using the small knob.

4. The Recall frequency and Emission mode is set on the Top Level HF Dis-play page in a similar manner.

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5. Press the LSK L1 twice (if tuning window has not been selected) to accessthe HF Main Display page.a. Use LSKs L1 and R1 and the control knobs to set Active and Recall HF

frequencies in the same manner used on the Top Level Display.b. Use LSK L2 to set the transmission power to LO, MED, or HI. Each press

of the key will cycle the radio to the next power level.c. Use LSK L3 to select the Preset page.d. Use LSK L4 to return to the Top Level Display page.e. Use LSK R2 to select the SIMPLEX or DUPLEX Mode. The DUPLEX

Mode refers to Half-Duplex operation where the transmit frequency is dif-ferent from the receive frequency. For this key to be active, the tuningwindow must be selected in either the Active frequency display by press-ing key L1, or selected in the Recall frequency display by pressing keyR1. Each press of key R2 cycles the radio between SIMPLEX and DU-PLEX.

f. Press and release LSK R3 to run the HF self-test. The white TEST leg-end will enlarge and change color to cyan. The Active frequency will bereplaced with the cyan message IN TEST, both on this page, the TopLevel Display page, and the Preset page. The test lasts approximately40 seconds. If successful, the page returns to normal. If unsuccessful, asix digit code will appear on the RTU. This code may be interpreted fromthe RTU installation manual, P/N 523-0780424-041116.

NOTEThe Active frequency will be inoperative after a self-test has been conducted. To reactivate the frequency, cycle it to the Recall win-dow, then back to the Active window.

6. Press the LSK L3 to select the Preset page.a. Seven Preset pages are available for setting up to 20 preset frequencies.

Press LSK R3 to select the tune window, then rotate either the large orsmall control knob to change the page number from 1 to 7.

b. Press LSK L1, L2, or R1 to activate the tuning window for the channel tobe preset.

c. Press LSK R2 to select either the SIMPLEX or DUPLEX mode for the fre-quency to be preset.

d. Then use the large and small knobs to set the frequency and Emissionmode.

e. The LSK L3 may be used to change the tuning mode of the radio (FREQ,PRESET, EMER, and MAR). Each press of the key selects the nextmode.

f. Press LSK L4 to return to the HF Main Display page.

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CDU1. Press the Tune Function Key to select the Tune page.2. Select page 2 of the Tune page by pressing the NEXT key.3. The active HF frequency is shown in the HF sub-display opposite LSK L4

along with the associated Emission mode.4. The Squelch setting is shown between LSKs 5 and 6. These LSKs may be

used to adjust the squelch from 0 to 3.5. The HF frequency may be tuned by entering it in the scratch pad and then

pressing LSK L4. The following procedures are used to enter the desiredfrequency or channel.a. Enter the frequency in MHz in the scratch pad using up to 6 digits, fol-

lowed by the desired Emission mode (e.g. 10.2345AM). The decimalpoint is optional if no decimal digits are required; however, if a decimalpoint is not used, add the letter F after the digit to distinguish is from anemergency channel.Example: 5F may be entered for the frequency 5.0000. (If the FREQ Tun-ing Mode of the radio has been selected, the F is not required.)

b. If only the Emission mode of the existing frequency needs to be changed,enter only the desired mode in the scratch pad (e.g. UV, LV or AM).

c. If a Half-Duplex operating mode is required, enter either the Transmittingor Receiving frequency, preceded by the letter T or R. Enter a space us-ing the SP key, then enter the second frequency preceded by T or R. TheEmission mode follows the last frequency with no space.Example: T10.1234(space)R15.1234UV.

d. When entering an Emergency or Maritime (ITU) channel, the letter E orM must follow the channel unless the Tuning Mode of the preset frequen-cy matches the existing Tuning Mode of the radio.For example, the following channels would be entered as shown: 6E =Emergency channel 6 and 408M = Maritime channel 0408.

6. Press LSK L4 to select the HF1 Control page.a. The active HF frequency and Emission mode is displayed adjacent to

LSK L1 and may be changed as described in item 5.b. The squelch may be adjusted using LSKs L3 and L4 as previously de-

scribed.c. The Tuning Mode may be changed using LSK L5. Each push of the LSK

cycles the Tuning Mode between FREQ, EMER, and MAR.d. The HF1 Presets page is selected using LSK L6.e. The transmission power of the HF radio is set using LSK R1. Each push

of the LSK cycles the power between LO, MED, and HI.

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f. Press LSK R2 to activate the self-test.1) The white TEST legend changes to cyan and enlarges.2) After approximately 14 seconds, the cyan TEST legend is replaced by

a green PASS message. This message will remain unless the fre-quency, power level, or Tuning Mode is changed. However, it is notnecessary to clear the PASS message in order to run another self-testor use the radio.

NOTEThe Active frequency will be inoperative after a self-test has beenconducted. To reactivate the frequency, cycle it to the Recall win-dow, then back to the Active window using the RTU.

7. Press LSK L6 to select the HF1 Presets page.a. The active frequency can be changed using LSK L1 as previously de-

scribed in step 5.b. Twenty preset frequencies and/or channels can be set on five preset

pages. Pages are selected using the NEXT or PREV keys.c. A preset frequency may be entered as a Simplex frequency, Half-Duplex

frequency, Emergency channel, or a Maritime channel using the proce-dures previously described.

d. An identifier can be associated with a preset frequency or channel. Theidentifier can be up to 12 characters long, must begin with a letter, butmay contain digits. If sufficient space is available in the scratch pad, theidentifier may be included as the last portion of the frequency entry, witha space separating the two. If sufficient space is not available, the iden-tifier may be added as a separate entry after the frequency has been en-tered.The general format is: T(freq)spaceR(freq)(emission mode)space(iden-tifier). The identifier will be displayed in cyan to the right of the preset fre-quency.

e. To load a preset frequency into the active window, use either the left orright LSK associated with the desired frequency. The frequency will beloaded into the CDU and RTU active frequency window.

More detailed information may be found in the following documents.

