Jet Cockpit Review

TheCockpitReview

B737-200

Canova Aviation PublicationsGilbert, Arizona

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The Cockpit Review®

B737-200

Canova Aviation Publications, Inc.

Revision 4, Copyright 1999, 2000Canova Aviation Publications, Inc.All rights reserved. Printed in the

United States of America.



B737-200

Canova Aviation Publications

Copyright 2000Canova Aviation Publications, Inc.All rights reserved. Printed in the

United States of America.

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B737-200

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B737-200

W.T. DanburyPhantom Jockey #16

McDonnell FH-1First 75 Navy Pilots To Fly Jets



B737-200

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B737-200

TheCockpitReview

B737-200

PREFACE

The material contained in this publication is based on information derivedfrom the Airplane Flight Manual, appropriate FAR’s, and other various non-copyrighted publications. It is important to remember that, in all instances,information from the manufacturer’s approved documentation for this air-plane or from applicable FARs shall always take precedence over the mate-rial in this review. This publication may not be current or compatible withyour airplane, therefore, Canova Aviation Publications does not guaranteethe accuracy of this document and assumes no responsibilities for its use.



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B737-200

Contents

Preface 7

Chapter I The Cockpit Review 13

Chapter II Aircraft Limitations 261

Chapter III Aircraft Schematics 277

Chapter IV Flight Training Profiles 300

Index 312



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Publisher’s NoteChapter One of the “Cockpit Review” has been designed and

formatted to follow a typical “oral” session that many airline and governmen-tal examiners use. We recommend that you prepare for the “oral” byreviewing this chapter while sitting in a CPT trainer. If a CPT trainer is notavailable, a cockpit wall panel would be helpful. Beginning with the over-head panel, verbally describe each panel using a clear voice. Present yourmaterial in an essay-type format. Try to verbally recall as much detailedinformation as possible about each of them. Then, review the “CockpitReview” and see if your presentation matches ours.

By orating and reviewing in this manner, your presentation anddelivery of the information will improve, giving the appearance of being wellprepared. Good luck with the oral!

Capt. Bob DanburyATP/CFII - B727, B737, DHC-8, Learjet, Citation



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How To Use The Cockpit Review

Flow Pattern Philosophy

• Most airlines have developed some type of “flow pattern” and“challenge/response checklist” philosophy. The concept of the flowpattern is to accomplish required tasks which may not be directly relatedto the normal checklist. After completing these flow pattern items, thechecklist is then used to cover critical items of flight.

• As previously mentioned, many oral examinations follow this flowpattern to evaluate a pilot’s total knowledge. The same flow pattern canbe also used in learning the aircraft systems. This is HOW and WHYthe Cockpit Review has been designed and presented in its currentformat. Always use the index to fine a particular subject matter.



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B737-200

Chapter I

The Cockpit Review



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DG/Slaved Switch

• The DG/Slaved switch is a two position switch. Selecting the DG posi-tion, supplies “non-corrected” directional gyro information to the com-pass system. Directional gyro data is normally corrected by the fluxvalve system. Selecting the switch to the SLAVED position, provides“corrected” directional gyro information to the compass system. Thefast synchronization process is also activated by the selection of theswitch to the DG position. A second method of initiating the fast syn-chronization process is the selection of the compass transfer switch.

Heading Control Knob

• Selecting the heading control knob to either the left or right index posi-tion, will cause the respective compass card to rotate to the desiredheading. Placing the heading control knob to the center index position,the heading control function will be inoperative. Selecting the controlknob to the first index mark, the compass card begins a slow rotation.The second index mark position, provides fast movement of the com-pass card.



B737-200

Synchronization Indicator

• The index alignment marks indicate the directional gyro function of thecompass system is being correctly slaved by the magnetic flux valvesystem. The synchronization indicator will be inoperative when the DG/Slaved switch is selected to the DG position.

Compass Transfer Switch

• The compass transfer switch provides the means to select the opposite(operable) compass system in the event of a compass failure or mal-function. The instrument transfer system may be inoperative for dis-patch provided the associated instruments are operating normally. Referto your MEL.

Compass System Review

• The synchronizing process of the flux valves and the compass systemcan also be seen on the synchronizing annunciator located on the RMI(as installed). If the RMI synchronizing annunciator needle is pointing at



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the DOT or CROSS indicators, it will indicate the compass system is outof synchronization.

• Fast synchronization of the compass can be accomplished by threedifferent ways:1. Reestablishing electrical power to the system.2. Using the compass transfer switch3. Selection of the DG/SLAVED switch to DG then back to the

SLAVED position.

Flight Control Switches

• The flight control switch is normally placed in the guarded ON position.This allows the respective hydraulic system pressure to pressurize theprimary flight control system.

• (Oral Topic) Selecting the Flight Control Switch to the STBY RUDposition, initiates the following actions:1. Activates the standby pump.2. Arms the standby rudder hydraulic low pressure light.3. Opens the standby rudder shutoff valve, thus pressurizing the

standby rudder power control unit. Corresponding hydraulic systempressure is isolated from the ailerons, elevators and rudder.



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• Selecting the respective flight control switch to the OFF position, iso-lates the corresponding hydraulic pressure from ailerons, elevators, rud-der, and the elevator feel computer.

• (Simulator Training Hint) A common aircraft system failure that manyinstructors and examiners give during simulator training is the failure ofthe standby hydraulic pump. During your preflight, check the operationof the standby hydraulic pump by placing the flight control switch to theSTBY RUD position. Observe the extinguishing of the low pressurelight and a momentary illumination (flicker) of the standby low pressurelight. Another operational check of the standby system is the applica-tion of foot pressure on the rudder pedals. Without hydraulic pressureavailable, the pedals are very hard to move.

Standby Hydraulic System

• (Oral Topic) The standby hydraulic system can be manually activatedby two means:1. Flight control switch.2. Alternate flaps switch.

• Selecting the flight control switch to the STBY RUD position, extin-guishes the standby hydraulic low pressure light. This indicates hydrau-lic pressure is now available to the standby rudder actuator. Standbyhydraulic power will now be available to the standby rudder, thrustreversers and the L.E. Devices. The equipment necessary to store,pressurize, deliver and filter standby hydraulic fluid is located in themain gear wheel well, on the keel beam and the aft wall.

Standby Hydraulic Low Quantity Amber Light

• The illumination of the standby hydraulic low quantity light indicates lowquantity condition within the standby hydraulic reservoir. The low quan-tity light is always armed. The light will illuminate when the fluid level inthe standby reservoir decreases below 50% full. The master cautionand the flight control annunciator will illuminate in conjunction with thelow quantity light.

• The standby hydraulic low quantity light may be inoperative prior todeparture providing the hydraulic quantity is checked prior to each flightsegment. Refer to your MEL.



B737-200

Flight Control Low Pressure Light

• The illumination of the flight control low pressure light indicates a lowhydraulic pressure condition. This condition is sensed by one of thethree low pressure sensors located within the corresponding aileron,elevator and rudder system. This low pressure warning function isdeactivated whenever the corresponding flight control switch is selectedto the STDBY RUD position and when the standby rudder shutoff valveactually opens. Upon the activation of any of the low pressure flightcontrol sensors, the FEEL DIFF light may also illuminate. The elevatorfeel computer and the elevator feel centering unit will sense a differen-tial pressure exceeding 25%, thus illuminating the FEEL DIFF light.

• The flight control low pressure lights may be inoperative for departureprovided other warning lights, hydraulic quantity and pressure indicatorsare operating normally. Refer to your MEL.

Alternate Flap Master Switch

• The alternate flap master switch is normally positioned in the guardedOFF position during flight operations. When selected to the ARM posi-tion, the following will occur:1. The trailing edge flap bypass valve will close.2. The standby pump will be activated.3. The standby hydraulic low pressure light will be armed.4. The alternate flaps position switch will be armed.

• (Oral Topic) The following conditions may require the use of the alter-nate flap master switch:1. Flaps fail to extend or retract in response to the movement of the

flap handle (no existing asymmetrical flap condition).

2. Hydraulic system A failure has occurred. It provides the means toactuate the flap bypass valve which prevents the possibility of ahydraulic lock of the flap drive unit. The electric motor operates theunit and will extend or retract the trailing edge flaps. The electricmotor can be described as a “high speed motor with low torquecapabilities”. The power source for this action is from the AC trans-fer bus. Caution must be exercise, since no asymmetric protectionis provided when operating the alternate flap system.



B737-200

3. Hydraulic system B failure has occurred. It provides the means foralternate flap extension. The standby hydraulic pump will power therudder power control unit.

4. The alternate master flap switch may also be used during a trailingedge flaps up landing operation. Example: a malfunction has oc-curred to the trailing edge flaps requiring a flaps up landing. Theleading edge devices may still operate normally when using thealternate flaps master switch. Follow the appropriate checklist pro-cedures as published for your aircraft.

• (Oral Topic) The alternate master flap switch should not be used duringasymmetrical trailing edge malfunctions. Do not attempt to move thetrailing edge flaps with this switch, as there is no asymmetric protection.

Alternate Flaps Position Switch

• The alternate flaps position switch when momentarily selected to theDOWN position, will cause the leading edge devices and trailing edgeflaps to extend. When selected to the UP position, the trailing edgeflaps retract. The leading edge devices will remain extended and can-not be retracted by the alternate flaps system.

• The alternate flap position switch may be used whenever the flaps fail toextend or retract in response to flap lever selection, and if no asymme-try condition exists. During alternate flap operations, plan a flaps 15landing. Approach planning is important, since flap extension and re-traction requires a considerable amount of time (approximately 2 min-utes) when using the alternate flap system. Observe the appropriatespeed limitation as listed in the checklist.

Flight Limitations: one cycle then 25 minutes OFF.

Asymmetrical Flap Protection System Review

• Asymmetrical flap protection has been provided for the wing trailingedge flaps. Anytime a difference is detected during the operation of thewing trailing edge flaps (left and right flaps), hydraulic power will auto-matically be removed from the flaps drive unit. This system uses theflap indicator associated with flaps asymmetry control circuit to provide



B737-200

the detection and protection capabilities. When the flap indicator point-ers are separated by a predetermined amount, the comparator switchcloses and applies power to the asymmetry shutoff relay. The ener-gized relay drives the trailing edge flap bypass valve to bypass andstops the hydraulic motor.

• Procedures call for the flaps to be moved to the detent nearest thesmallest actual flap position. Do not attempt to move the trailing edgeflaps using the alternate flap position switch, as there is no asymmetricprotection.

Flight Spoiler Switch

• Selecting the flight spoiler switch to the OFF position, closes the respec-tive flight spoilers shutoff valve. For all practical purposes, this switch isused for maintenance purposes only. These switches control the flightspoilers only, and has no effect on the operation of the ground spoilers.

• The flight spoilers are hydraulically actuated in reference to movementof the ailerons. The spoiler mixer receives input from the aileron systemand speed brake lever position. This allows the flight spoilers to aug-ment lateral control when simultaneously being used as speedbrakes.The flight spoiler system has a blow-down check valve in the pressureline that allows the spoilers to retract at high speeds.



B737-200

• Two flight spoilers are located on the upper surface of each wing.Hydraulic System A provides power to the inboard spoilers, and SystemB provides power to the outboard spoilers. There is no backup hydrau-lic power source for the outboard or inboard flight spoilers.

Yaw Damper Warning Light and Switch

• The illumination of the yaw damper warning light indicates that the yawdamper has been disengaged and/or the yaw damper is inoperative.Should the light remain illuminated after the yaw damper switch hasbeen selected OFF then ON, positioned the switch to OFF.

• The yaw damper system consists of a yaw damper coupler, rate gyro,and a yaw damper actuator in the rudder power control unit. The yawdamper has no turn coordination function. No rudder pedal movementis felt from yaw damper system operation. Total amount of ruddermovement is 3 degrees. Airspeed signals from the air data computerdecreases the amount of yaw damper rudder deflection at higher air-speeds.

• The yaw damper uses System B hydraulic pressure, loss of hydraulicpressure does not cause yaw damper disengagement or the illuminationof the amber Yaw Damper Warning Light. The yaw damper will disen-gage should Flight Control Switch B be selected.

• (SP77) With the yaw damper inoperative, the AFM restricts the use ofthe autopilot roll channel above 30,000’.

• (SP177) Aircraft equipped with SP177 autopilots, the 30,000’ limitationwill not apply. Yaw damper circuit breakers are located on the P6-2panels.

Feel Differential Pressure Light

• The illumination of the feel differential pressure light indicates excessivedifferential pressure in the elevator feel computer. This differential oc-curs when a 25% or more difference has been computed between theSystem A and System B hydraulic pressures. This warning indicationwill be armed only when the trailing edge flaps are up. No crew action isrequired during flight operations.



B737-200

• The elevator feel and centering unit provides artificial feel to the pilotand for centering of the elevator control system. The mach trim actua-tor is mounted on top of the feel and centering unit. Inputs to the feelcomputer are:1. Hydraulic A and B system pressures.2. Static pressure.3. Pitot pressure.4. Stabilizer position.

Mach Trim Fail Light

• The illumination of the mach trim fail light indicates mach trim dualchannel failure. Limit airspeed to .74 MACH. A single channel failurecauses the mach trim fail light to illuminate only when the master cau-tion annunciator recall function is activated. Light will extinguish whenmaster caution system has been selected to reset.

• Both mach trim channels may be inoperative for departure providing theAFM .74 MACH limitation is not exceeded. Refer to your MEL.

Mach Trim Test Button

• Selecting the mach trim test button, activates the mach yrim actuatorand moves the elevator up approximately 3o. The mach trim light willilluminate during the test. This test is normally performed by mainte-nance personnel.

Mach Trim System Review

• The mach trim system provides stability at mach speeds above .715 to.84 Mach. The functions of the system are automatic and has a directrelation to airspeed. The elevator is “programmed” to move with respectto the stabilizer. As speed increases, the center of lift moves rearward,thus balancing the nose downward. This is known as mach tuck. Themach trim system prevents the downward trend by adjusting the "pro-grammed movement" upward with the elevator. With the selection offlaps at a position other than zero, the mach trim light will be deacti-vated.



B737-200

Vertical Gyro Transfer Switch

• (As installed) The vertical gyro switch enables the selection of theopposite vertical gyro in the event of a malfunction and/or attitude fail-ure. The switch should be selected to the side of the instrument failure(gyro flag indication).

• (As installed) There are three vertical gyros installed to produce attitudereference information. They are labeled Vertical Gyro No.1, No. 2, andAuxiliary Vertical Gyro. Whenever vertical gyro No.1 or No.2 is unableto provide proper attitude reference, the vertical gyro switch may beused to select the auxiliary vertical gyro to provide information.

• The red INOP FLAG may be installed which indicates AUX VG failure.With the INOP flag displayed, do not select the switch. Caution shouldbe used when selecting the different modes with the autopilot con-nected, selection will disengage the autopilot. Vertical gyro No.1 orNo.2 may be inoperative for departure providing the auxiliary verticalgyro operates normally and the vertical gyro switch has been selectedto the AUX position on the control panel. Refer to your MEL.



B737-200

Engine Fuel Valve Closed Light

• The engine fuel valve closed light has two modes of illumination. Whenthe engine fuel shutoff valve is in the transit mode, the light will beilluminated a bright blue. A dim illumination of the engine fuel valve lightindicates the valve is in the closed position. With the light extinguished,the valve is in the open position. The engine fuel shutoff valves requires28V DC power from the hot battery bus operate the light and valve.

• On basic aircraft, the engine fuel valve light works in conjunction withthe fire switch handle. When selected, the light illuminates. On ad-vanced aircraft, the light operates in a similar manner, but with the fuelcutoff lever. The two fuel shutoff valves are located on the front sparsoutboard of the engines.

Fuel Temperature Indicator

• The fuel indicator is a resistance ratiometer type unit. The dial iscalibrated from -560C to 560C and requires 28V AC power to operate.The fuel temperature bulb (probe) is located in the No.1 tank, near theaft spar section of the fuel tank. The bulb itself is an armored, resis-tance-type probe which projects into the tank through the rear spar.



B737-200

• (Oral Topic) The No.1 wing fuel tank has been selected as the site ofthe temperature probe because it contains the coldest fuel. This isbecause the hydraulic system A heat exchanger is located in that tank.System A heat exchanger is smaller than system B heat exchanger(located in the right wing tank). Therefore, less heat is absorbed intothe fuel tank, thus producing the coldest fuel.

Fuel System Limitations

• The maximum limitation for fuel temperature is 490C. The minimumfuel temperature limitation is the freezing point plus 30.

• The maximum fuel quantity is 10,120 lbs. per each main wing tank and16,351 lbs. for the center tank. Total fuel is 36,591 lbs.

• The maximum fuel lateral imbalance between main tank No.1 and No.2is 1,500 lbs. for taxi, takeoff, and flight. With a maximum of 1,300 lbs.imbalance for landing.

• Fuel loading for the main tanks must be full if center tank contains morethan 1000 lbs. With less than 1000 lbs. in the center tank, partial maintank fuel may be loaded, provided the effects of balance have beenconsidered.

• Fuel usage must be planned to use center tank to depletion, followed bymain tank fuel. It is advised that maintenance must be contacted priorto adding any type of fuel other than Jet-A.

Fuel Filter Icing Light

• The illumination of the amber fuel filter icing light indicates an iced orcontaminated fuel filter. Procedures call for the selection of the startswitches to the ON position, the fuel pump switches selected ON, andboth fuel heat switches be activated for a one-minute. The caution lightshould extinguish prior to the end of the one-minute cycle. One fullcycle is recommended to restore the filter to its maximum filtrationcapacity. Should the light remain illuminated, and with fuel tempera-tures at or below 00 C, position both fuel heat switches to ON for oneminute every 30 minutes of flight.



B737-200

• The B737-200 uses 13th stage, pneumatic fuel heater to increase thefuel temperature. This helps to prevent the blocking of the fuel filter dueto icing. The illumination of the filter icing light indicates the filter isblocked with possible ice.

Fuel Heat Valve Open Light

• The blue fuel heat valve open light has three modes of operation (illumi-nation). A bright illumination indicates the fuel heat valve is in a transitmode or in disagreement with fuel heat switch position. A dim illumina-tion of the fuel heat valve light indicates the valve is in the openedposition. And, with the valve open light extinguished, the fuel heat valveis in the closed position.

• (Oral Topic) To confirm the fuel heating process, observe the illumina-tion of the fuel heat valve open light and the increase of oil temperature(approximately 5-100 increase), as displayed on the oil temperature indi-cator.

Fuel Heat Switch

• The fuel heat solenoid switch controls the respective engine fuel heatvalve by allowing 13th stage bleed air to heat. By increasing the fueltemperature, it helps to prevent the blocking of the fuel filter due toicing. When outside air temperatures are 00C or colder, normal proce-dures calls for fuel heat applications of one minute after engine starting.The switch automatically moves to OFF after one minute.

• A fuel heat valve failure is recognized by the valve open light remainingilluminated bright blue. This indicates that the fuel heat valve is indisagreement with the fuel heat switch. Observe the following oil tem-perature limitations: maximum temperature of 1570C, maximum con-tinuous temperatures between 1200C to 1570C is limited for 15 minutesof operations, and maximum continuous temperatures of 1200C with notime limit. Engine operation may be continued provided that the engineoil limits are maintained within the required limits. Should the fuel heatvalve fail to open when selected (fuel filter icing light illuminated), moni-tor the engine fuel flow and use caution due to the possibility of enginepower loss due to fuel icing.



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• Dispatch with an inoperative fuel heat timer (auto cycle function) isallowed providing the associated fuel heater valve light operates nor-mally. During flight operations that require fuel heat, manually operatethe fuel heat switch for one minute cycles. Remember, oil cooling willbe degraded by lengthy fuel heat operations since the engine oil is fuelcooled. Refer to your MEL for details.

• Dispatch with an inoperative fuel heat valve is allowed providing variousMEL procedures are followed. Considerations for departure fuel tanktemperatures, enroute ambient temperatures, and recommended alti-tude/speed schedules are followed per MEL charts. Refer to your MELfor details.

Crossfeed Valve Open Light

• The blue crossfeed valve open light has two modes of illumination.When the crossfeed valve is in the transit mode, the light will be illumi-nated a bright blue. A dim illumination of the crossfed valve light indi-cates the valve is in the opened position. With the light extinguished,the valve is in the closed position.

• With an inoperative crossfeed valve light, the aircraft may be dispatchedproviding the fuel quantity indicator for the main tanks are operatingnormally. Refer to your MEL.

Crossfeed Selector

• The purpose of the crossfeed valve selector is to provide the means ofdirecting fuel to both engines from a single tank. When selected open,engine No.1 and No.2 fuel feed lines are connected. Power for thevalve and light operation is from 28V battery bus, with the circuitbreaker location on the P6 panel.

• Crossfeed valve failure is indicated by the bright illumination of thecrossfeed valve open light. Procedures for a failed valve in the CLOSEDposition require fuel balance to be maintained by varying the thrust. Afailed valve in the OPEN position, fuel balance may be maintained byselective use of the fuel pumps.

• The crossfeed system is located on the forward leading edge spar.



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Manual Defuel Valve(Located inboard of fueling station, right wing)

Fuel Transfer

• Fuel transfer can be accomplished by using the fuel boost pumps, rightwing fueling station, and the manual defueling valve. The six fuel boostpumps, two per tank, deliver fuel under pressure from the No.1 wingtank, No.2 wing tank and the center tank.

• The manual defueling valve is located outboard of the No.2 engine andjoins the engine feed system with the fueling station. A red handle valveselector is used to open the valve manually. With the handle in theOPEN position, the access panel can not be closed. AC power isrequired to operate the boost pumps and valves.

Fuel Tank Transfer

1. Sender tank boost pumps on.2. Crossfeed selector valve open.3. No. 2 start lever idle.4. Defueling valve open.5. Receiver tank fueling valve open.6. Fuel transfer monitor.7. Transfer complete, reverse the procedures.



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Aircraft Defueling

• (Oral Topic) Defueling can be accomplished by using the fuel boostpumps and the single point fueling station during ground operationsonly. Suction defueling method may also be used as an alternatemethod of defueling the aircraft. The manual defueling valve must beopen in all cases of defueling.

• Procedures require the fueling hose to be attached to the fueling recep-tacle. The fuel flows from the operating boost pumps, through the No.2engine fuel shutoff valve, through the manual defueling valve, then tothe fueling manifold station. A fueling hose connected to the fuelingstation directs the fuel to an external storage tank or truck. AC electri-cal power is required to operate the boost pumps and valves. MinimumNo.2 tank fuel for hydraulic B pump operation is 1676 lbs. Flaps do nothave to be extended during this operation. Refer to your AFM forfurther details.

Defueling

1. Sender tank boost pumps on.2. Crossfeed valve open (required for tank No.1 defueling).3. No.2 start lever open.4. Defueling valve open.5. Defueling levels monitor.

Aircraft Refueling

• Normal refueling requires 115V AC and 28V DC from the ground powercart, APU, or from the battery. This is required for the operation of thefuel quantity indicators and the refueling valve circuits. Fueling proce-dures require filling the wing tanks full first with equal amounts. Whenadditional fuel is required, load center tank next.

• Maximum nozzle pressure is 50 PSI, this will be approximately 300 USgallons (2010 lbs) per minute. The main tanks may also be filledthrough overwing ports. The center tanks can then be filled using thetransfer procedure. The fuel tanks can be filled to any desired amountby using the fueling control station located on the right wing.



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• Float switches located in the respective fuel tank sense tank quantitiesand can automatically stop the fueling process as selected on the fuelcontrol station. When the fueling station access door is opened, thefueling power switch is actuated and the panel is illuminated by whitelights.

• A auxiliary fueling power control switch has been provided as an alter-nate method of powering the system. By selecting this switch, 28 VDCpower is provided to illuminate the fueling bay, valve position switches,and the press-to-test function of the position lights.



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Fuel Boost Pump Power Source

• (Oral Topic) The above schematic displays the power source for eachfuel boost pump. Please note that the respective aft boost pump ispowered by the associated transfer bus. The left forward and rightcenter tank boost pump is powered by the No.1 main bus. The rightforward and left center tank boost pump is powered by the No.2 mainbus. The boost pump power sources are arranged to ensure that inthe event of a single generator failure, at least one pump in each tankwill remain powered.



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Center Tank Fuel Pump Low Pressure Light

• The illumination of a center tank fuel pump low pressure light is afunction of low pressure and pump switch position. The fuel pumpswitch must be in the ON position to arm the low pressure light. Theillumination of both fuel low pressure lights will cause the illumination ofboth the master caution light and the fuel annunciator light.

• With only one low pressure light illuminated, the fuel annunciator lightwill illuminate only on RECALL selection.

• Crew action for one center tank fuel pump low pressure light illuminated,calls for the crossfeed selector to be selected to the OPEN position.This will prevent a fuel imbalance. For both center tank low pressurelights illuminated, center tank fuel will be unusable. During takeoff, withcenter tank quantity less than 2000 lbs, the low pressure lights mayilluminate due to high deck angles. Center tank pump lights may beinoperative for dispatch providing the center fuel quantity gauge is oper-ating normally. Refer to your MEL.



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Center Wing Fuel Boost Pumps

• The center fuel tanks have two AC powered centrifugal impeller typepumps used to deliver fuel under pressure to the engines. The pumpsmay also be used for fuel transfer and fuel tank defueling (groundoptions only). The center wing fuel boost pumps are located in the wing(not in the center wing tank), one on either side of the fuselage. Thepower source for the right CWBP is the generator bus No.1 and the leftCWBP is the generator bus No.2. Therefore, if engine No.1 is lost, youwill lose power to both the forward No.1 WTBP (wing tank boost pump)and the right CWBP. Refer to the boost pump power source section foradditional information.

• The left center wing boost pump is considered the “AFT” pump, sincethe fuel pickup point is in the aft section of the center tank. The actualphysical location of the pump is in the left wing, next to the main tank aftboost pump. The right center wing boost pump is considered the“FWD” since the fuel pickup point is the forward section of the centertank.

• The CWTB pumps are not “override” type boost pumps. They aredesigned with a spring-loaded-closed flapper valve that is less restrictivethan those in the main fuel tanks. These check valves open at a lowerdifferential pressure (1.5 psi) than the check valves in the No.1 andNo.2 tanks (12.5 psi). This ensures that center tank fuel is used beforemain tank fuel, thus preventing reverse fuel flow to the respective boostpump. Continual dry running of the CWBP's has been demonstrated tobe detrimental and has been a major cause of premature pump remov-als and/or pump failures.

• The one center wing boost pump may be inoperative for dispatch pro-viding the center tank fuel is not required. The tank remains empty, andall zero fuel weight factors must be reviewed. During flight operations,with one CWBP inoperative, open the crossfeed valve to prevent fuelimbalance. If both CWBPs are inoperative, close the crossfeed valveand plan your flight with the center tank fuel as unusable. Refer to yourMEL.



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Main Tank Fuel Pump Low Pressure Light

• The illumination of the main tank fuel pump low pressure light is afunction of low pressure only. The illumination of both fuel low pressurelights will cause the illumination of both the master caution light and thefuel annunciator light. With only one low pressure light illuminated, thefuel annunciator light will illuminate only on recall selection. The powersupply for the pressure indication circuit is 28V DC power from the P6circuit breaker panel.

• The illumination of only one main tank fuel pump low pressure lightrequires no immediate crew action. Sufficient fuel pressure is availablefor normal operations. If both main tank low pressure lights are illumi-nated, caution should be taken at altitudes above 30,000’. Thrust andengine conditions may deteriorate.

Fuel Pump Switch

• Selecting the fuel pump switch to the ON position, causes the activationof the respective fuel pump. Fuel pumps are capable of suction feedingin the event that normal electrical fuel pump action is not available. Theengine pumps draw fuel through a bypass valve located in either theNo.1 or No.2 tank. These bypass valves may also be used for suctiondefueling.

• The boost pump power sources are arranged to ensure that in the eventof a single generator failure, at least one pump in each tank will remainpowered. The main tank aft pumps are powered from the transferbusses so that at low fuel level, if a generator is lost, both aft pumpsremain powered. The remaining pumps are powered from the mainbusses. The main tank forward pumps are installed on the front spars.The main tank aft and center tank pumps are located in bays in themain tanks.



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• One fuel boost pump may be inoperative in each tank for dispatchproviding a minimum of 4800 lbs of fuel is carried in the respectivetank. The reason for this extra 4800 lbs is that in the event of asecond boost pump failure, the engine suction feed bypass valve willremain covered with high deck angles during takeoffs and go-arounds. Review to your MEL for reference to procedures requiringthe selection of the crossfeed valve to the OPEN position when therespective tank reaches the 5800 lbs level point. Another consider-ation with respect to an inoperative boost pump and fuel loading areAFM symmetry limitations. Extra considerations include proper fuelheat application procedures with both boost pumps inoperative.

Note: Contact your maintenance representative as to the type of pumpsinstalled in your aircraft. Inoperative procedures will differ for type offuel boost pump installed.

Fuel Tanks

• The fuel tanks consists of three tanks for fuel storage and two ventsurge tanks for temporary fuel storage. The fuel tanks are largerthan the noted fuel capacity to allow for expansion and vent space.The surge tanks are normally empty.

• The fuel main tanks are designed as part of the primary wing structure,located between the front and rear wing spars and between the upperand lower wing skin. The center tank is contained within the fuselageand is divided into three cavities by spanwise beams.

• The fuel vent system is designed to prevent damage to the fuel tank byproviding positive venting (regardless of the attitude of the aircraft). Thefuel vent system helps in decreasing fuel evaporation and provides apositive head pressure on the fuel.

• (Oral Topic) The purpose of the surge tank is to collect fuel overflowpassing through the vent channels. This fuel overflow is then returnedthrough a surge tank drain into the center tank. The surge tank islocated at the end of each wing tank.