Collins HF-9000 High-Frequency Communications System Operator’s Manual,P/N 523-0774344-008217

Collins RTU-4200/4210/4220 Radio Tuning Unit Pilot’s Guide, 2nd edition, datedMay 1, 1996, or later edition

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AVTECH CSD-714 SELCAL DECODER (IF INSTALLED)

The SELCAL (Selective Call) Decoder system consists of a CSD-714 decodermounted in the aft avionics compartment and a control panel mounted on the ped-estal in the cockpit. The decoder is powered by the No. 3 Avionics Bus through a1-amp circuit breaker, placarded SELCAL, located on the right circuit breaker pan-el. The control panel, placarded SELCAL, consists of a PUSH OFF switch/annun-ciator, and a red test button placarded SELF TEST. Backlighting of the controlpanel is adjusted using the OVHD PED & SUBPANEL rheostat located on theoverhead panel. The white HF SELCAL annunciator on the switch is illuminatedwhen a SELCAL code is received. The annunciator may be tested using the an-nunciator PRESS TO TEST button located on the glareshield. The illumination ofthe annunciator is adjusted using the ANNUN PUSH-BRT rheostat located on theoverhead panel. The decoder is capable of receiving 16 tone codes.

When a SELCAL code is received by the decoder, the HF SELCAL annunciatoron the control panel will begin to blink and a single tone will be heard over thespeakers (speaker switch on or off) and the headsets. The annunciator can be ex-tinguished by pressing the switch/annunciator.

The SELCAL decoder may be manually tested by pressing the SELF TEST buttonon the control panel. When the button is pressed, a single tone will be heard overthe speakers and headsets and the white HF SELCAL annunciator will blink forapproximately 5 seconds, then extinguish.

The SELCAL decoder can function in normal operation only when the HF radio isset to the AM Emission Mode.

IFIS-5000 INTEGRATED FLIGHT INFORMATION SYSTEMThe Rockwell Collins IFIS-5000 Integrated Flight Information System (IFIS) sup-plements the EFIS-3000 Electronic Flight Instrument System portion of the ProLine 21 system in the Model B200GT/B200CGT by providing advanced displayfunctions to aid the pilot in charting and weather information. Features which maybe used with the IFIS-5000 are as follows:

• Electronic Charts - Provides Jeppesen approach, departure and airportscharts. This function requires a subscription from Jeppesen.

• Graphical Weather (optional) - Provides current and forecasted weatherfrom either XM Radio or Universal Weather. This function requires a sub-scription service from XM Radio or Universal Weather.

• Enhanced Map Overlays - Provides additional overlays on the FMS planningmap (PLAN) and present position map (PPOS). These overlays includeGeographic features, Geo-Political boundaries, Airways, and Airspace.

The following line replaceable units comprise the IFIS-5000 System:• FSU-5010 File Server Unit provides the processing and storage for all IFIS

functions.

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• AFD-3010E Adaptive Flight Display MFD (part of EFIS-3000) displays IFISfunctions.

• Weather Data Interface (optional): • XM Weather - XM radio receiver and top-mounted satellite antenna to

receive and process satellite-based weather from XM Radio or,• Universal Weather - Collins Communications Management Unit (CMU),

Collins VHF com and bottom-mounted VHF blade antenna to receiveand process weather data from Universal Weather.

• CCP-3000 Cursor Control Panel located in the pedestal controls the IFISfunctions via MFD on-screen menus. Dedicated controls are provided forchart selection, quick MFD format access keys and MFD menu controls. Ajoystick is used for panning across the charts.

More detailed information may be found in the following documents:

Collins IFIS-5000 Integrated Flight Information System Operator’s Guide,Document 523-0806347-005117, dated 4 April 2007, or later revision.

Collins Corporate Datalink System CMU-4000 / RIU-40X0 Operator’s Guide,Document 523-07090499, dated 14 July 2003, or later revision.

RADIO TUNING SYSTEMThe Radio Tuning System provides the control, displays, and sensors for VHFvoice communication, HF voice communication (if installed), VOR/ILS/DME, ADF,transponder tuning, and TCAS II (if installed). The system consists of the RadioTuning Unit (RTU) located in the center instrument panel, and a radio tuning unitcontained in the CDU which is located in the pedestal. The RTU is considered theprimary method of tuning and the CDU as the secondary method of tuning. Thetuning capabilities of the CDU are accessed using the TUNE page. If Dual CDUsare installed, only the left CDU (CDU 1) has radio tuning capabilities.

An RTU/CDU TUNE reversionary switch is located on the reversionary panel.With this switch in the NORM position radios may be tuned using either the RTUor the CDU. Should the RTU become inoperable, the RTU/CDU TUNE switchmust be placed to the CDU position, and the radios tuned using the CDU. Like-wise, if the CDU should become inoperable, the RTU/CDU TUNE switch must beplaced in the RTU position and the radios tuned using the RTU.

If radio tuning capability is lost from both the RTU and the CDU, the EMER TUNEswitch, located on the reversionary panel, may be pushed to tune the No. 1 Comto the emergency frequency of 121.5 MHz. Activation of the switch is denoted bythe illumination of the annunciator, 121.5, located on the switch.

The radio tuning unit in the CDU has the capability of automatically tuning the VHFNAV receivers in order to improve the calculation of airplane position by the FMS.This auto tune function is selected on the NAV CONTROL page of the CDU. Theauto tune function is automatically cancelled if any of the following occur:

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• DME HOLD is selected.• A NAV receiver is manually tuned using either the RTU or the CDU.• The FMS is deselected as a NAV source.• A NAV receiver fails.

If a malfunction occurs in the auto tune function, it may be manually disabled usingthe RMT TUNE switch, which is located on the reversionary panel. Moving thisswitch from the NORMAL position to the DISABLE position will disable the autotuning function of the CDU.

The RTU is protected by a 2-amp circuit breaker, placarded RTU, located on theright circuit breaker panel. The CDU (and its radio tuning unit) is protected by a 3-amp circuit breaker, placarded CDU NO. 1.