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Fuel Tank Capacities

Tank No.1 1499 (10,120 lbs)Tank No.2 1499 (10.120 lbs)Center Tank 2313 (16,351 lbs)Total 5311 (36,591 lbs)



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Electrical System Description

• AC power (alternating current) is normally supplied by two engine drivengenerators for ground and flight requirements. The generator’s rotorsare rotated at constant speeds. The CSD’s (constant speed drive)provide a constant frequency. A third generator, driven by the APU, canalso provide power for ground and flight requirements. The constantfrequency produced by the APU generator is controlled by a speedreduction control system and by various gearboxes. AC power can alsobe supplied by a fourth source, through the AC external power recep-tacle.

• AC power is connected to the main busses through the use of generatorbreakers and external power contactors. AC power can also be sup-plied to the transfer busses through the use of transfer relays. Externalpower can also power the external AC bus and the ground service bus.Control and indication of the AC electrical system is from the P5 over-head panel in the flight deck.

• DC power (direct current) is supplied from three different sources.Those sources are the battery, three transformer rectifiers, and from theexternal power system. This power is connected to the two DC bussesand the battery bus by relays. The battery charger is supplied withpower from the 115V AC system. Control and indication of the DCelectrical system is from the P5 overhead panel in the flight deck.



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• Standby AC and DC busses are normally powered from the respectiveAC and DC system. As an alternate power source, the battery canprovide power to the standby system.

DC Ammeter

• The DC ammeter displays amperage indications of the selected sourceby the DC selector. The ammeter will display indications from TR No.1,TR No.2, TR No.3, and/or the battery. The STBY PWR position and theBAT BUS position (as indicated on the panel), will not show amperageindications. During preflight, the TRs (transformer rectifiers) may bechecked by observing a positive amperage indication on the DC amme-ter.

DC Voltmeter

• The DC voltmeter displays voltage indications of the selected source bythe DC selector. The voltmeter will display indications from TR No.1,TR No.2, TR No.3, battery, standby power bus, and the battery bus.

Frequency Meter

• The AC frequency meter displays frequency of the power source se-lected by the AC meter selector. The frequency of the generator isdependent on the speed of the CSD. Frequency will be indicated only



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when the generator is electrically excited. For dispatch purposes, thefrequency meter may be inoperative providing the APU generator is notrequired for flight operations (ground operations not included). Refer toyour MEL for details.

AC Voltmeter

• The AC voltmeter displays voltage of the source selected (130V scale)from the Phase B line current. The voltmeter also displays residualvoltage of generator selected when residual volts switch is pressed.(30V scale)

DC Meters Selector

• (Oral Topic) The DC meters selector selects the DC source for the DCvoltmeter and DC ammeter. When selected to the STBY PWR or BATBUS positions, amperage will not be indicated since these are notsources of power. For aircraft shutdown and/or flight termination, selectthe DC meters selector off the BAT position. This will prevent a batterydischarge since the DC voltmeter will continue to require power from thebattery to operate the display indication. The TEST position is used bymaintenance and connects the voltmeter and frequency meter to apower system test module for the selection of additional readings.

AC Meters Selector

• The AC meters selector selects the AC source for the AC voltmeter andfrequency meter display indications. The TEST position is used bymaintenance to monitor the selections of the power system test modulein the P-6 panel. Both the AC and DC selectors should be placed indifferent position for different configurations of flight operations. Duringthe standby power check, the DC and AC meter selectors should beplaced in the STBY PWR positions to check for proper voltage andfrequency. For normal inflight operations, the DC and AC meter selec-tors should be placed at the BAT and STBY PWR positions respec-tively. For aircraft shutdown and/or flight termination, select the DCmeters selector off the BAT position. This will prevent a battery dis-charge since the DC voltmeter will continue to require power from thebattery to operate the display indication.



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Residual Volts Switch

• (Oral Topic) The primary purpose of the residual volts switch is todetermine if a generator has been disconnected. The residual voltsswitch when selected, displays residual voltage of the selected genera-tor on the 30 volt scale of the AC voltmeter. Pressing the switch with adisconnected generator causes the voltmeter to display zero volts. Anindication of 10 to 15 volts on the voltmeter shows evidence that thegenerator is still rotating. The permanent magnets within the generatorproduces the voltage displayed. The permanent magnets are mountedon the exciter generator frame. These magnets provide a “built-in”residual voltage which accrues voltage buildup and eliminates the needfor field flashing. As an operational suggestion, maintenance normallyrequest an residual volts reading before and after CSD disconnects.

Galley Power Switch

• Selecting the galley power witch to the ON position provides electricalpower to the galleys. Galley power is available only when generator busNo.1 and No.2 are powered. Circuit breakers for galley power arelocated in the galley and on the P-6 circuit breaker panel. This switchshould be considered during the Electrical Smoke or Fire procedure. Byselecting the galley power switch to the OFF position, removes highload items from the aircraft’s electrical system.

• Galley power availability is part of the automatic load shedding protec-tion. With a generator bus failure, the respective B system hydraulicpump and galley power is lost. If the opposite hydraulic pump is alreadyswitched off, the remaining hydraulic pump power source will transfer tothe main bus that is powered.

Battery Switch

• (Oral Topic) The battery switch ON/OFF positions have several areasof importance with reference to battery bus operation and powersources for that bus. Selecting the battery switch to the OFF position,will cause the battery bus not to be powered. Even with both enginesoperating (generators on line), selecting the battery switch to the OFFposition will only de-energize the battery bus. Think of this switch as a"battery bus switch". Selecting the battery switch to the ON position(No.2 main bus energized), causes the No.3 TR to furnish power to the



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battery bus. This is known as the primary power source for the batterybus. If the No.2 main bus is not energized, the hot battery bus powersthe battery bus. This is known as the alternate power source for thebattery bus.

• (Oral Topic) The OFF position of the battery switch has a direct affecton the operation of the APU. Selecting the battery switch to the OFFposition (with the APU is operating), will cause an auto-shutdown of theAPU. This automatic APU shutdown feature is a ground item only,during flight operations, APU operations will not terminate. The batteryswitch also provides the AC system with control and protection func-tions. For example, the battery switch must be in the ON position toprovide fire detection capabilities.

Battery

• (Oral Topic) The primary purpose of the battery is to provide DC powerto the standby buses when normal DC power supply has been disruptedfrom the TRs (transformer rectifiers). The secondary purpose of thebattery is for starting the APU. A minimum of 22 volts is required forAPU starting.

• A fully charged battery has sufficient capacity to provide power for aminimum of 30 minutes. The typical 20 cell nickel-cadmium battery islocated in the electronics compartment. The battery contains harmfulfluids, extreme caution should be used when handling the battery. Incase of spilled electrolyte fluids, clean your hands with water or with a3% boric acid solution.

• (Oral Topic) Following the loss of both AC generators, the battery willprovide power to the battery bus, DC standby bus, hot battery bus, andthe switched hot battery bus. The battery must be above the minimumvoltage to operate units supplied by the bus. If the APU is the onlyoperating generator, connect it first to the No.2 bus (as it will power TRNo.3). If the APU cannot be connected to the No.2 bus, connect it tothe No.1 bus. Loss of both engine driven generators is normally indi-cated by the illumination of the TRANSFER BUS OFF, BUS OFF, andthe GEN OFF BUS lights. Various other instrument warning lights, andMaster Caution System lights will also be illuminated.



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• (Oral Topic) The condition or quality of the battery (duration time foremergency power use), can be determined by the type of chargingphase the battery is in. Remember, a battery is completely chargedwhen pulsing charges are indicated on the ammeter. Therefore, you willhave approximately 30 minutes of emergency power available with a"fully charged battery".

Battery Charger

• (Oral Topic) The primary source of power for the battery charger isprovided by the ground service bus, via the No.1 generator bus. Thealternate power source for the charger is from the No.2 main bus, viathe charger relay. When the battery power is low and requires morethan 26 amps of charging from the battery charger, the charger will actlike an unregulated transformer rectifier. When the battery is completelycharged (battery charger is delivering less than 26 amps charging cur-rent), the battery charger reverts to a pulsing charger. Battery chargeroperation can be checked by selecting the DC meter to BAT. Observeeither a steady charge or a pulsing charge on the ammeter. Pulling thebattery charger CB will enable you to observe the actual voltage of thebattery. The charger is rated 40 amps with forced air cooling.

• The electrical system has been designed with a charger relay feature.This provides the means of a power source transfer for the batterycharger in the case of No.1 generator bus failure (power source for theground service bus). The charger relay closes and the No.2 main busnow becomes the power source for the battery charger.

• During APU starting, AC power to the battery charger will be interruptedto prevent heavy power draw from the charger. When the APU reaches50% RPM, various relays will relax and the charger will revert to normalfunctions.

TR (Transformer Rectifier)

• The purpose of a TR (transformer rectifier) is to convert 115V AC, 400Hz, 3-phase power to 28V DC. The aircraft has three main TR unitswhich are located on the E3-1 shelf in the E/E compartment. Each unithas been rated at 65 amps with cooling and 50 amps without cooling.



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• The transformer rectifiers are labeled as TR No.1, No.2, and No.3. TheNo.1 TR is supplied with power from the No.1 transfer bus and providesDC power for the No.1 DC bus. No.2 TR is supplied power from theNo.2 transfer bus and provides DC power to the No.2 DC bus. No.3 TRis supplied power from the No.2 main bus and provides DC power to thebattery bus and acts as an alternate DC power source for TR No.1 orTR No.2 failure.

• (Oral Topic) A common oral question that examiners may ask withreference to transformer rectifiers and the electrical metering panel isthe source of displayed information. For example, the selector hasbeen placed to either TR 1 or TR 2, what is the source of the informa-tion that is being displayed on the ammeter and/or voltmeter? Thecorrect response, with TR 1 or TR 2 selected, the ammeter will displayinformation from the TR unit itself and the voltmeter will display informa-tion from the respective bus. TR 3 is different, it will display voltage andamperage information obtained from the TR 3 unit only.

• (Oral Topic) Another common oral question with reference to the elec-trical metering panel is the display indications of a failed TR unit. Afailed TR 1 or TR 2 unit would be indicated by zero amps and normalbus voltage. A failed TR 3 would be indicated by zero amperage andzero voltage.

• Two of the three transformer rectifiers are required for dispatch. TRNo.2 may be inoperative provided that all DC buses and all generators(including the APU generator) operate normally and the APU generatorcan be electrically connected to either bus. Refer to your MEL.



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CSD - Constant Speed Drive

• The aircraft generators produced AC power that must be of a constantfrequency. To achieve this constant frequency, the CSD has beenprovided to control the generators at an constant speed, thus providingconstant frequency. Example of the relationship between speed andfrequency: 6000 RPM of generator speed, corresponds to the electricalfrequency of 400Hz. The CSDs are a hydromechanical unit, internal oilis used as an operating fluid and for cooling purposes. The APU doesnot have a CSD. The APU generator operates at a constant speedthrough the use of internal gearboxes.

• Operating conditions of the CSD can be observed on the generatordrive oil temperature indicator. Oil temperature is measured as it entersand leaves the CSD. Temperature of the oil entering the CSD is indi-cated on the IN scale. Temperature differential between outlet and inletis indicated as RISE (out temperature minus in temperature). The "IN"temperature sensor for the CSD is downstream from a fan air cooler.This temperature reflects the ability of the fan air to cool the CSD oil.The "RISE" temperature is a "comparison" of the IN and OUT tempera-tures, and only reflects how hard the CSD is working.

Generator Drive Low Oil Pressure Light

• The illumination of the amber generator drive low oil pressure lightindicates the CSD oil pressure is below the minimum operating limit of120 PSIG. With the illumination of the light, a malfunction in the CSDcan be assumed. The CSD should be disconnected before additionaldamage occurs. The illumination of the low pressure light will alsoilluminate the master caution and ELEC annunciator lights.

• The generator drive low oil pressure light may be inoperative for dis-patch provided the frequency meter and the respective CSD high oillight and oil temperature indicator operates normally. Refer to yourMEL.



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Generator Drive High Oil Temperature Light

• The illumination of the amber generator drive high oil temperature lightindicates the CSD oil has exceeded the operating temperature limit of1570C. Two causes for high oil temperatures are low oil quantity andimproper internal mechanical operation of the CSD. With high CSD oiltemperatures, CSD failure may occur. To prevent damage to the CSD,procedures calls for the CSD to be disconnected.

• The illumination of the high oil temperature light will cause the illumina-tion of the master caution and ELEC annunciator lights. The generatordrive high oil temperature light may be inoperative for dispatch providedthe frequency meter, the respective CSD low oil pressure light, and/orthe oil temperature indicator operates normally. Refer to your MEL.

Generator Drive Disconnect Switch

• The generator drive disconnect switch controls a 28V DC disconnectsolenoid. This disconnect solenoid has been designed to prevent thepossibility of any voltage pickup that may inadvertently trip the CSD. Byselecting the guarded switch to the UP position, the CSD will be me-chanically disconnected from the engine. The re-connection of the CSDcan only be accomplish on the ground by maintenance personnel, repo-sitioning the external reset handle. Electrical power for disconnecting



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comes from the battery bus. The residual volts button is normally usedduring the CSD disconnect procedures. Selecting the residual voltsbutton, causes the display of “zero volts” on the AC voltmeter. After theCSD has been disconnected, the low pressure light will remain illumi-nated.

Generator Drive Temperature Switch

• The generator drive temperature switch may be used to select either theRISE or IN temperatures. These temperature values are displayed onthe generator drive oil temperature indicator. Two variable resistancetype temperature bulbs (probes) measure the oil temperature on eitherside of the CSD oil cooler.

Generator Drive Oil Temperature Indicator

• The generator drive oil temperature indicator displays temperature asselected by the generator drive temperature switch. A higher thannormal temperature may indicate either excessive generator load or apossible poor mechanical condition of the generator drive unit. Lack ofadequate cooling may also cause the temperature to increase.

• A generator drive oil temperature indicator may be inoperative for dis-patch provided the frequency meter, the respective CSD low oil pres-sure light, and/or the high oil temperature light operates normally. Referto your MEL.

CSD Limitations

• The maximum CSD oil temperature is 1570C. The maximum CSD oiltemperature when selected to the RISE position is 200C.

Standby Power Switch

• The standby power switch is a three position, guarded switch (BAT,OFF, AUTO). When selected to AUTO (right guarded position), thestandby AC bus will be powered by the No.1 transfer bus. The standbyDC bus will be powered by the No.1 DC bus. When selected to OFF(center position), the static inverter is not powered, thus not poweringthe standby AC bus. At this time, the standby power off light will be



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illuminated indicating the standby buses are not powered. When se-lected to BAT (left position), the R328 relay is energized, thus poweringthe static inverter and providing power to the standby AC bus. Thestandby DC bus will receive power directly from the battery bus.

• The AUTO position has some additional features that should be noted.During normal in-flight and ground operations, the standby AC bus ispowered by the No.1 transfer bus. The standby DC bus is powered bythe No.1 DC bus. With the loss of AC power (in-flight only), the standbyAC bus is automatically powered by the battery bus through the staticinverter. The standby DC bus is automatically powered directly by thebattery bus. With the loss of AC power (ground only), there is noautomatic transfer of power.

• The battery can furnish power to the standby bus equipment for aminimum of 30 minutes. The battery bus is powered by the hot batterybus, regardless of the battery switch position.

Standby Power Off Light

• The illumination of the amber standby power light indicates the AC and/or DC standby buses are not powered after a loss of both generatorbusses. The master caution light and ELEC annunciator will illuminate.The standby power switch should be selected to BAT position. This willprovide power to the standby buses from the battery. A fully chargedbattery will provide a minimum of 30 minutes of standby power.



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Generator AC Ammeter

• The generator AC ammeter displays engine generator load in amperes.The maximum engine driven generator load limit is 111 amps. Theammeter may be inoperative for an inoperative generator. Refer to yourMEL.

Ground Power Available Light

• (Oral Topic) The illumination of the ground power available light indi-cates the external power bus is powered by a ground power supplydevice. The light will remain illuminated as long as the ground powerdevice is plugged in. The light provides no quality assurances that thepower is correct (frequency/amps), just that it is connected to the air-craft.

Ground Power Switch

• The ground power switch is a three position (OFF, ON, and neutral)switch. The switch is spring-loaded to the neutral position. Whenselected to the OFF position, ground power will be disconnected fromboth generator buses. When momentarily moved to the ON position,external power will be connected to the both generator buses.



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• The following items occur when connecting ground power to the aircraftbuses:1. Removes previously connected power source from both generator

buses.2. Closes external power contactor and connects ground power to both

generator buses if power quality is correct.3. Switches the ground service bus to the generator bus No.1.4. Deactivates the ground service switch.5. Allows the battery to be charged from the external AC power supply.

Ground Power

• AC external power receptacle provides 115V AC, 3-phase power froman external power source (cart or ground power unit). The AC recep-tacle is located on the right side of the airplane, forward of the nosewheel well. The receptacle has four long AC pins, for phases A, B, C,and N (neutral/ground) and two short DC interlock pins E and F. Thereason for the short DC pins is to prevent arcing or flashing (by the ACpins) should the external power cable be inadvertently removed whileAC power is being applied. Power for the E and F pins will disconnectbefore the AC pins disconnection, thus providing power interruption pro-tection for the AC circuits prior to activation.



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• The 28V DC power receptacle is located near the battery in the elec-tronic compartment. When the external DC power is connected, itoperates in paralleled with the battery and will power all circuits nor-mally supplied by the battery. In the event that the airplane battery isdepleted, the APU can be started by using DC external power. Exter-nal DC power is not intended to be used as a power source for batteryrecharging. Battery re-charging should always be accomplished by thebattery charger.

• The DC power receptacle consists of two large pins (positive and nega-tive) and a third small pin for correct orientation of the external connec-tor. Before connecting the power receptacle, voltage should rangebetween 24-28V DC only.

Ground Service Switch

• The ground service switch is located on the forward flight attendant’spanel. The switch has been designed to provide power to the groundservice items. Such items include various service lights and interiorservice outlets. The ground service power system makes it unneces-sary to energize the main aircraft’s AC and DC buses when power isrequired for ground service items.



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Ground Service Bus

• (Oral Topic) The ground service bus provides power to:1. Battery charger.2. Equipment cooling switch (normal position).3. Service lights.4. Service outlets.

Bus Transfer Switch

• The bus transfer switch is a two position switch (AUTO - OFF). Theswitch is normally in the AUTO guarded position. This allows the auto-matic transfer of the transfer bus (essential electrical loads) upon failureof the associated generator bus. This automatic function also allows TR3 to supply the No.1 DC bus if TR No.1 fails. A TR failure can bedetected by a zero reading on the DC ammeter of the selected TR.



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• Selecting the bus transfer switch to the OFF position, the followingitems occur:1. Isolates the transfer buses by preventing the operation of the bus

transfer relays, and opens the TR 3 disconnect relay.2. Prevents the battery charger from switching to its alternate source

of power (main bus No.2)3. With normal power on the generator buses No.1 and No. 2, opens

TR 3 disconnect relay.4. Transfer of alternate power for the "B" pumps.5. Transfer of alternate power for the standby hydraulic pump.

Automatic Load Shedding

• Automatic load shedding should be discussed at this time with referenceto bus transfer. The concept of load shedding provides the capability toreduce power demands automatically during single AC generator opera-tions. Automatic load-shedding will turn off all nonessential, high drawelectrical equipment such as:1. Galley power.2. Respective hydraulic system B electric pump. If one system B hy-

draulic pump switch is already selected OFF, the remaining systemB hydraulic pump power supply will be transferred to the main busthat is powered. The standby hydraulic pump also has power trans-fer capabilities if the No.1 bus is not powered.

Transfer Bus Off Light

• The illumination of the amber transfer bus off light indicates the transferrelays are not energized. Therefore, the transfer bus is inactive. Themaster caution light and ELEC annunciator will illuminate.

• Non-normal procedures calls for the bus transfer switch to be selectedto the OFF position, then to AUTO. Check the standby power off light isextinguished, this will ensure the integrity of the standby buses foressential equipment power.

• With the failure of the AC transfer bus (transfer bus light illuminated),the emergency instrument flood lights will automatically activate.



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BUS OFF Light

• The illumination of the amber BUS OFF light indicates the generator buson the respective side is inactive and its respective engine generatorbreaker, APU generator breaker and external power contactors areopen. Master caution light and ELEC annunciator will illuminate. Therespective illumination would indicate the following:1. (Generator No.1) Loss of power to the generator bus No.1, the

main bus and the ground service bus.2. (Generator No.2) Loss of power to the generator bus No.2 and the

main bus No.2.

• Non-normal procedures calls to ensure the generator switch has beenselected ON. Should the BUS OFF light remain illuminated, start theAPU and place on line (if available). If both generator buses are notpowered, the galleys and one system B hydraulic pump are not pow-ered.

• (Oral Topic) Illumination of the BUS OFF light indicates the respectivebus is inactive. With this type of a failure, you will have the followingvisual indications:1. Illumination of the respective low fuel boost pump pressure lights.2. Galley power will be inoperative.3. Illumination of the respective BUS OFF, GEN OFF BUS lights.4. Illumination of the respective window heat lights.5. Illumination of the respective hydraulic System B pump low pres-

sure lights.6. Respective landing lights will be inoperative.7. Illumination of the equipment cooling off light.

Generator Off Bus Light

• The illumination of the blue generator off bus light indicates the genera-tor is not supplying the generator bus and the respective generatorbreaker is open.



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Dual Generator Loss

• The illumination of the GEN OFF, BUS OFF, and the TRANSFER BUSOFF lights indicates the possible loss of both engine driven generators.Upon loss of all generators, the electrical system will automaticallyswitch to standby power. Essential radios, flight instruments, and navi-gation equipment will be powered by the standby system. Non-normalprocedures directs your attention to reduce electrical loads before re-selecting a generator and/or the activation of the APU generator. Thefollowing is a summary of those actions.

1. Select the galley power switch to OFF.2. Select the bus transfer switch to OFF.3. Select system B Hydraulic pumps switches to OFF.4. Re-select the generator switches to ON.5. If either or both BUS OFF lights remain illuminated, bring on line the

APU. If the APU generator is the only operating power source,connect it to the No.2 bus first, so that the TR 2 and TR 3 arepowered. If the APU will not be connect to the No.2 bus, thenconnect it to the No.1 bus.

Generator Switch

• The generator switch is a three position, spring loaded (OFF-ON)switch. It enables the engine driven AC generator to be connected tothe respective generator bus when power quality is correct. If thegenerator was de-excited, it will connect the field power supply to theexciter. Selecting the switch to the OFF position, the generator will bede-excited and thereby disconnect itself from the generator bus.

• Either engine driven generator may be inoperative for dispatch providedthe APU generator is operating normally and the APU fuel heater oper-ates normally when fuel temperature less than 320C is anticipated. Ifthis can not be accomplished, then the flight duration must be less thanthat required for the fuel to cool to the above temperature value. Referto your MEL.



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• Each generator has an generator control unit (GCU), located in the P6panel. The GCU provides protection for the following items:1. Overvoltage 130 +/- 3 volts2. Undervoltage 100 +/- 3 volts3. Overfrequency 430 +/- 5 Hz4. Underfrequency 365 +/- 5 Hz5. Overcurrent 170-175 amps6. Differential Current 20-30 amps

APU Generator Off Bus Light

• (Oral Topic) The illumination of the blue APU generator off bus lightindicates that the APU is at the governed speed of 95% rpm, the APU isnot supplying a generator bus, the generator frequency is above 380Hz, and the APU generator is ready to accept a load.

APU Generator Switch

• The three position, spring loaded (OFF-ON) switch enables the APUgenerator to be connected to the desired generator bus when the powerquality is correct. If the APU generator is de-excited, it will connect thefield power supply to the exciter. When selected to the OFF position,the APU generator will disconnect itself from the respective generatorbus. If the other generator bus is not utilizing the APU generator, thenthe APU generator de-excites itself.

• (Oral Topic) The two APU generator switches work the same on theground with respects to connecting the APU generator to the respectivebus. During flight operations, only one bus can be powered from theAPU generator. However, it is possible to depart with both bussespowered by the APU, but it is not recommended due to the increasedelectrical loads on the APU generator.

• The APU generator is rated at 40 KVA inflight and 45 KVA while on theground.



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APU Limitations

• The following limitations apply to the APU generator. Some airlineshave deleted these numbers from their limitation sections of their manu-als. They have been included here as technical reference only.1. APU generator limit (ground) 125 amps.2. APU generator limit (flight) 111 amps.

APU Low Oil Quantity Light

• The illumination of the blue APU low oil quantity light indicates the APUoil quantity is insufficient for extended operations. The warning light isdisarmed when the APU switch is in the OFF position.

• (Oral Topic) The total oil quantity of the APU lubrication system is 1.5gallons. The illumination of the low quantity light only indicates the APUoil quantity is insufficient for extended operations. The actual oil re-maining within the system is at or below 1.25 quarts.

• The APU low quantity light may be inoperative for dispatch provided theAPU oil is checked prior to each departure. This quantity check allowsthe APU to be used during ground and flight operations. Refer to yourMEL.



B737-200

APU Oil Pressure Light

• The illumination of the amber APU oil pressure light indicates the APUoil pressure is low, causing the APU to initiate an automatic shutdown.This occurs only after the start cycle has been completed and at anytime during normal APU operations with oil pressures below 45 psig. Itis normal for the APU oil pressure light to illuminate during the startingcycle and remains illuminated until the APU oil pressure is normal. Thislight is disarmed when the APU switch is in the OFF position. The APUoil pressure light may be inoperative for dispatch provided the APU isnot used in flight. For ground operations, APU operations are permittedprovided the automatic shutdown features are operating normally. Re-fer to your MEL.

APU High Temperature Light

• The illumination of the APU high temperature light indicates the APU oiltemperature is excessive, causing the APU to initiate an automatic shut-down. This light is disarmed when the APU switch is in the OFFposition. With actual high oil temperatures, the high temperature lightcan be reset by placing the APU switch to OFF. The temperature of theoil must also decrease below temperature switch limit for the reset. Thetemperature sensor senses the oil temperature at the pressure pumpoutlet and is designed to illuminate the amber caution light when the oiltemperature reaches 255oF (124oC).

• The APU high temperature light may inoperative for dispatch providedthe APU is not used in flight. For ground operations, APU operationsare permitted provided the automatic shutdown features are operatingnormally. Refer to your MEL.

APU Fault Light (as installed)

• The illumination of the APU fault light indicates the APU oil temperature isexcessive, causing the APU to initiate an automatic shutdown. This light isdisarmed when the APU switch is in the OFF position. With actual high oiltemperatures, the APU Fault light can be reset by placing the APU switchto OFF. The temperature of the oil must also decrease below temperatureswitch limit for the reset. The temperature sensor senses the oil tempera-ture at the pressure pump outlet and is designed to illuminate the ambercaution light when the oil temperature reaches 124oC (255oF).



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APU Overspeed Light

• The illumination of the amber APU overspeed light indicates the APUspeed is excessive (greater than 110% RPM), causing the APU toinitiate an automatic shutdown. After shutdown, the APU cannot berestarted without resetting the overspeed circuit. The APU overspeedreset switch is located in the E/E compartment, must be reset beforeanother start attempt. The APU overspeed light is always armed, re-gardless to the position of the APU switch.

• The overspeed light also illuminates if an APU start is aborted prior toreaching governed speed, but extinguishes following a normal start.Light illuminated during APU shutdown indicates overspeed shutdownprotection is lost.

• The APU overspeed light may be inoperative for dispatch provided theAPU is not used in flight. For ground operations, APU operations arepermitted provided the automatic shutdown features are operating nor-mally. Refer to your MEL.



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APU Generator Ammeter

• The APU generator ammeter displays APU generator load current. Ob-serve the following limitations: ground operations = 125 amps; flightoperations = 111 amps. The above limitations have been removed fromsome airline’s AFMs, observe your airline’s limitations.

APU Exhaust Temperature Indicator

• The APU exhaust temperature indicator displays APU exhaust gas tem-perature. The thermocouple is located in the exhaust of the APU en-gine. Power supply of 0-5 volts DC is used only for the indicator integralillumination.

• The APU EGT indicator may be inoperative for dispatch provided allwarning and caution lights operate normally. The APU may only beused to provide the aircraft with electrical power, and for the starting ofonly one engine. Passengers are not permitted on board until the APUhas been shutdown. This is to provide extra protection and to guardagainst possible APU fire conditions. Refer to your MEL.

• Caution should be exercised during APU starting with regards to inad-vertent sticking of the APU start switch. This may affect the operationof the EGT indicator and provide erroneous indications.



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APU Start Switch

• The APU start switch is a three position switch (OFF-ON-START). Se-lecting the OFF position with the APU operating, initiates an APU shut-down. This shutdown feature simulates a 110% RPM overspeed andautomatically secures the APU.

• (Oral Topic) The ON position of the APU start switch is the normalposition for APU (running) operations. Selecting the APU start switch tothe START position begins the automatic start sequence. The air inletdoor and the fuel valve opens. As the door reaches the full openposition, the automatic APU starter engagement begins.

APU Start Sequence

• (Oral Topic) Some examiners may ask the start sequence of the APU.Listed below is the normal start sequence for the APU.

1. Position the battery switch to ON (minimum 22 volts). With batteryvoltage below 18 volts, certain battery switch relays may not close,thus preventing the APU from starting. The APU motor is normallypowered directly from the battery.

2. Fuel boost pump ON (not mandatory). However, with the fuel boostpump ON, it will help extend the service life of the APU fuel controlunit.

3. APU bleed switch OFF.

4. Complete the APU Fire Test.

5. APU switch is momentarily selected to the START position (observethe following):• Low oil pressure light illuminated.• Full scale negative deflection on DC ammeter for APU starter

engagement. The fuel valve opens and supplies fuel to theAPU. At 4%, fuel is injected. The APU inlet door is in the fullopened position. A battery charger relay opens to isolate thebattery charger.

• Low oil pressure light extinguished at approximately 37% RPM.



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• Monitor EGT. Max-760C.• Normal fuel burn: 250/300 lbs per hour.• Maximum starter operation time: 90 seconds.

• At 50% RPM, the starter is de-energized. If the APU rpm does notreach the starter cutout circuit requirement within 90 seconds, the startis automatically terminated. Observe DC ammeter for starter cutout.An increase above normal values indicates the charger is operating andrecharging the battery.