For further information, refer to the Collins Pro Line 21 Avionics and Flight Man-agement System Operator’s Guides.

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TERRAIN AWARENESS AND WARNING SYSTEM PLUS(TAWS+) (IF INSTALLED)Refer to Operator’s Guide Collins Pro Line 21 Avionics System with optional IFISfor the Beechcraft King Air, P/N 523-0808535, dated June 1, 2007, or later revi-sion.

Refer to Aviation Communication and Surveillance Systems (ACSS) TAWS+ Pi-lot’s Pilot’s Guide, Document No. 8006772-001, dated March 2007, or later revi-sion.

The information in the ACSS Pilot’s Guide is applicable except in the four areasnoted below:

1. The radio altitude of 245 feet depicted in the “Too Low Flaps” envelopeshown in Figure 1-16 (Too Low Terrain/Gear/Flaps Envelope for Mode 4)on page 1-29 of the ACSS Pilot’s Guide should read 195 feet.

2. The terrain display patterns shown in Figure 2-2 (Terrain Display Patternsfor Various Terrain Elevations) on page 2-5 of the ACSS Pilot’s Guideshould be as follows:

BB07C084547AA.AI

LOW DENSITYYELLOW

HIGH DENSITYRED

HIGH DENSITYYELLOW

HIGH DENSITYGREEN

LOW DENSITYGREEN

BLACK

+ 2000 FT

+ 1000 FTFPA

REF. ALT.

- MTCD FT

- 1000 FT

- 2000 FT

30 SEC.

TERRAIN COLORATION

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3. The following table replaces Table 2-1 (Terrain Display Patterns for VariousTerrain Elevations) on page 2-7 of the ACSS Pilot’s Guide:

Terrain Elevation(typical values)

Terrain Color Dot Density

2000 ft or more above the airplane’s reference altitude

Red High

1000 ft to 2000 ft above theairplane’s reference altitude

Yellow High

Maximum of 500 ft (MTCD) below the airplane’s referencealtitude to 1000 ft above the airplane’s reference altitude

Yellow Low

Maximum of 500 ft (MTCD) to1000 ft below the airplane’sreference altitude

Green High

1000 ft to 2000 ft below theairplane’s reference altitude

Green Low

2000 ft or greater below theairplane’s reference altitude

Black N/A

Invalid terrain cell Magenta SolidTerrain Caution Yellow SolidTerrain Warning Red SolidAvoid Terrain Red/Black Black with red

cross-hatch

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4. The following table replaces Table 2-2 (Obstacle Display Patterns for Vari-ous Obstacle Elevations) on page 2-9 of the ACSS Pilot’s Guide:

The following information supplements the two documents listed above:

The TAWS+ is powered from the No. 4 Dual-Fed Bus and is protected by a 3-ampcircuit breaker, placarded TAWS, located on the copilot’s circuit breaker panel.

The system operates by accepting a variety of airplane parameters as inputs, ap-plying alerting algorithms, and providing the flight crew with data (aural alert mes-sages, visual annunciations and displays) in the event that the boundaries of anyalerting envelope are exceeded. The system performs the following functions:

1. Processes inputs from the airplane systems.2. Correlates the current airplane position to terrain and airport databases.3. Projects the terrain hazards along the projected flight path ahead of the air-

plane.4. Generates an image of potential terrain hazards for display to the flight

crew.5. Generates applicable alerts and warnings, both aurally and visually, to the

flight crew.

Obstacle Elevation(typical values)

Obstacle Color Dot Density

2000 ft or more above the airplane’s reference altitude

Red High

1000 ft to 2000 ft above theairplane’s reference altitude

Yellow High

Maximum of 500 ft (MTCD) below the airplane’s reference altitude to 1000 ft above the airplane’s reference altitude

Yellow Low

Maximum of 500 ft (MTCD) to 1000 ft below the airplane’s reference altitude

Green High

1000 ft to 2000 ft below theairplane’s reference altitude

Green Low

2000 ft or greater below theairplane’s reference altitude

Black N/A

Obstacle Caution Yellow SolidObstacle Warning Red SolidAvoid Obstacle Red/Black Black with red

cross-hatch

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The TAWS+ is designed to be fully compatible with airplane operations. Unwantedalerts will be very rare if the pilot maintains a situational awareness with respectto terrain. There is normally no requirement for the pilot to input to the system, ex-cept for preflight self-testing.

Basic Ground Proximity Warnings

The following operating modes generate the following cautions and warnings aspart of the basic (non-enhanced) ground proximity warning system. Several of themodes present an initial caution to alert the crew of a developing problem that re-quires their attention. If the situation worsens, a warning is generated to alert thecrew to provide increased correction.

The following messages and aurals are provided for TAWS+ Modes 1 - 6. SeeSection 3, EMERGENCY PROCEDURES, and Section 3A, ABNORMAL PROCE-DURES, for the appropriate procedure associated with each caution and warningalert.

Mode Function PFD Caution Message(Yellow)

Caution Aural

PFD Warning Message

(Red)

WarningAural

1 Excessive Descent

Rate

GND PROX “SINK RATE” PULL UP “PULL UP”

2 Excessive Closure on

Terrain

GND PROX “TERRAIN, TERRAIN”

PULL UP “PULL UP”

3 Altitude Loss After Takeoff

GND PROX “DON’T SINK, DON’T SINK”

N/A N/A

4a Unsafe Terrain Clearance

GND PROX “TOO LOW GEAR”

N/A N/A

4b UnsafeTerrain

Clearance

GND PROX “TOO LOW FLAPS”

N/A N/A

5 Excessive Glideslope Deviation

GND PROX “GLIDESLOPE” N/A N/A

6a Excessive Bank Angle

N/A “BANK ANGLE”

N/A N/A

6b Altitude Callouts

N/A “500, 200, 100, 50, 40, 30, 20,

10”

N/A N/A

6f Selected MDA N/A “MINIMUMS, MINIMUMS”

N/A N/A

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GPWS Functions

The following equipment is required to be operational for the proper functioning ofMode 1 through 6 of the ACSS TAWS+ system:

1. TAWS+ Warning System Computer.2. Radio Altimeter.3. Vertical Speed from the Air Data Computer.4. Airspeed from the Air Data Computer.5. Glideslope and Localizer Deviation.6. Landing Gear Position.7. Flap Position.8. Roll attitude from Pilot’s Attitude System (for “BANK ANGLE” voice mes-

sage).9. Decision Height System (for “MINIMUMS” voice message).