• At governed speed of 95% RPM, the ignition exciter is terminated andthe APU GEN OFF bus light will illuminate, the APU is now ready toaccept a load.

APU Alternate Starting - Power Source

• Normal power source for APU starting is the battery. An alternatepower source for starting the APU can be provided by a DC groundauxiliary cart through the 28V DC power receptacle. This receptacle islocated near the battery in the E/E compartment.

APU Aborted Starts

• After any unsuccessful ground start of the APU, do not attempt anotherground start (AD 90-05-02). A second/subsequent ground start attemptis permitted if a qualified observer is present after any unsuccessfulground start of the APU. That observer is required to watch for anypossible "unacceptable torching" or any other non-normal indications.With multiple aborted start attempts, five minutes of cooling is requiredbetween the second and third attempt. Wait one hour after the thirdattempt.

APU Inflight Start Attempts

• Inflight APU starts may be attempted at any altitude up to 35,000’, butaltitudes below 25,000’ are recommended. With at least one generatoroperating normally, subsequent start attempts (up to a maximum offour) can be made at lower altitudes.



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APU Fuel Control Unit

• Many airlines are upgrading their B737 APU’s with the timed accelera-tion fuel control unit (TAFCU). This modification is designed to reduceEGT during APU acceleration, thus improving hot section replacementtimes. This modification has been accomplished by scheduling less fuelto the ignitors during the starting cycle. Slower acceleration is the resultof this fuel scheduling. For TAFCU equipped APU’s, it is normal forAPU starting time to range from 40-70 seconds (cold aircraft: 135 sec-onds).

• Fuel for the gear-driven fuel control unit (FCU) is available from theNo.1 Fuel Tank. The fuel is automatically heated, if required, to preventicing. The fuel solenoid valve opens when oil pressure is sufficient toinitiate ignition.

APU Limitations

1. Maximum EGT. 7600C2. Maximum continuous. 7100C3. Maximum altitude pneumatic use. 17,000’4. Maximum altitude pneumatic/electrical. 10,000’5. Maximum altitude electrical load. 35,000’6. APU generator limit: ground: 125 amps

flight: 111 amps

APU Notes

• The APU is a gas turbine consisting of a two-stage centrifugal compres-sor directly coupled to a single-stage radial inflow turbine. The turbineshaft is geared to the accessory drive section and provides power fordriving the engine accessories and the generator. Electrical power fromthe airplane battery and fuel from the No.1 tank are used to start andoperate the APU. The APU supplies bleed air for engine starting and airconditioning. The APU also provides an auxiliary AC power source froma self-contained AC Electrical Generator.

• The standard APU on B737 aircraft was designed by Airesearch, desig-nated the GTCP 85-129E. The unit weights 313 pounds and is locatedin the tail section of the aircraft. Maximum fuel burn is rated at 340 PPHwith maximum loads on a standard day. The rated speed (at sea level -



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steady state) is 41,000 RPM, with an output of 6000 RPM at the gen-erator drive. Maximum EGT for starting or transient conditions is 7600C,with 7100C as maximum continuous. Maximum oil temperature is1240C. Automatic internal control units monitor EGT and will restrictpneumatic loads to favor electrical loads during periods of high demand.

• (Oral Topic) The battery switch must be ON during normal APU opera-tions. Positioning the battery switch to OFF, while the airplane is on theground, will automatically shut down the APU. During flight operations,the selection of the battery switch to OFF, does not shutdown the APU.When securing the aircraft, the APU must be completely secured for atleast 20 seconds to allow for the closure of the APU air inlet door priorto the selection of the battery switch to the OFF position. Reason forthis, the APU must be secured since fire detection is removed by theclosure of the battery switch.

• Some airlines require two minutes of normal APU operation prior tousing it as a pneumatic source. The manufacturer recommends at leastone minute prior to use. The same two minute requirement of normalAPU operation with no pneumatic load prior to shutdown is desired.The manufacture recommends at least one minute of APU operationwith no pneumatic load prior to shutdown.

• The APU generator is rated at 50 KVA inflight and 60 KVA while on theground.

• (Oral Topic) The start sequence is partially controlled by the "multi-speed" or the "speed switch". This internal switch is used to sequencevarious APU events such as :1. Starter cutoff at 50% RPM.2. Load circuit arming at 95% RPM.3. Fuel solenoid valve power shutoff at 110% rpm.

• (Oral Topic) Automatic APU shutdown protection is provided for thefollowing items:1. Overspeed (pops CB in the E/E compartment).2. Low oil pressure.3. High oil temperature.4. APU fire.5. Hung start (90 second timed).



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• (Oral Topic) APU bleed valve located on the air conditioning controlpanel, must be closed when:1. Ground air is connected and the isolation valve is opened.2. L.H. engine bleed valve is opened.3. Isolation Valve, R.H. engine bleed valves are opened.4. APU bleed valve may be opened during engine start, but avoid

engine power above idle.

• C.A.A. certified aircraft (Canadian or British etc.), are equipped with aDC power APU boost pump that provides fuel during starting. With theactivation of the APU speed switch (95%), the pump is deactivated.

APU Winter Operations

• During winter operations, special considerations must be reviewed. In-let area icing can be detected by higher than normal EGT indicationsand reductions in duct pressure. Operations in extreme cold conditionsmay require the use of both packs.

• Special precautions must be observed with the use of de-icing fluidsaround the APU inlet area. Care must be taken to prevent this fluidfrom entering the APU, therefore, the APU should be shutdown whende-icing that portion of the aircraft. Ensure the APU door is free ofimpacted snow and ice prior to APU operations.

APU Fire

• An APU fire is recognized by the fire warning bell ringing and the APUfire warning light illumination. The APU fire warning handle should bepulled and rotated. To manually unlock the APU fire handle, press theoverride and pull. By pulling the APU fire handle, this action performsbackup protection to the APU automatic shutdown feature. The follow-ing items occur:1. Closes the APU bleed air valve.2. Closes the APU inlet door.3. Trips the generator field.4. Arms the fire extinguisher system.5. Closes the fuel valve.



B737-200

APU Fire Detection

• The Kidde sensor fire detection loop is a single loop that signals a firewarning due to a general overheat or localized fire condition. The firedetector sensor-loops are mounted on the bottom of the APU engine (4000Fsensor), one on the outside leading edge of the exhaust pipe (7500F sen-sor), and one on the outside of the exhaust heat shield (3600F sensor).

• (Oral Topic) The power supply for the fire detection system (circuit andlight) is 28V DC from the battery bus. A fire short circuit discriminator isinstalled to prevent shorts from causing false fire warnings. Testing for firesends an artificial electronic signal to the fire warning detector. The detec-tor is not actually heated. A short in the fire circuit is indicated by theillumination of the APU DET INOP light. This light is located on the aftelectronic control panel.

APU Ground Control Panel

• The APU ground control panel is located on the aft bulkhead of the rightmain wheel well. It provides visual and aural fire warning and extinguisheroperation/control from outside of the aircraft. When a fire is detected, thehorn and light will operate alternately and the APU will automatically shut-down. Selecting the horn cutout switch will stop the horn from soundingand the red light will remain illuminated, visually warning of the fire condi-tion. The selection of the fire handle will arm the extinguisher system andwill shutdown the APU should the automatic shutdown features fail to se-cure the APU.



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Circuit Breaker Light Control

• The circuit breaker light controller when rotated clockwise regulates theintensity of the P-6 and P-18 circuit breaker panels illumination. Indi-vidual lights may be inoperative for dispatch provided the remaininglights are sufficient to illuminate all instruments and panel switches.Variable intensity control switches may also be inoperative providedthat this feature is unnecessary. Refer to your MEL.

Panel Light Control

• The panel light controller when rotated clockwise regulates the intensityof the forward and aft overhead panel lights. Individual lights may beinoperative for dispatch provided the remaining lights are sufficient toilluminate all instruments and panel switches. Variable intensity controlswitches may also be inoperative provided that this feature is unneces-sary. Refer to your MEL.



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Equipment Cooling Switch

• The equipment cooling switch is a two position (NORMAL-ALTERNATE)switch. When selected to the NORMAL position, the normal cooling fanis activated. Power source for “normal fan” is provided by the groundservice bus. Selecting the switch to the ALTERNATE position, thealternate cooling fan is activated. The power source for “alternate fan”is from the No.2 main bus.



B737-200

Equipment Cooling Off Light

• The illumination of the amber equipment cooling OFF light indicates noairflow from the respective cooling fan. A no airflow detector in theducting just forward of the fans consists of a heating unit located belowa thermal switch. Loss of airflow causes a thermal switch to close,illuminating the equipment cooling OFF light. Non-normal procedurescalls for the selection of the alternate fan which should restore airflow tothe equipment.

• Ground operations and operations with low differential (less than 2.5psi) causes the air to be dumped overboard through a flow control valvein the bottom of the aircraft. In-flight operations with high differentialpressures (greater than 2.5 psi), causes the warm air to be routedforward to the forward cargo compartment. This warm air insures ad-equate heating at the higher altitudes.

• The flow control valve used in the equipment cooling system, operatesunder the influence of differential pressure and “aerodynamic” forceswhich either opens or closes the valve, thus allowing the proper direc-tion of the air from the equipment cooling system.



B737-200

Emergency Exit Lights

• The guarded emergency exit light switch is a three position (OFF/ARMED/ON) switch. When selected to the OFF position, prevents theactivation of the emergency lights system when the airplane electricalpower fails or is turned off. Selected to the ARMED position, theemergency lights system activates all interior and exterior emergencylights automatically should DC power fail or if AC power has beenselected off. Placing the emergency exit switch to the ON position,manually illuminates all emergency lights.

• The emergency exit switch when selected to the ON position will illumi-nate all emergency lights. The passenger cabin emergency exit lights(flight attendants switch) may override the flight deck controls and illumi-nate the emergency lights.

• (Oral Topic) The emergency lights, located throughout the passengercabin, are powered by individual nicad batteries with a charging, moni-toring and voltage regular circuit. If electrical power to the 28V DC busNo.1 fails or if AC power has been turned off, the emergency exit lightswill illuminate automatically.



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• The emergency battery packs contain six nicad batteries and a chargingpackage to maintain the batteries charge. The system is designed toprovide illumination for approximately 15-20 minutes and has incorpo-rated within the system a device that prevents total battery discharge ifthe system is inadvertently left on.

• The charging of the emergency lights batteries occurs regardless of theposition of the emergency light switch, either in the OFF or ARMEDpositions. The battery pack electronics unit provides 350 milliamperesbattery charging current when the DC bus No.1 is powered.

Emergency Exit Lights Not Armed Light

• The illumination of the amber emergency exit lights NOT ARMED lightindicates the emergency exit lights switch is not in the ARMED position.

Passenger Cabin Emergency Exit Lights

• The passenger cabin emergency exit light switch, located on the aftflight attendant’s station, is a two position switch (NORMAL-ON). Whenselected to the ON position, illuminates all interior and exterior emer-gency lights.

• (Oral Topic) The cabin emergency exit light switch can be used tooverride the flight deck control of the emergency lighting system andilluminate the emergency lights. This can be accomplished wheneverthe flight deck controls have been selected to the ON position. Thedesign concept is to provide the means of bypassing the cockpit controlof the emergency lights in the case of an actual emergency or with thefailure of the flight deck emergency light controller.



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No Smoking Passenger Warning Switch (as installed)

• The no smoking passenger warning switch controls the illumination ofthe no smoking sign portion of the passenger notice system. The nosmoking switch is labeled OFF, AUTO, and ON. Selecting the respec-tive switch to OFF, extinguishes the cabin no smoking sign. Selectingthe AUTO position, will automatically illuminate the passenger informa-tion sign when the landing gear is extended. When the landing gear isretracted, the light will be extinguished. Selecting the control switch toON, manually controls the illumination of the warning lights. A lowchime sound has been incorporated within the system to sound anytimea change has occurred, either automatically or manually.

• The no smoking sign must be readily legible to each passenger seat,cabin attendant seat, or lavatory. If this is not possible, then the seatmust be blocked and placard with DO NOT OCCUPY. The sign may beinoperative for dispatch, provided the PA system operates normally andan acceptable procedure is used to inform the passengers of thesepassenger requirements. Refer to your MEL for detailed proceduresconcerning the appropriate required announcements.



B737-200

Fasten Belt Passenger Warning Switch (as installed)

• The fasten belt passenger warning switch is located on the overheadcontrol panel. The dual three-position switch controls the illumination ofthe fasten seat belt sign portion of the passenger notice system. Thefasten belt switch is labeled OFF, AUTO, and ON. Selecting the re-spective switch to OFF, extinguishes the cabin fasten belt sign. Select-ing the AUTO position, will automatically illuminate the passenger infor-mation sign when the landing gear is extended. When the landing gearis retracted, the light will be extinguished. Selecting the control switchto ON, manually controls the illumination of the warning lights. A lowchime sound has been incorporated within the system to sound anytimea change has occurred, either automatically or manually.

• The fasten belt sign must be readily legible to each passenger seat,cabin attendant seat, or lavatory. If this is not possible, then the seatmust be blocked and placard with DO NOT OCCUPY. The sign may beinoperative for dispatch, provided the PA system operates normally andan acceptable procedure is used to inform the passengers of thesepassenger requirements. Refer to your MEL for detailed proceduresconcerning the appropriate required announcements.

Attendant Call Switch

• The attendant call switch is located on the overhead control panel.Selecting the attendant call switch, sounds a two-tone chime in thepassenger cabin and illuminates both pink overhead annunciator mastercall lights. The annunciator lights will stay illuminated until the reset hasbeen made by the cabin attendant.

Ground Call Switch

• Selecting the ground call, sounds an alert horn in the nose wheel well.The switch will continuously sound until the press-switch is released.

Cockpit Call Light

• The cockpit call light is located on the overhead control panel. Whenilluminated, indicates the flight deck is being called by the flight atten-dants or by the ground crew. The cockpit call light remains illuminateduntil the captain call or pilot call switch is released.



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Rain Repellent Switch

• Some airlines have deactivatedthis system for their fleets. Thefollowing system description hasbeen added since many examin-ers will require knowledge of thesystem operation for type ratings.

• (Oral Topic) The rain repellentswitch is located on the overheadcontrol panel. When the respec-tive press-select switch is momen-tarily activated, a time released(.17 seconds) application of 5cc ofrepellent is applied to the No.1windows.

• (Oral Topic) The rain repellentused in this system is a BoeingAircraft patented fluid called RainBoe, Type III. The container is apressurized can containing 550ccof fluid when full. The power sup-ply for the solenoid valve opera-tion of the rain repellent system is28V DC power. The left switch ispowered by the DC bus No.1, andthe right switch is powered by DC bus No.2. The repellent fluid ispackaged in one pressurized disposable container which is replacedwhen empty. This bottle is located on the P-18 panel aft of theCaptain’s seat. There is a level line marked on the receptacle housingadjacent to the sight gauge.

• For preflight considerations, verify the float is above the reference lineand the shutoff valve is in the vertical position. When the sight gaugefloat is at or below the level line, the bottle should be replaced. With thefloat at the replacement level (mark), about 10 applications of fluidremains.



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• (Oral Topic) Rain repellent is used in conjunction with the windshieldwipers to improve visibility during heavy rain. The repellent fluid isrecommended only in moderate or heavy rain. Rain repellent should notbe used on dry windshields because the undiluted solution will restrictvisibility and the use of the wipers will only increase the smearing.Caution should be exercised to avoid the application of the rain repellentfluid to a dry windshield. The fluid will dry to a hard film. To removethis film, mild detergent, water, and a soft cloth may be used. In case ofheavy, built-up layers of film, additional rain repellent may be applied toact as a solvent. This new fluid must then be washed off with waterimmediately.

• The rain repellent system may be inoperative for dispatch provided thesystem is not required and procedures do not require its use. Refer toyour MEL.

Windshield Wiper Selector

• The windshield wiper selector is located on the overhead center panel.The selector is a four-position rotary controller that controls the wind-shield wipers on the No.1 windows. Each wiper is operated by a sepa-rate system to ensure that clear vision through one of the windows ismaintained in the event of a system failure. Although two independentsystem are installed, both wiper systems are electrically operated andcontrolled by a common controller switch.



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• Selecting the switch momentarily to the PARK position will help stow thewiper blades to the lower edge of the window. Moving the selector fromthe OFF position to LOW, will initiate the operation of the wipers at 130strokes per minute. The HIGH position will operate the wipers at 160strokes per minute. Caution should be exercised to not operate thewindshield wipers on a dry window.

• The windshield wipers may be inoperative for dispatch provided theaircraft is not operated in precipitation within 5 nautical miles of yourarrival and departure airports. Refer to your MEL.



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Window Heat ON Light

• The illumination of the green window heat light indicates the windowheat controller is applying heat to the respective window. The windowson each side of the cockpit are provided with electrical heating for anti-icing and defogging. Heating of the windows improves the impactstrength of the windows for bird-strike protection. Air from the air condi-tioning system can be used to defog the No.1 cockpit windows.

• Electrical power from the respective AC generator bus provides powerto the four window heat control units. These units are located in the E/Ecompartment. Flight compartment windows No.1, No.2, No.4, and No.5(left and right sides) are heated. Window No.3 is warmed by vented airbetween the two acrylic panes of the window. The No.1 window canalso be heated from the air conditioning system by the windshield aircontrol system.

• (Oral Topic) The windows are paired to a power source to insureadequate visibility with the loss of one generator. Windows L1, L4, L5,and R2 are powered by the No.1 generator bus. Windows R1, R4, R5,and L2 are powered by the No.2 generator bus.



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Window Overheat Light

• The illumination of the amber window overheat light indicates electricalpower to the respective window has been removed. The window over-heat light will also illuminate if an overheat condition occurs. As windowtemperatures reach 620C (1450F), the overheat sensors de-energizespower to the window and illuminates the amber light. Resetting thewindow overheat system is accomplished by momentarily positioningthe window heat switch to OFF. With the illumination of this light, thefollowing items will occur: the respective window heat light will be extin-guished, the Master Caution Lights and the Anti-ice System annunciatorlight will illuminate. The window overheat light may also illuminate if theelectrical power to the window has been interrupted.

• The No.1 or No.2 window heater may be inoperative for dispatch pro-vided the aircraft is not operated in known icing conditions, the wind-shield defog system operates normally, and the airspeed is limited to250 kts below 10,000’ MSL. Refer to your MEL.

Window Heat Switch

• The window heat switch is a two-position OFF/ON controller. Selectingthe switch to ON, signals the window heat controller to apply heat (fiveor more watts of power) to the associated window. This power ismodulated to maintain the window temperature at 430C (1100F). Theswitch labeled FWD, controls power to the No.1 windows. The switchlabeled SIDE, controls power to No.2, No.4 and the No.5 windows(eyebrow windows). The eyebrow windows are thermostatically con-trolled and the main windows are constant temperature powered.

• Standard airline operating procedures calls for the window heat switchesselected to ON at least 10 minutes prior to takeoff. This limitation isalso listed in many airline MEL’s.

Window Heat Test Switch

• The window heat test procedure is designed to test the overheat andpower functions of the window heat system. The OVHT position whenselected, simulates an overheat condition which causes the amber over-heat lights, the Master Caution Lights and Anti-Ice System annunciatorlights to illuminate. Reset of the system can be accomplished by mo



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mentarily positioning the window heat switch to OFF then ON. Theillumination of the overheat lights indicates that the overheat circuits areoperating properly. The overheat light(s) may extinguish immediately orremain illuminated for as long as 70 seconds.

• The PWR TEST position provides a confidence test when any of thegreen ON lights fail to illuminate after the selection of the window heatswitch to the ON position. The test mode will force the window control-ler to full power, bypassing the normal temperature control. If any greenlight remains extinguished during the power test, window heat protectionmay be lost. Do not hold the PWR TEST switch on for extendedperiods of time, overheating of the windows will occur. It is permissibleto perform the test with the other green lights illuminated. Overheatprotection is available during the power test mode. For inoperativewindow heat procedures, refer to the MEL notation in the previoussection.

• Non-normal procedures for a window overheat condition requires theselection of the affected window heat switch to the OFF position. Allow2-5 minutes before selecting the window heat switch to ON again.Should the light illuminate again, select switch to OFF and limit airspeedto 250 kts maximum below 10,000. For windshield defogging, positionthe windshield air controls to the PULL position. Allow extra time fordefogging in humid climates (30-45 minutes application prior to touch-down).



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Windows

• The aircraft windows are constructed of laminates consisting of glass,vinyl and acrylic. Windows No.1, No.2, No.4 and No.5 are electricallyheated by power passing through a conductive coating. On windowsNo.1 and No.2, a conducting coating is located near the outer surface.This provides heating to prevent ice buildup on the respective windows.On windows No.4 and No.5, the conductive coating is located near theinside surface of the window. This helps prevent window fogging.

• Arcing-delaminated-shattered-cracked window condition is recognizedby arching, substantial delamination, shattering or cracking of any cock-pit window. Window 4 is the only window having a middle glass pane.A failed middle pane appears shattered and transparency is virtuallylost. Non-normal procedures calls for the window heat switch to beselected OFF and limit maximum airspeed to 250 kts. If Window No.1,No.2, No.4, or No.5 is affected, set cabin altitude selector to 10,000’ andset the pressurization mode selector to STBY. Reduce pressure differ-ential by limiting flight altitude. Follow recommended differential pres-sure placards located in your operation manuals or emergency check-lists.

Pitot Static Heat Switch

• The pitot static heat switch is a two-position control switch. Selectingthe switch to the ON position, applies power to heat the respectivesystem. Electric power is supplied to the resistance type heaters insideall probes and vanes. Selecting to the OFF position, removes all electri-cal power from the pitot. The power sources for the pitot system are115V AC Bus No.1, bus No.2, the transfer bus No.1, transfer bus No.2for the switches and 28V DC bus No.1, and bus No.2 & battery bus forthe indicating lights. Alternate static ports are not heated.

• The respective pilot’s or copilot’s pitot/static heater probe (upperprobes), may be inoperative for dispatch providing the aircraft is onlyoperated during day VMC conditions and if the aircraft is not operated invisible or known/forecasted icing conditions. The appropriate staticsource selector switch should be selected to the alternate static source.Refer to your MEL.



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• The respective No.1/No.2 auxiliary pitot-static probe heater may be in-operative for dispatch providing the corresponding auxiliary heater oper-ates normally and the aircraft is not operated in known or forecast icingconditions. Refer to your MEL. The No.1 auxiliary pitot-static heaterprobe is the right lower probe. The No.2 auxiliary pitot-static heaterprobe is the left lower probe.

Probe Heater Lights

• There are two types of probe heater lights (green and amber) installedon the B737-200 aircraft. Aircraft installed with green lights, the illumi-nation of these lights indicates the electrical power is supplying theheating unit and the heating unit is operating normally.

• Aircraft installed with amber probe heater lights have a different lightdescription. The illumination of this amber light indicates the electricalpower is not supplying the heating unit and the heating unit is notoperating normally.

Heater OFF Light

• The amber heater OFF light is not installed on the control panels wherethe probe heater lights are amber in color. The illumination of theheater OFF light indicates one and/or both pitot static heat switches arein the OFF position. The illumination also indicates heat is not beingapplied to the Captain’s or First Officer’s primary or auxiliary pitot staticprobes. Respective green lights are extinguished. The Master Cautionlights and the Anti-Ice Annunciator light will illuminate upon detection ofthe heater OFF light illumination.



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Wing Anti-Ice Valve Open Light

• The blue wing anti-ice valve open light has two levels of light intensity.When illuminated a bright blue, indicates the corresponding wing anti-icecontrol valve is in transit, or if not in transit, the position of the valve is indisagreement to the position of the wing anti-ice switch or the electricalconnector is disconnected from the valve. The dim illumination of thelight indicates the corresponding wing anti-ice control valve is open(switch in the ON position). With the light extinguished, the correspond-ing wing anti-ice control valve is closed (switch is in the OFF position).The wing anti-ice valve open light may be inoperative for dispatch pro-vided the valve operates normally before operating into known or fore-casted icing conditions. Refer to your MEL.

Wing Anti-Ice Switch (three-position switch)

• (Oral Topic) The wing anti-ice switch is a three position, spring-loadedswitch (GRD TEST-OFF-ON). When selected to the OFF position, thewing anti-ice control valves will close and the wing anti-ice valve openlights will be extinguished. Selecting the switch to ON, the controlvalves will open. The wing anti-ice valve open lights will illuminatebright and then a dim intensity when the valves are fully opened. Whenselected to the OFF position, the wing anti-ice control valves will close



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and the wing anti-ice valve open lights will be extinguished. Selectingthe switch to ON, the control valves will open. The wing anti-ice valveopen lights will illuminate bright and then a dim intensity when thevalves are fully opened.

• (Oral Topic) Selecting the wing anti-ice switch to GRD TEST, openswing anti-ice control valves, only if the aircraft is on the ground. If ducttemperature exceeds 930C during the ground test, the wing anti-icecontrol valves will close. The air-ground sensor prevents the wing anti-ice control valves from operating on the ground except during groundtest.

• The wing anti-ice valve is located in each wing leading edge outboard ofthe strut. The power source to operate the wing anti-ice valves is 115VAC (motors) from the No.1 transfer bus. Control for the valve operationis 28V DC from the battery bus.

• The wing anti-ice valves may be inoperative for dispatch provided thevalve is manually closed for engine start, the respective manifold isdepressurized when outside air temperature is above 500F, the respec-tive engine bleed thrust limits are followed, and pressurization/air condi-tioning requirements are followed when one or both manifolds are de-pressurized. No wing anti-ice decrements are published for takeoff,therefore the engine bleed for the affected manifold must be OFF fortakeoff. Refer to your MEL.

Wing Anti-Ice Switch (two-position switch)

• The wing anti-ice switch is a two position switch (OFF-ON). Whenselected to the OFF position, the wing anti-ice control valves will closeand the wing anti-ice valve open lights will be extinguished. Selectingthe switch to ON, the control valves will open. The wing anti-ice valveopen lights will illuminate bright and then a dim intensity when thevalves are fully opened.

• (Oral Topic) During the takeoff roll, with the switch ON, the air-groundmode switch will cause the switch to trip OFF at lift-off. This function isautomatic, and requires no crew action.



B737-200

• During ground operations, the wing anti-ice switch may be selectedON, providing the thrust on both engines is below the takeoff warningposition, and the temperature inside both distribution ducts is below thethermal protection temperature limit of (1250C). The thrust setting andduct sensor temperature logic are both bypassed during inflight opera-tions. The wing anti-ice control valves are motor-operated using ACpower. Criteria for ground use is the same as for the use of engine anti-ice, “BOTH ON” at the same time.

• (Oral Topic) The wing anti-ice system provides protection for the lead-ing edge slats by using bleed air ducted from the main pneumatic mani-fold. Wing anti-ice protection does not include the leading edge flaps.Bleed air flows through the wing distribution duct in the leading edge,through a telescoping duct to each slat, and then exhausted overboard.This protection is effective with the slats in any setting. Prolongedoperation in icing conditions with the leading edge and trailing edgeflaps extended is not recommended. As a recommended procedure,limit trailing edge retraction to the Flaps 15 position. After shutdown,perform a postflight inspection for ice accumulation during the landingroll.



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• Failure of the wing anti-ice valve in the OPEN position, with tempera-tures above 100C TAT, requires the isolation valve to be closed. Thisaction prevents the isolation valve from opening when non-normal pro-cedures directs the selection of the affected engine bleed air switch tothe OFF position. By selecting the bleed air switch to OFF, unwantedwing anti-ice operation will be prevented.

• The only one wing anti-ice valve may be inoperative for dispatch pro-vided the respective manifold is depressurized when outside air tem-perature is above 500F, the respective engine bleed thrust limits arefollowed, and the pressurization/air-conditioning requirements are fol-lowed when one or both manifolds are depressurized. The reason foronly one valve limitation is due to the location of the wing anti-ice valveand to the requirement to provide sufficient air for engine starting withthe valve failed in the open position. Refer to your MEL.

Cowl Valve Open Light

• The blue cowl valve light has three modes of illumination. When illumi-nated a bright blue, indicates the respective control valve is in transit, orif not in transit, the position of the valve is in disagreement with theposition of the respective engine anti-ice switch. The illumination of thelight to that of a dim intensity indicates the respective control valve isopen (switch is in the ON position). Hot 13th. stage bleed air andambient air from the engine inlet is mixed and applied to the nose cowllip areas. Temperature is automatically regulated and controlled. Withthe cowl valve open light extinguished, indicates the respective controlvalve is closed (switch is on the OFF position).

• One nose cowl anti-ice valve may be inoperative in the CLOSED posi-tion for dispatch provided all remaining anti-ice valves operate normallyand the aircraft is not operated in known or forecasted icing conditions.Refer to your MEL.

• One nose cowl anti-ice valve may be inoperative in the OPEN positionfor dispatch provided various thrust limits are reduced by .03 EPR,enroute climb weights are reduced, all remaining valves operate nor-mally, operating temperatures for cowl valves are limited to 500F maxi-mum ambient temperatures, and further limitations as applied for -15and -17 equipped aircraft are followed as outlined in the MEL. Refer toyour MEL.



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• During flight operations, failure of the engine cowl valve may be indi-cated by the bright illumination of the cowl valve open light. If total airtemperature is above 100C , limit thrust of the affected engine to 80%N1 if possible. With the engine cowl valve failed in the closed position,avoid icing conditions if possible.

Left or Right Engine Anti-Ice Valve Light

• The respective left or right blue engine anti-ice valve light has threemodes of illumination. A bright blue illumination, indicates the respec-tive control valve is in transit, or if not in transit, the position of the valveis in disagreement with the position of the associated engine anti-iceswitch. The illumination of the light to that of a dim intensity indicatesrespective control valve is opened. Eighth stage bleed air is applied tothe engine inlet guide vanes, nose dome areas, and to the Pt2 probe.With either valve inoperative, adequate anti-icing capabilities are stillavailable through the opposite valve. With the cowl valve open lightextinguished, indicates the respective control valve is closed (switch isin the OFF position).