The following Mode 6 advisory callouts are enabled: “FIVE HUNDRED”, “TWOHUNDRED”, “ONE HUNDRED”, “FIFTY”, “FORTY”, “THIRTY”, “TWENTY”,“TEN”, and “MINIMUMS”.

The following three pushbutton switch/annunciators for TAWS+ are located direct-ly in front of the pilot:

Switch/AnnunciatorColor

ACTIVEState

Function

FLAPOVRD AMBER

Pressing the switch disables the “TOO LOWFLAPS” portion of the GPWS Mode 4b alertboundaries and also selects the Mode 2benvelope. The “ACTIVE” annunciatorilluminates when the switch is pressed.

G/SINHIB

AMBER

Illuminates to indicate the GPWS Mode 5glideslope alert has been inhibited. Pressingthe switch on the ground initiates self-test.The “ACTIVE” annunciator illuminates ambermomentarily when pressed and thenextinguishes when released. However, theglideslope alerting will remain inhibitedalthough the “ACTIVE” legend will beextinguished. The inhibit function is enabledbelow 2000 ft AGL and disabled at 30 ft AGLor after ascending above 2000 ft AGL.

TERRINHIB AMBER

Pressing the switch deselects all CPAfunctions of the TAWS+ system. The“ACTIVE” annunciator illuminates when theswitch is pressed.

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If terrain data penetrates the caution or warning envelopes, then the correspond-ing aural and visual alerts are generated.

Terrain display can be selected manually at anytime. Areas of terrain sufficientlyclose to the airplane that do not penetrate the terrain caution or warning envelopesare depicted by areas of red, yellow or green dot patterns. The color and dot den-sity vary based on terrain elevation relative to the airplane. Magenta coloring isused to indicate areas where terrain information is unavailable.

NOTEIf the terrain is within 15 NM of the nearest airport, terrain with an el-evation of less than the nearest runway elevation + 400 feet is dis-played as black.

The following annunciators, aural alerts, and aural warnings are provided for theCPA function of TAWS+. See Section 3, EMERGENCY PROCEDURES, andSection 3A, ABNORMAL PROCEDURES, for the appropriate procedure associ-ated with each alert and warning.

Function/Mode PFD Caution Message (Yellow)

Caution Aural

PFD Warning Message

(Red)

WarningAural

Terrain Alerting and Display (TAD)

GND PROX “CAUTION TERRAIN, CAUTION TERRAIN”

PULL UP “TERRAIN, TERRAIN, PULL UP, PULL UP”

GND PROX “CAUTIONOBSTACLE

,CAUTION

OBSTACLE”

PULL UP “OBSTACLE,OBSTACLE,

PULL UP,PULL UP”

GND PROX “CAUTION TERRAIN”

or“CAUTIONOBSTACLE

”

PULL UP “AVOIDOBSTACLE”

or“AVOID

TERRAIN”

Premature Descent Alerting (PDA)

GND PROX “TOO LOW, TERRAIN”

N/A N/A

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TAWS+ Functions

The following equipment must be operational for the proper functioning of the pre-dictive features of the TAWS+:

1. TAWS+ Warning Computer.2. Heading from the No. 1 Compass System (used for Terrain display only).3. GPS position. 4. Terrain and Airport Database.

The following predictive features are available:1. Terrain Display.2. Collision Prediction and Alerting (CPA).

Obstacles

Obstacles in the database are defined as those man-made obstacles, such astowers, buildings and antennas. An obstacle has a height of at least 250 ft. If ac-tivated, the obstacle database has the same alert displays as the terrain database(“CAUTION OBSTACLE”, “OBSTACLE AHEAD, PULL UP”, “AVOID OBSTA-CLE”) and can be inhibited by selection of the TERR INHIB switch/annunciator.

NOTEWith Terrain selected to be displayed on any cockpit display, usuallywhen display ranges of 200 NM or greater are selected, it may bepossible for a short line of red pixels to appear in front of the wing ofthe displayed aircraft symbol. These red pixels are not associatedwith a TERRAIN or OBSTACLE warning and have a different ap-pearance to the solid red square associated with a TERRAIN or OB-STACLE warning. This display anomaly may disappear andreappear at random intervals. If this occurs, select a lower range dis-play and ensure the red area disappears from the display. If any redis displayed and a TERRAIN or OBSTACLE warning is issued, suit-able avoiding action must be taken.

TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM(TCAS I) (IF INSTALLED)The Goodrich or L3 Communications Skywatch HP Traffic Alert and CollisionAvoidance System, Model SKY899, is to be used for aiding visual acquisition ofconflicting traffic. The system includes a transmitter-receiver computer (TRC) anda directional antenna mounted on the top of the fuselage. The installation receivespressure altitude from the pilot’s or copilot’s encoding altimeter through the No. 1or No. 2 transponder. The system also receives inputs from the right weight-on-wheels switch, the right landing gear downlock switch, radio altimeter, and head-ing input from the No. 1 compass. The system is powered from the No. 3 AvionicsBus, and is protected by a 5-amp circuit breaker, placarded TCAS.

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The SKY899 is an active system that operates as an aircraft-to-aircraft interroga-tion device. The system interrogates up to 35 different aircraft transponders in a35 nm radius in the same way ground-based radar interrogates aircraft transpon-ders. When the SKY899 receives replies to its interrogations, it computes the re-sponding aircraft’s range, relative bearing, relative altitude, and closure rate. TheSKY899 then predicts collision threats and plots the eight most threatening aircraftlocations.

The SKY899 is controlled with the following:• Operating Mode Button: This switch/light is placarded ON/STBY. ON is illuminat-

ed when the system is in the operating mode. The switch/light will be blank whenthe system is in the standby mode. On the ground, this switch can be used tochange the operating mode between ON and STBY. In flight, this switch is inac-tive and the system is continuously in the ON mode.

• Display Range Knob: The display range is controlled through the range knob lo-cated on the DCP.

• Vertical Display Mode/Test Button: This push-button is placarded TEST/ALT. Onthe ground, pressing this button will begin an internal self-test. This test shouldbe conducted in accordance with Section 4, NORMAL PROCEDURES, beforethe first flight of the day. In flight, this button acts as a Vertical Display Mode con-trol, allowing the pilot to toggle the display between ABOVE, BELOW, ABOVE/BELOW and Normal.

The SKY899 will display the following features:• Solid Yellow Circle: This is the Traffic Advisory (TA) symbol that the SKY899

generates when it predicts that an intruder aircraft may pose a collision threat.This is accompanied by the aural alert “TRAFFIC TRAFFIC”.

• Solid Cyan Diamond: This is the Proximate Traffic symbol that is generatedwhen intruder traffic is detected within 6 nm and ±1200 feet, but does not posea threat.

• Open Cyan Diamond: This is the symbol for Other Traffic and is generated torepresent an intruder aircraft that has been detected but has not generated atraffic alert.

• Solid Yellow Semicircle: This is the Traffic Advisory (TA) symbol that theSKY899 generates when it predicts than an intruder aircraft may pose a collisionthreat but is out of the current display range.

• Vertical Trend Arrow: A vertical trend arrow will be shown to the right of the trafficsymbol to indicate that traffic is ascending or descending at a rate greater than500 fpm. An arrow is not shown for non-altitude reporting aircraft.

• Data Tag (Example +04): A two-digit number representing the relative altitude,in hundreds of feet, of the intruder aircraft. A positive data tag is displayed abovethe traffic symbol and a negative data tag is displayed below the traffic symbol.If the intruder aircraft is at the same altitude as your airplane, 00 is displayedabove the traffic symbol.

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• Four altitude display modes are available:• Look-up Mode (ABOVE): Displays traffic detected within +9,000 feet to -

2,700 feet of your airplane.• Normal Mode: Displays traffic detected within ±2,700 feet of your airplane.• Look-down Mode (BELOW): Displays traffic detected within +2,700 feet to -

9,000 feet of your airplane.• Unrestricted Mode (ABOVE/BELOW): Displays traffic detected within

±9,000 feet of your airplane.

When the TCAS self-test is conducted as described in the Normal Procedures, thefollowing test pattern will be displayed on the MFD.• Traffic Advisory (solid yellow circle) will appear at 9 o’clock, range 2 miles, 200

feet below and climbing. • Proximate Traffic (solid cyan diamond) will appear at 1 o’clock, range 3.6 miles,

1000 feet below and descending.• Other Traffic (open cyan diamond) will appear at 11 o’clock, range 3.6 miles, fly-

ing level 1000 feet above.

The SKY899 has the following automatic features:• Using the right weight-on-wheels switch, the SKY899 will automatically switch

from STBY Mode to Operate Mode in the 6 nm range and ABOVE Mode will beselected approximately 8 to 10 seconds after takeoff.

• Using the right weight-on-wheels switch, the SKY899 will automatically switchfrom Operate Mode to STBY Mode approximately 24 seconds after landing.

• Using the radio altimeter input, the SKY899 will inhibit the aural alerts below 400feet AGL to minimize pilot distraction.

Refer to the Pilot’s Guide for the Skywatch HP Traffic Alert/Advisory System, Mod-el SKY899, Goodrich or L3 Communications PN 009-11901-001, Rev A, datedAugust 29, 2001, or later revision.

TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM(TCAS II) (IF INSTALLED)The Collins TCAS-4000 is a TCAS II system designed to protect a volume of air-space around the TCAS II-equipped airplane by warning the pilot of the threat ofother transponder-equipped airplanes penetrating that airspace. The system in-terrogates Mode C and Mode S transponders in nearby airplanes and analyzestheir replies to identify potential and predicted collision threats. The system advis-es the pilot when to climb, descend, or maintain altitude to avoid passing too closeto, or colliding with, the threat airplane. When an intruder airplane is equipped withTCAS II, the system coordinates avoidance maneuvers with this airplane usingthe data link capability of the Mode S transponders.

If traffic gets within 25 to 45 seconds (depending upon altitude) of the projectedClosest Point of Approach (CPA), it is considered an intruder and a Traffic Advi-

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sory (TA) is issued. This TA calls attention to what may develop into a collisionthreat using visual and aural alerts. The visual alert consists of a solid yellow circledepicting the intruder on the traffic map and a yellow flashing TRAFFIC messageon the PFDs. The aural alert consists of the voice message, “TRAFFIC, TRAF-FIC”. These alerts promote mental and physical preparation for a possible maneu-ver that may follow, and assists the pilot in achieving visual acquisition of theintruding aircraft.

If the intruder gets within 20 to 35 seconds (depending upon altitude) of the CPA,it is considered a threat, and a Resolution Advisory (RA) is issued. This RA pro-vides a recommended vertical maneuver using modified instantaneous verticalspeed indicators (IVSIs) and voice messages to provide adequate vertical sepa-ration from the threat aircraft (a Corrective RA) or prevents initiation of a maneuverthat would place the TCAS II aircraft in jeopardy (a Preventive RA). In addition tothe voice messages, e.g. “CLIMB, CLIMB”, the threat aircraft is depicted as a solidred square on the Traffic Map, and a red flashing TRAFFIC message is displayedon the PFDs.