• One of the four engine anti-ice valves may be inoperative in theCLOSED position for dispatch provided all remaining anti-ice valvesoperate normally and the aircraft is not operated in known or forecastedicing conditions. One of the four engine anti-ice valves may be inopera



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tive in the OPEN position for dispatch provided various thrust limits arereduced by .03 EPR, enroute climb weights are reduced, all remainingvalves operate normally, and further limitations as applied for -15 and -17 equipped aircraft are followed as outlined in the MEL.

Engine Anti-Ice Switch

• The selection of the engine anti-ice switch to the ON position opens therespective engine anti-ice valve and illuminates the cowl valve openlight. Observe a decrease on the engine’s EPR gauge when selectingthis switch. Movement of the switch to the OFF position, closes thevalve and the cowl valve open light extinguishes. The engine anti-icesystem may be operated on the ground and infight whenever icingconditions exist, except during climb and cruise when the temperature isbelow -400C SAT.

• Indications of ice accumulation on the fan blades at low thrust settingsmay be seen as increased engine vibration levels. Ice may be shed athigh RPM settings. To clear an engine, advance the thrust lever to 70%N1. If after one minute, the vibrations continue, consider possible en-gine shutdown procedures. Many airline manuals have procedures tofollow with reference to fan blade icing and shedding techniques forengine ice vibrations. Refer to yours.



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• (Oral Topic) During icing conditions, the following aircraft limitationsapply: a minimum of 40% N1 RPM with TAT temperatures between 00

and 100C; 55% N1 RPM with TAT temperatures below 00C, and withTAT temperatures -6.50C, maintain 70% N1 RPM or higher. Do notoperate engine anti-ice when the TAT is above 100C. For operations onthe ground while in icing conditions, it is recommended that enginethrust be increased to 80% N1 for 15 seconds for every 10 minutes.

• (Oral Topic) Occasionally, examiners may inquire about the relation ofthe engine bleed valve and the engine anti-ice system. Refer to therespective engine pneumatic schematic, you will note that the engineanti-ice valves are located upstream of the engine’s bleed valves.

Icing Conditions

• Icing conditions exist when OAT is 100C or below during ground opera-tions, takeoff, initial climb or go-around, or; TAT temperature is 100C, orbelow inflight and visible moisture in any form is present: such as cloudsor fog with visibility less than one mile, rain, snow, sleet, ice crystals,etc., or when standing water, ice or snow is present on ramps, taxiwaysor runways.



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Ground Interconnect Switch

• The selection of the ground interconnect switch to the OPEN positionallows System B pressure to be connected to System A for groundfunctional checks. The ground interconnect valve will only open if theparking brake is set and the airplane is the ground mode.

• (Oral Topic) A commonly asked oral topic concerns the position of theparking brake lever and the ground interconnect switch. With theground interconnect switch in the OPEN position, releasing the parkingbrake lever will automatically close the ground interconnect valve.

• The ground interconnect valve may be inoperative for dispatch providingthe valve is in the closed position. The electrical power to the valvemust be removed and the override lever must be placed in the closedposition. These procedures are normally accomplished by your mainte-nance personnel. Refer to your MEL.



B737-200

Hydraulic Overheat Light - Electric Pump

• The illumination of the amber overheat light indicates the hydraulicpump and/or fluid used to cool and lubricate the respective electricmotor driven pump has overheated. Refer to the schematic section ofthis manual, you will notice two sensors are used for this light . Eithersensor will illuminate the light. Non-normal procedures calls for theselection of the affected pump switch to be selected OFF. Normalsystem pressure can be maintained with single pump operation. Lightsmay be inoperative for dispatch. Refer to your MEL.

Hydraulic Pump Low Pressure Light

• The illumination of the hydraulic pump low pressure light indicates theoutput pressure of the respective pump is low. The Master Cautionlights and the Hydraulic Annunciator light will also illuminate for lowhydraulic pressure. The engine driven low pressure warning circuit iswired to the engine fire handle. When either fire handle is pulled, thefluid flow of the associated engine driven pump is terminated and thelow pressure light is deactivated. Extended operation with no fluid tothe engine driven pump will cause damage to the pump because of lackof fluid cooling to the internal parts of the windmilling action of the pump.

• (Oral Topic) A commonly asked oral topic is the location of the lowpressure sensors and the possibilities of a failed check valve. The lowpressure light sensor is located downstream of the pump. When select-ing a pump ON, if you notice both low pressure lights are extinguished,this may indicate a failed check valve on the opposite side of selectedpump.

• The hydraulic low pressure lights may be inoperative for dispatch pro-viding the output of the respective pump is checked before each depar-ture. There is a procedure for the flight crew to follow to accomplish thischeck. After starting the engine, ensure the respective “A” systemhydraulic pump is activated. Verify that the system is pressurized. For“B” system pressure lights, the engine does not have to be started.Select the respective switch to the ON position and verify that thesystem is pressurized. Refer to your MEL.



B737-200

Engine Driven Pump Switch

• (Oral Topic) The selection of the respective engine driven hydraulicpump switch to the ON position de-energizes the blocking valve in pumpto allow pump pressure to enter the selected system. The pump switchshould remain ON at the end of the flight to prolong the life of thesolenoid. The selection of the switch to the OFF position energizes theblocking valve to block the hydraulic pump output.

• The depressurization function of the engine driven hydraulic pump maybe inoperative on both pumps. Refer to your MEL.

Electric Motor Pump Switch

• The selection of the B system electric pump switch to the ON positionprovides power to the respective electric motor driven pump. Minimumfuel for ground operations of “B” pumps is 1676 pounds in the No.2 wingfuel tank. The hydraulic system B heat exchanger is located in the No.2wing fuel tank and is larger than system A exchanger. The heat ex-changer must be covered with a minimum of 1676 pounds of fuel fornormal operation.



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Hydraulic System

• Hydraulic power is provided by three independent sources; System A,System B and the Standby System. System A pressure is powered bytwo engine driven pumps located on each engine. System B pressure isprovided by two electrically driven hydraulic pumps. The standby sys-tem pressure is provided by the standby electric driven hydraulic pump.Nominal operating pressure for each hydraulic system is 3000 psi.Each hydraulic system has a fluid reservoir located in the main wheelwell area. The reservoirs are pressurized by 13th. stage bleed airwhich is directed into the system A reservoir. Fluid balance lines inter-connect all reservoirs. This provides a constant pressure to ensurepositive fluid supply.



B737-200

• System A hydraulic fluid flows from the reservoir through shutoff valvescontrolled by the engine fire warning switches. Pulling a fire switch willshut off the flow of fluid to the respective pump and deactivates thehydraulic pump low pressure light. Engine windmilling will cause inter-nal pump damage after a short period of time. Do not confuse thisaction with the selection of the "A" pump to the OFF position. Pumpfiltering and cooling occurs when the pump has been selected to thisposition. This provides cooling and lubrication for the internal parts ofthe hydraulic pump. The system A heat exchanger is located in theNo.1 main fuel tank and must be covered with fuel for the operation ofthe pumps. Both system A pumps are required for dispatch, only thedepressurization function may be inoperative. Refer to your MEL.

• System A hydraulic components consist of the following items:a. Inboard brakes.b. Inboard flight spoilers.c. Ground spoilers.d. Ailerons.e. Elevators.f. Rudder.g. Trailing edge flaps.h. Leading edge devices.i. Landing gear.j. Nose wheel steering.k. Thrust reversers .l. Inboard autobrakes (analog).



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• (Oral Topic) Hydraulic system B is connected to the system A reservoirand the standby reservoir by balance lines. The purpose of the balancelines is for servicing and pressurization of the hydraulic systems. Thehydraulic system B heat exchanger is located in the No.2 wing fuel tankand is larger than system A exchanger. The heat exchanger must becovered with a minimum of 1676 lbs of fuel for normal operation. Thetwo electric powered pumps have check valves that isolate each other.

• Only one of the two system B electric hydraulic pumps may be inopera-tive for dispatch providing the pressure indicator and the thrust revers-ers operates normally. Refer to your MEL.

• System B hydraulic components consist of the following items:a. Outboard brakes.b. Outboard flight spoilers.c. Ailerons.d. Elevators.e. Rudder.f. Yaw damper.g. Autopilot.h. Outboard autobrakes (analog).i. Inboard/Outboard autobrakes (digital).



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• The standby hydraulic pump has been provided as a backup system tosystem A and system B. The standby system is connected by a bal-ance line as described in the system B outline. Only one hydraulicpump powers the standby system. The standby system may be acti-vated manually by the selection of either flight control switch to STBY orby the selection of the alternate flaps master switch to ARM. Reviewthe flight control system description for further information pertaining tothe standby hydraulic system. The standby hydraulic pump must beoperational for dispatch. Refer to your MEL.

• The standby hydraulic components consist of the following items:a. Leading edge devices (extend function).b. Thrust reversers.c. Standby rudder.



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Hydraulic Malfunction: Loss of System A Pressure

• System A hydraulic malfunctions may be indicated by a loss of pressureor a loss of fluid. Crosscheck System A quantity. With quantities belowthe 1.85 level, a possible hydraulic leak within the system may exist.Loss of System A pressure is indicated by the System A pressure atzero and the illumination of the following lights:

a. Master caution lights.b. FLT CONT & HYD annunciator lights.c. System A low pressure lights.d. System A flight control low pressure light.e. FEEL DIFF PRESS light (flaps up).

• (Oral Topic) Non-normal procedures requires the selection of the Sys-tem A flight control switch to the STBY RUD position. This will activatethe standby pump, which will power the rudder power control unit. Thenext item on the checklist calls for the selection of the System A hydrau-lic pumps switch to the OFF position. With the System A hydraulicpumps selected OFF, you will lose the following items:

a. Ground spoilers.b. Inboard flight spoilers.c. Nose wheel steering.d. Autopilot A.

• With the loss of system A pressure, prepare for a Flaps 15 landing withVref+15. Give considerations for proper approach planning sincemanual gear and alternate flap extension procedures are required. Thethrust reversers will be operational using standby pressure. Inboardbrakes will have accumulator pressure only.



B737-200

Hydraulic Malfunction: Loss of System B Pressure

• System B hydraulic malfunctions may be indicated by a loss of pressureor a loss of fluid. The first indications of a System B hydraulic malfunc-tion may be a drop in System A quantity to the 1.85 gallon mark.Crosscheck System B low quantity light for illumination. With the illumi-nation of the low quantity light, the fluid in the reservoir has droppedbelow the half level (.65 gallons). Loss of System A pressure is indi-cated by the System A pressure at zero and the illumination of thefollowing lights:

a. Master caution lightsb. FLT CONT & HYD annunciator lights.c. System B low pressure lights.d. System B flight control low pressure light.e. FEEL DIFF PRESS light (flaps up).f. System B low quantity light.

• Non-normal procedures requires the selection of the System B flightcontrol switch to the STBY RUD position. This will activate the standbypump, which will power the rudder power control unit. The next item onthe checklist calls for the selection of the System B hydraulic pumpsswitch to the OFF position. With the System B hydraulic pumps se-lected OFF, you will loose the following items:

a. Outboard flight spoilers.b. Yaw damper.c. Autobrakes.d. Autopilot B.

• With the loss of system B pressure, there are no special approachpreparations or considerations for there are no abnormal flap or landinggear conditions to deal with. The thrust reversers will be operational.Outboard brakes will have accumulator pressure only. Autobakesshould be selected to the OFF position.



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Hydraulic Malfunction: Loss of Standby System

• Standby hydraulic system malfunctions may be indicated by a loss ofpressure or a loss of fluid. Crosscheck the standby low quantity light orthe standby low pressure lights for illumination. Loss of the standbysystem pressure is indicated by the illumination of the following lights:

a. Master caution lights.b. FLT CONT annunciator light.c. Standby low pressure light.d. Standby low quantity light.

• With the failure of hydraulic system A and the standby hydraulic system,the thrust reversers and the leading edge devices will be inoperative.Failure of system A, B, and the standby system will cause the loss ofrudder use. If the standby hydraulic system develops a leak, you willobserve system A quantity indicating 1.84 gallons and system B quan-tity should be full. The balance line from system B reservoir is attachedto the top portion of the reservoir, therefore, the leak will be indicated bythe low quantity warning light and the lower than normal quantities onthe system A quantity gauge.

Hydraulic Malfunction: Manual Reversion

• The loss of both hydraulic systems (A & B) is known as manual rever-sion. System A and B pressure will be at zero and will cause theillumination of the following lights:

a. Master caution lights.b. FLT CONT and HYD annunciator lights.c. System A & B low pressure lights.d. System A & B flight control low pressure lights.



B737-200

• Non-normal procedures requires the selection of the System A andSystem B flight control switches to the STBY RUD position. This willactivate the standby pump, which will power the rudder power controlunit. The next item on the checklist calls for the selection of the SystemA and System B hydraulic pump switches to the OFF position. Withboth System A and System B hydraulic pumps selected OFF, you willlose the following items:

a. Ground spoilers.b. Inboard and outboard flight spoilers.c. Nose wheel steering.d. Autopilots.e. Autobrakes.f. Yaw damper.

• With the loss of system A and B pressure, prepare for a Flaps 15landing with Vref+15. Give considerations for proper approach planningsince manual gear and alternate flap extension are required. The thrustreversers will be operational using standby pressure. Inboard and out-board brakes will have accumulator pressure only.

• With the loss of both hydraulic systems, the ailerons are controlledmanually. Bank angle should be limited to 200 degrees of bank. Flylarge landing patterns with a long straight-in approach. Keep thrustchanges to a minimum. Rapid thrust applications result in maximumnose up pitch forces. Fly a normal landing profile, do not make a flatapproach. On touchdown, thrust reverser operation will be slow; applysteady brake pressure, do not modulate the brakes. Do not attempt totaxi the aircraft, nose wheel steering will be inoperative and you arelimited by the capacity of the brake accumulators for stopping.



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Cockpit Voice Recorder (C.V.R.)

• There are several different types of cockpit voice recorders installed, allproviding the same basic functions. The cockpit voice recorder usesfour independent channels to record cockpit audio on a 30 minute con-tinuous-loop tape. Recordings older than 30 minutes are automaticallyerased. One channel records cockpit area conversations using the areamicrophone. The other channels record individual audio selector paneloutput (headset) audio and transmissions for the captain, first officer,and for the first jumpseat observer. The cockpit voice recorder systemmay be inoperative for dispatch providing repairs to the system arecompleted within three flight days and the Flight Data Recorder (FDR)operates normally. Refer to your MEL.

C.V.R. Area Microphone

• The cockpit voice recorder area microphone is designed to receivecockpit area conversations anytime 115V AC is applied to the aircraft.

C.V.R. Erase Switch

• The cockpit voice recorder erase switch is operative only when theaircraft is on the ground and the parking brake has been set. Selectingthe switch for more than 2-18 seconds (depending on type of recorderinstalled), all four channels are simultaneously erased.



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C.V.R. Test Switch

• The cockpit voice recorder test switch when selected for more than 5seconds, initiates an operational check of all four channels. Observethe monitor light or the monitor indicator (as installed) for proper opera-tion. For aircraft installed with the monitor indicator, observe two needledeflections of full scale movement into the green band for the 4 chan-nels. Partial deflection of the needle may indicate only a particularchannel is not functioning normally, contact your maintenance depart-ment. A tone may also be heard through a headset plugged into theheadset jack.