The TCAS II system consists of a TCAS II receiver-transmitter, two TCAS direc-tional antennas, and two diversity Mode S transponders with antennas. The sys-tem receives altitude and vertical speed information from the pilot’s Air DataComputer (ADC1). If that system fails, information is provided by the copilot’s AirData Computer (ADC2). Radio altitude information is provided from the radio al-timeter, and heading information from the pilot’s AHS. The system also receivesinputs from the right weight-on-wheels switch and right landing gear downlockswitch. The TCAS II system generates vertical guidance commands that are dis-played on the pilot’s and copilot’s IVSIs in the form of vertical red and green bands.Vertical speeds located next to the red band are to be avoided. The vertical speedassociated with the green band (either descending or climbing) is the verticalspeed the pilot should attain. Intruder targets are displayed on the MFD on theTCAS Only Map, or may be overlaid on the Present Position Map. Aural alerts aresounded over the speakers, whether or not they are selected on, and also overthe headsets. Controls for the TCAS II system are integrated into the RTU and theCDU. Either unit may be used to control the TCAS system. The TCAS II systemis powered by the No. 3 Avionics Bus and is protected by a 5-amp TCAS circuitbreaker located on the right circuit breaker panel. Power is applied to the systemwhen the Avionics Master switch is turned on.

MFD DISPLAYS AND CONTROLS

The TCAS Traffic Only Map may be selected by pressing the TFC line select keyfor more than 1 second. The TCAS Traffic Only Map will be displayed in the 10 nmrange. The range of the display may be adjusted from 5 nm to 50 nm using theRANGE knob on the Display Control Panel. The TFC key may also be used to se-lect the TCAS Traffic Display on or off.

Once the Traffic Only Map has been selected using the TFC key, the FORMATkey may be used to select the Plan Map, the Present Position Map, or the TCASOnly Map.

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The following messages appear along the right side of the display when appropri-ate. They are listed as they would appear from top to bottom:

ABS INOP(white)

If the Absolute Altitude Mode is selected and the airplane isbelow 18,000 feet P.A., this display is presented.

ALT XXX(cyan)

If the Absolute Altitude Mode is selected and the airplane isabove 18,000 feet P.A., this display will show airplane pressurealtitude in thousands and hundreds of feet.Example: 23,000 feet = 230.

ABOVE/BELOW(white)

These messages indicate the operating altitude volume of theTCAS system. These messages will be shown as ABOVE,ABOVE BELOW, BELOW, or will be blank. The operatingvolume of each display is as follows:ABOVE = -2700 ft to +9900 ftBELOW = +2700 ft to -9900 ftABOVE/BELOW = -9900 ft to +9900 ftBlank = -2700 ft to +2700 ft

OFF(cyan)

This message indicates that the OTHER TRAFFIC symbol hasbeen selected OFF.

TFC(cyan or white)

This legend indicates that the TCAS II system has beenselected for display (cyan), or has been selected OFF (white).

TCAS TEST(cyan)

This message indicates that the TCAS II is in the Test Mode.(Color is white if TCAS display has not been selected.)

TCAS OFF(cyan)

This message indicates that the Standby Mode of the TCASsystem has been selected, the standby mode of thetransponder has been selected, or that the Mode C has beenselected OFF. (Color is white if TCAS display has not beenselected.)

TA ONLY(cyan)

This message indicates that the TA Only Mode has beenselected. It will always be displayed on the ground. Themessage will change color from cyan to yellow and flash whena TA is issued by the TCAS. (Color is white if TCAS display hasnot been selected.)

TCAS FAIL(yellow)

This message indicates a TCAS fault has been detected.

TA or RA with no bearing data

Two lines are provided for the first two detected TAs or RAswithout valid bearing data. Each line of data will be yellow forTA data or red for RA data. The format of each line will includethe range of the intruder followed by the relative or absolutealtitude, if available, and a rate-of-climb or descent directionarrow if applicable.

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When the TCAS self-test is conducted as described in Section 4, NORMAL PRO-CEDURES, the following test pattern will be displayed on the MFD:• Traffic Advisory (solid yellow circle) will appear at 9 o’clock, range 2 miles, 200

feet below, and climbing. • Proximate Traffic (solid cyan diamond) will appear at 1 o’clock, range 3.6 miles,

1000 feet below, and descending.• Other Traffic (open cyan diamond) will appear at 11 o’clock, range 3.6 miles,

1000 feet above, and in level flight.• Resolution Advisory Traffic (solid red square) will appear at 3 o’clock, range 2

miles, 200 feet above, and in level flight.

PFD DISPLAYSThe PFD does not display traffic unless in the reversionary mode. The followingTCAS messages and displays are provided.The following messages will be displayed, one at a time, just below the lower rightcorner of the EADIs:

The following messages will be displayed on the right side of the PFD opposite thethird Line Select Key. They are identical to those shown on the MFD.• TCAS TEST (white)• TCAS OFF (white)• TA ONLY (white)

During a Resolution Advisory, red or red and green bands will be displayed on theIVSI. There are two types of RAs; corrective and preventive.

If a corrective RA is issued, red and green bands will be displayed. The greenband indicates the rate-of-climb or descent required for the pilot to obtain in re-sponse to the RA. The red bands indicate the rate-of-climb and descent the pilotis to avoid during the response to the RA.

If a preventive RA is issued, normally only a single red band will be displayed in-dicating the vertical speeds to be avoided. If intruders exist above and below theairplane, it is possible to have a green band covering the lower rates-of-climb and/or descent followed by two red bands indicating the higher rate-of-climb and de-scent to avoid.

During the TCAS self-test, the IVSI will display the following test pattern:• A red band will extend from 0 fpm to the bottom of the display.• A green band will extend from 0 fpm to +300 fpm.• A red band will extend from +2000 fpm to the top of the display.

TRAFFIC(yellow or red)

This message will be yellow for a TA and red for an RA. It willblink approximately 6 times and then become steady.

TCAS FAIL(yellow)

This message is identical to the one shown on the MFD.

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SYSTEM CHARACTERISTICS

Only the TA Only Mode is available during ground operations. The RA Mode isavailable after takeoff above approximately 1150 feet.

The Traffic Display may be overlaid on the radar or EGPWS display on the MFD.

EGPWS and radar displays are not available on the TCAS Traffic Only map.

EGPWS voice alerts have priority over TCAS II voice messages. During such oc-casions, the TCAS II will automatically switch to the TA Only Mode with no TCASvoice messages.

The TCAS II surveillance may not function at distances less than 900 feet.

CLIMB and INCREASE CLIMB RAs are inhibited with flaps extended beyond theApproach position.

CLIMB and INCREASE CLIMB RAs are inhibited above 32,000 feet P.A.

When below approximately 1000 feet, the TCAS II will automatically revert to theTA Only Mode.

All RA and TA voice messages are inhibited below 600 feet AGL while climbingand below 400 feet AGL while descending.

DESCEND RAs are inhibited below 1200 feet AGL while climbing and below 1000feet AGL while descending.

INCREASE DESCENT RAs are inhibited below 1450 feet AGL.

Failure of the radio altimeter will cause the TCAS II system to be inoperable.

VOICE MESSAGES

The following voice message accompanies a TCAS II Traffic Advisory (TA).

The following voice messages accompany TCAS II Resolution Advisory Traffic(RAs).

VOICE MESSAGE PILOT RESPONSE“TRAFFIC, TRAFFIC” Gain visual contact with traffic. Check the TCAS II display

for range and bearing of the traffic if necessary. Assessthe threat and prepare to execute the evasive maneuverif a Resolution Advisory is subsequently issued.

VOICE MESSAGE PILOT RESPONSE

“CLIMB, CLIMB, CLIMB” (corrective)

Change vertical speed to 1500 fpm climbing, or asindicated by the green band on the IVSI.

“CLIMB, CROSSING CLIMB, CLIMB, CROSSING CLIMB”(corrective)

Same as previous except that this message indicatesthat flight paths will cross at some altitude.

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SYSTEMS DESCRIPTION

AVIONICSTRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM(TCAS II) (IF INSTALLED)SYSTEM CHARACTERISTICS

For airplanes with Rockwell Collins, Inc. TCAS II System (Version 7.1) installed,all RAs are inhibited below 1100 feet AGL while climbing and below 900 feet AGLwhile descending and will revert to TA only mode.

VOICE MESSAGES

For airplanes with Rockwell Collins, Inc. TCAS II System (Version 7.1) installed,the following resolution advisory (RA) will be annunciated aurally and visually asappropriate.


“CLIMB, CLIMB” Climb at the rate depicted by the green (fly to) arcor line on the IVSI or other suitable indicator.

“DESCEND, DESCEND” Descend at the rate depicted by the green (fly to)arc or line on the IVSI or other suitable indicator.

“MONITOR VERTICAL SPEED”

Ensure that vertical speed is out of the illuminatedIVSI red arc or line, or other suitable indication.

“LEVEL OFF, LEVEL OFF” Reduce vertical speed to zero feet per minute. Agreen arc or line will be illuminated beginning atzero feet per minute.

“CLEAR OF CONFLICT” Expeditiously return to the applicable clearance,unless otherwise directed by ATC.

“CLIMB, CROSSING CLIMB, CLIMB, CROSSING CLIMB”

Climb at the rate depicted by the green (fly to) arcor line on the IVSI or other suitable indicator.

“DESCEND, CROSSING DESCEND, DESCEND, CROSSING DESCEND”

Descend at the rate depicted by the green (fly to)arc or line on the IVSI or other suitable indicator.

“MAINTAIN VERTICAL SPEED, MAINTAIN”

Maintain the existing climb or descent rate asdepicted by the green (fly to) arc or line on the IVSIor other suitable indicator.

“MAINTAIN VERTICAL SPEED, CROSSING MAINTAIN”

Maintain the existing climb or descent rate asdepicted by the green (fly to) arc or line on the IVSIor other suitable indicator.

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For airplanes with Rockwell Collins, Inc. TCAS II System (Version 7.1) installed,the following voice messages annunciate enhanced TCAS II maneuvers whenthe initial RA does not provide sufficient vertical separation. The tone andinflection connotes increased urgency.


“INCREASE CLIMB, INCREASE CLIMB”

Increase climb at the rate depicted by the green (flyto) arc or line on the IVSI or other suitable indicator.

“INCREASE DESCENT, INCREASE DESCENT”

Increase descent at the rate depicted by the green(fly to) arc or line on the IVSI or other suitableindicator.

“CLIMB — CLIMB NOW, CLIMB — CLIMB NOW”

Immediately climb at the rate depicted by the green(fly to) arc or line on the IVSI or other suitableindicator.

“DESCEND — DESCEND NOW, DESCEND — DESCEND NOW”

Immediately descend at the rate depicted by thegreen (fly to) arc or line on the IVSI or other suitableindicator.

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“INCREASE CLIMB,INCREASE CLIMB”(corrective)

This follows a CLIMB voice message. The climbingvertical speed is typically increased to 2500 fpm asshown by the green band on the IVSI.

“ADJUST VERTICAL SPEED, ADJUST”(corrective)

Reduce climbing vertical speed to that shown on theIVSI.

“DESCEND, DESCEND NOW”(corrective)

This follows a CLIMB voice message. This messageindicates that a reversal of vertical speed from a climbto a descent is needed to provide adequateseparation.

“DESCEND, DESCEND, DESCEND”(corrective)

Change vertical speed to 1500 fpm descending, or asindicated by the green band on the IVSI.

“DESCEND, CROSSING DESCEND, DESCEND,CROSSING DESCEND”(corrective)

Same as previous except that this message indicatesthat flight paths will cross at some altitude.

“INCREASE DESCENT,INCREASE DESCENT”(corrective)

This follows a DESCENT voice message. Thedescending vertical speed is typically increased to2500 fpm as shown by the green band on the IVSI.

“ADJUST VERTICAL SPEED, ADJUST”(corrective)

Reduce descending vertical speed to that shown onthe IVSI.

“CLIMB, CLIMB NOW”(corrective)

This follows a DESCEND voice message. Thismessage indicates that a reversal of vertical speedfrom a descent to a climb is needed to provideadequate separation.

“CLEAR OF CONFLICT” Resume normal flight. Apparent conflict of airspacehas been resolved.

“MONITOR VERTICAL SPEED”(preventive)

Be alert for approaching traffic. Ensure that the IVSIneedle does not enter the area of the red band.

“MAINTAIN VERTICAL SPEED”(preventive)

Maintain present vertical speed and direction. Ensurethat the IVSI needle does not enter the area of the redband.

“MAINTAIN VERTICAL SPEED, CROSSING, MAINTAIN”(preventive)

A flight path crossing is predicted, but being monitoredby the TCAS II. Maintain present vertical speed anddirection. Ensure that the IVSI needle does not enterthe area of the red band.

“ADJUST VERTICAL SPEED, ADJUST”(preventive)

Indicates a weakening of the RA. This allows the pilotto start returning to an assigned altitude.


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WEATHER RADAR SYSTEMThe Weather Radar System is a fully-integrated system that detects precipitation,moisture-based turbulence and ground feature returns in front of the airplane,which can be displayed on the PFDs or the MFD and features the following oper-ating modes:

AUDIO SYSTEMThe audio system consists of dual DB-700 audio amplifiers, a crew interphonesystem, cabin PA, and a pilot and copilot audio control panel. The audio controlpanels are located on the instrument panel adjacent to each PFD and have almostidentical functions as described in the following. Refer to the INSTRUMENT PAN-EL illustration earlier in this section for a layout of the switches.

TRANSMIT (XMIT) SELECT SWITCH

The XMIT select switch has corresponding positions for selecting the Commtransceiver to be used for transmissions and includes 1 (for COMM 1), 2 (forCOMM 2), PA (for the cabin public address), HF (if installed), and TEL (if in-stalled). The switch also incorporates a master VOL control knob.

TRANSCEIVER AND RECEIVER AUDIO CONTROLS

Each Com transceiver and Nav receiver has an audio control incorporating a pull-on/ push-off switch and a rotary volume control. The audio controls are used toenable the audio of incoming radio receptions and include COMM 1, COMM 2,NAV 1, NAV 2, DME 1, DME 2 (if installed), TEL (for telephone) (if installed), ADF,HF (if installed), and MKR (for marker beacon). When the audio control is pulledout, the audio for the selected transceiver or receiver can be heard over the head-phones and speaker, if selected on. (When the audio control is pushed in, its vol-ume level remains fixed at the last setting.) Rotating the audio control after it hasbeen pulled out increases or decreases the volume. The audio controls include a

MODE DESCRIPTIONWX (Weather) Mode The basic weather detection mode. Depicts areas of

precipitation with four different colors, determined byreflectivity strength.

MAP Mode Used to depict the display colors on the PFD/MFD asaccentuated ground features.

WX+T (Weather Plus Turbulence) Mode

Contains WX mode features and includes detection ofprecipitation-related turbulence targets, including windshifts that contain precipitation.

TURB (Turbulence) Mode

Shows areas of precipitation-related turbulence within50 nautical miles.

TARGET (Target/Turbulence Alert) Mode

Used as an alert for precipitation and/or precipitation-related turbulence.

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white band that becomes visible when the control is pulled on to indicate the activeposition.

NOTEIf only one DME is installed, the DME 2 control monitors channel 2of DME 1. If two DMEs are installed, the DME 2 control monitorschannel 1 of DME 2.

MIC SWITCH

In the NORM position, the hand mic and boom mic are active. In the OXY position,the mic in the oxygen mask is active and the cockpit speakers are automaticallyselected ON.

PA (PUBLIC ADDRESS) AUDIO CONTROL

This control functions in a manner nearly identical to the Comm and Nav audiocontrols. The only difference is that the control does not incorporate a pull-on/pushoff feature, but instead is always on. When transmissions are made using the PAfunction of the XMIT select switch, rotating this control increases or decreases thevolume of the audio from the cabin speakers.

AUTO COMM SWITCH

The Auto Comm switch works in conjunction with the XMIT switch and the AudioControls of the radio transceivers. If the Auto Comm switch is selected, the AudioControl associated with the radio selected with the XMIT switch does not have tobe on (pulled out) in order to hear incoming audio. If the Auto Comm switch is notselected, the Audio Control associated with the radio selected with the XMITswitch must be pulled out in order to hear incoming audio.

SPEAKER (SPKR) SWITCH

When the SPKR switch is selected, the pilot and copilot overhead speakers areenabled. The audio volume from the speakers is controlled by the master VOLknob located on the XMIT switch and the Audio Controls.

NOTEThe volume controls adjust the audio volume from both the speak-ers and the headsets. It is not possible to adjust the volume of thespeakers independently of the volume of the headset.

INTERPHONE (INPH) AUDIO CONTROL

This control functions in a manner similar to the Comm and Nav audio controls.Pulling the control out enables the cockpit interphone system. Pushing the controlin disables the system. Rotating the control increases and decreases the volumeof the interphone system.

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NOTEThe same control on the copilot’s audio panel does not incorporatethe pull-on/push-off switch. Only the pilot’s control can turn the sys-tem on or off.

ALTERNATE AUDIO SWITCH

The Alternate Audio switch is normally left in the NORM position. If the audio am-plifier fails, the audio panel will not function normally and the Alternate Audioswitch must be placed to the ALTN position. When the ALTN mode is selected,the pilot can transmit and receive on COMM 1 using a boom mic, cockpit speakeror headphones. The volume of radio receptions is not controllable.

NOTETransmissions may be made on COMM 2 and the PA, but COMM 2receptions are not possible.

VOICE - BOTH - IDENT SWITCH

This switch controls the input filtering of the NAV 1, NAV 2, ADF 1, and ADF 2 (ifinstalled) controls. In the VOICE position, Morse codes are filtered out and onlyvoice transmissions are heard. In the BOTH position, both voice and Morse codetransmissions will be heard. In the IDENT position, only Morse code transmissionswill be heard.

February, 2012


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