C.V.R. Monitor Light (as installed)

• The illumination of the white cockpit voice recorder monitor light indi-cates proper operation of the recorder. The light will illuminate twiceduring the operational check, thus verifying proper operation of thecockpit voice recorder.

C.V.R. Headset Jack

• The cockpit voice recorder headset jack may be used with a headset tomonitor tone transmission during the operational test of the cockpitvoice recorder. The headset jack may also be used to monitor playbackor voice audio. The headset jack may also be labeled as HEADSET600 OHMS.

C.V.R. Monitor Indicator

• The cockpit voice recorder monitor indicator has been installed with apointer needle. The deflection of this needle confirms actual recording.The erasure on all four channels may also be confirmed by observingthe indicator needle movement. During the operational test, the pointerrises into the green band after a one second delay, indicating the test isin progress.



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Cabin Altimeter-Differential Pressure Indicator

• The cabin altimeter-differential pressure indicator is located within thecabin control panel. The indicator has two scales of pressurizationreference. The inner scale indicates cabin altitude in feet. The outerscale indicates differential pressure between cabin and ambient in psi.There is also two types of controllers available, each displaying differentpsi/dif values. To determine the type of controller display installed inyour aircraft, observe the location of the yellow arc. The yellow arc willbe located either at 7.5 psi/dif or 7.8 psi/dif.

• The cabin differential pressure indicator may be inoperative for dispatchprovided the cabin altitude indicator is operating normally and the use ofthe cabin altitude-cabin differential chart is used to convert cabin altitudeto cabin differential pressure. Flight operations while unpressurizedneed not follow the above mentioned procedures. Refer to your MEL.

Altitude Horn Cutout Switch

• The altitude horn cutout switch is located within the cabin control panel.Selecting the switch-button, silences the intermittent cabin altitude warn-ing horn. The altitude warning horn sounds when the cabin reaches10,000’ altitude. The 28V DC power supply for the cabin altitude warn-ing circuit is provided from the battery bus.



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Cabin Rate of Climb Indicator

• The cabin rate of climb indicator is located within the cabin controlpanel. The indicator displays the cabin rate of climb or descent in feetper minute. The indicator may be inoperative for dispatch provided theautomatic and standby pressurization systems are operating normally.Refer to your MEL.

Cabin Altitude Warning System

• (Oral Topic) The cabin altitude warning system provides aural warningwhen the cabin altitude exceeds 10,000’ above sea level. When thecabin reaches 10,000’, a pressure switch closes, causing the warninghorn to sound. The pressure switch is located on the ceiling of thelower nose compartment. The warning horn is located inside the centercontrol stand. It is the same horn used for the flight control and landinggear warning systems.

Pressurization Limit Placard

• The pressurization limit placard is located within the cabin control panel.The placard displays the maximum cabin differential pressure limitationfor takeoff and landing (.125 psi).



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• The limitation of .125 psi refers to the technique of preventing a pres-sure bump during the takeoff. While taxing for takeoff, the FLT/GRDswitch is placed to FLT. This signals the cabin controller to partiallyclose the outflow valve to control the cabin pressure to .1 psid, thusholding the cabin pressure at approximately 189 feet below runway.The outflow valve will be near the closed position at rotation, therefore,preventing a pressure bump at rotation.

Auto Fail Light

• (Oral Topic) The illumination of the amber AUTO FAIL light provides avisual warning that a failure has occurred within the automatic pressur-ization portion of the pressurization control system. Pressurization con-trol will automatically transfer to the standby mode.

• A successful auto mode transfer is indicated by the illumination of thegreen STANDBY light. The STANDBY mode should be selected by thepressurization mode selector prior to takeoff so the transfer does notcause cabin pressure fluctuations. To extinguish the auto fail light,position the mode selector to STBY.



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• (Oral Topic) Any one of the following three items may cause the illumi-nation of the AUTO FAIL light:1. Loss of auto AC power. Failure of AC power supply to auto mode

circuits for more than 14.9 seconds. Low AC power supply mayalso cause this type of failure.

2. Excessive rate of cabin pressure change (+/-1800 feet per minute).3. High cabin altitude (13,875’).

Off Schedule Descent Light

• The illumination of the amber OFF SCHED DESCENT light indicatesthe aircraft has initiated a descent with the controller in the ascentschedule mode and the aircraft has not reached within 0.25 PSI of theselected flight altitude placed in the FLT ALT controller. The .25 psi isapproximately 1000' below set altitude. This commonly occurs when alower cruise altitude has been flown and the pressurization controllerhas been originally setup for a higher cruise altitude. During initialdescent, the OFF SCHED DESCENT light will illuminate. This warningcircuit is disarmed once the aircraft is within 0.25 PSI standard airpressure of FLT ALT set. The circuit does not become active againuntil after the aircraft lands.

• (Oral Topic) The purpose of OFF SCHED DESCENT mode is to warnthe crew if a descent is started before flight altitude is reached and thecontroller circuits are not ready to establish a descent schedule forlanding at destination airport.

Pressurization Standby Light

• (Oral Topic) The illumination of the green STANDBY light indicates thepressure controller is operating in the standby mode. With the controlleroperating in the standby mode, all control signals to the outflow valveare directed to the DC actuator. Should the standby mode fail, thegreen STANDBY light will be extinguished, thus indicating the standbycircuits are no longer controlling the outflow valve. No automatic trans-fer to another mode is available. The outflow valve will remain in thelast position. Crew action is to select another mode of operation (MANAC or MAN DC) with the mode selector switch.



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Pressurization Manual Light

• The illumination of the green MANUAL light indicates the pressure con-troller is operating in the manual mode. When the pressurization modeselector is manually selected to either the AC or DC positions, thecircuits are armed in the control panel, controller, and the outflow valve.

Flight Altitude Indicator

• The flight altitude indicator is located within the AUTO section of thepressurization control panel. The indicator displays the selected cruiseflight altitude. During preflight, the flight crew will set the flight planaltitude for which the aircraft will use during cruise in the FLT ALTcontroller. The pressurization controller will use this information to com-pute the ascent schedule of flight and for establishing the cabin cruisepressure for the remaining of the flight.

Flight Altitude Selector

• The “PUSH and ROTATE” flight altitude selector is located on theAUTO section of the pressurization control panel. The selector is usedto set the appropriate cruise flight altitude. The settings are from 0 to40,000’ in 100 foot increments



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Landing Altitude Indicator

• The landing altitude indicator is located on the AUTO section of thepressurization control panel. The indicator displays the selected desti-nation field elevation. During preflight, the flight crew will set the desti-nation field elevation for which the aircraft will use during the final phaseof the flight. The pressurization controller will use this information tocompute the descent schedule. This readout is capable of being setfrom -990’ (below S.L.) to 13,990’. The large diameter control sets tothe nearest 1000‘. The small diameter control sets to the nearest 10’.

Cabin Rate Selector

• The cabin rate selector is located within the STANDBY section of thepressurization control panel. The selector is used to set the desiredrate for cabin pressure change. The index mark is normally known asthe PIP mark. The PIP reference mark equals a cabin altitude rate ofchange of approximately 300 ft/min. The DECR mark reference equalsa cabin altitude rate of change of approximately 50 ft/min. The INCRreference mark equals a cabin altitude rate of change of approximately2000 ft/min.

Cabin Altitude Indicator

• The cabin altitude indicator is located within the STANDBY section ofthe pressurization control panel. The indicator is used to display thedesired cabin altitude during standby pressurization operations. Duringpreflight, the flight crew will set 200’ below runway elevation. Afterdeparture, the crew will insert the cabin altitude required for cruise. Achart located under the control panel has been provided to determineproper cabin altitude/flight altitude. Prior to descent, the crew will insertagain the altitude of 200’ below destination field elevation.

Cabin Altitude Selector

• The cabin altitude selector is located within the STANDBY section of thepressurization control panel. The selector is used to select the desiredcabin altitude during standby pressurization operations. The large diam-eter control sets 1000’ increments and the small diameter control sets10’ increments.



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Outflow Valve Position Indicator

• The outflow valve position indicator is located on the MANUAL sectionof the pressurization control panel and is used to indicate the position ofmain cabin outflow valve. The valve position indicator operates in allmodes. The indicator circuit utilities 115V AC power for potentiometerfeedback voltage.

Outflow Valve Switch

• The spring-loaded to center outflow valve switch is located within theMANUAL section of the pressurization control panel. The switch is usedto control the position of main cabin outflow valve.

• (Oral Topic) In AUTO and STANDBY modes of operations, the outflowvalve switch is nonfunctional. Electrical power is made available to theswitch in MANUAL mode only. By selecting the switch to either theCLOSED or OPEN positions, the outflow valve will move in the desireddirection until the switch is released. Failure of electrical power to eitherthe manual AC actuator or the manual DC actuator, the outflow valvewill remain in the last position.

Flight-Ground Switch

• The FLT/GRD switch is located within the bottom right corner of thepressurization control panel. Selecting the switch to the GRD positionon the ground, drives the pressurization outflow valve full open at acontrolled rate and depressurizes the aircraft. After takeoff, the previ-ously mentioned function is inhibited, and begins to functions the sameas the FLT position mode. During preparation for departure, the FLT/GRD switch is placed to FLT. This signals the cabin controller topartially close the outflow valve to control the cabin pressure to .1 PSID.This will hold the cabin pressure at approximately 189’ below runway.The outflow valve will be near the closed position at rotation, therefore,preventing a pressure bump at rotation. The FLT/GRD switch is usedonly in AUTO or STBY modes.



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Pressurization Mode Selector

CHECK The CHECK position of the pressurization mode selector tests theAUTO FAIL circuits by electronically actuating the circuit for anexcessive rate of change of cabin altitude. The AUTO FAIL andSTANDBY lights will illuminate.

AUTO The AUTO position of the pressurization mode selector is thenormal selection for flight operations. All settings are made duringpreflight and system operations are automatic for all flight phases.This mode uses the AC motor for control of the main outflowvalve. Aircraft altitude is sensed directly from the static ports.The barometric correction for these pressures comes from theCaptain’s altimeter. The AUTO mode of operation has 5 phasesof pressurization events: ground unpressurized, ground pressur-ized, ascent, cruise, and descent.

STBY The STBY position of the pressurization mode selector is used asan alternate mode of operation that bypasses the AUTO mode ofthe pressurization controller. Cabin altitude setting must be madefor each phase of flight. The STBY mode uses the DC motor forcontrol of the main outflow valve. Aircraft altitude is sensed di-rectly from the air data computer (ADC). The barometric correc-tion for these pressures comes from the First Officer’s altimeter.

AC MAN The AC MAN position of the pressurization mode selector is usedto manually control the position of the outflow valve. This isaccomplished by using the outflow valve switch, that directly con-trols the AC actuator. Electrical power is made available to theoutflow valve switch in MANUAL mode only. The pressure con-troller receives electrical power for AC MAN operations from the115V AC Transfer Bus.

DC MAN The DC MAN position of the pressurization mode selector is usedto manually control the position of the outflow valve. This isaccomplished by using the outflow valve switch, that directly con-trols the DC actuator. Electrical power is made available to theoutflow valve switch in MANUAL mode only. All auto and standbycircuits are bypassed. The pressure controller receives electricalpower for DC MAN operations from the 28V DC standby bus.



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Note: The automatic or standby modes of operation may be inoperative fordispatch provided the AC MAN and DC MAN actuators are operatingnormally. This must be verified before flight that both actuators areoperational by observing the outflow valve indicator for full movementto the open and closed positions. All modes of pressurization maybe inoperative for unpressurized flight provided the outflow valveremains open and extended operations over water are prohibited.Refer to your MEL.

Auto Mode Flight Profile

• The auto mode flight profile provides the means of controlling cabinpressure automatically from settings selected prior to the start of theflight. By selecting the appropriate flight altitude and destination fieldelevation, the controller will automatically establish the correct cabinpressure for climb, cruise and descent.

• Auto pressurization procedures:1. Prior to start of flight:

a. AUTO mode selected.b. Flight altitude set in FLT ALT readout.c. Destination elevation set in LAND ALT readout.d. Set -200’ below destination field altitude in CAB ALT.e. After engine start, place FLT/GRD to FLT position.

2. After Landing:a. Set the FLT/GRD switch to GRD.

• By placing the FLT/GRD switch to FLT, the outflow valve is driven bythe controller to maintain cabin pressure at approximately 189’ belowthe runway. After takeoff, the ascent schedule is controlled to maintainproportional changes from 189’ below departure elevation to the cruisealtitude set. When the airplane reaches cruise altitude, the controllerchanges from the ascent schedule to an isobaric schedule. An isobaricschedule is a constant pressure schedule based on the above men-tioned PSID. Maximum climb rate is 500 fpm, maximum descent rate is350 fpm (AUTO Mode). The dual differential pressure control systemcontrols cabin differential pressure at 7.5 psi for flights below 28,000’and at 7.8 psi for flights above 28,000’.



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• At the beginning of the descent phase, the controller switches fromisobaric to a descent schedule. The cabin descends to an altitudeapproximately 300’ below the LAND ALT set. The reason for the 300‘(0.15 PSID) is to ensure that the outflow valve is closed to avoid apressure bump on landing. On landing, the ground sensing relay sig-nals to the controller and the cabin pressure (via the outflow valve) tobring the cabin back to 189’ below the runway. As the plane taxis, thecrew changes the FLT/GRD switch to GRD, the controller signals theoutflow to open fully, bringing the cabin altitude to a field elevation.

Standby Mode Flight Profile

• The standby pressurization mode was designed as a backup in case ofa failure to the automatic pressurization mode. The automatic transferfeature from the auto mode to the standby mode has been provided incase of that failure. The standby mode provides the means of control-ling cabin pressure with the manual insertion of cabin altitude settingsfor each phase of flight. At the beginning of the flight, the controller isselected to the STANDBY mode. This arms the controller standbycircuits, thus directing all control signals to the DC actuator of the out-flow valve. The STANDBY light will illuminate, indicating that the con-troller circuits are armed. In the case of a STANDBY mode failure,there is no automatic transfer to the MANUAL mode of operations. Ifthe STANDBY green light extinguishes, the standby circuits are nolonger controlling the outflow valve. The outflow valve will remain in thelast position selected.

• Standby Pressurization Procedures:1. Prior to start of flight:

a. Select STANDBY mode (green light illuminates).b. Set PIP mark on the CABIN RATE control.c. Set -200’ field elevation on CABIN ALT readout.d. After engine start, place FLT/GRD to FLT position.

2. After T/O:a. Select the setting from the placard and place in CABIN ALT .

3. Top of Descent:a. Set -200’ below destination.

4. After Landing:a. Set the FLT/GRD switch to GRD.



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• A placard located under the pressurization control panel, has been pro-vided that displays proper cabin altitude vs. flight altitude information.The index mark (PIP mark) located on the CABIN RATE control, setsthe desired rate of change to approximately 300 ft/min. The minimum/maximum rate of ascent and descent of the cabin rate controller is 50'to 2000' respectively. The desired cabin schedule is maintained by themaximum differential limiting circuit. This circuit compares ambientpressure to the desired cabin pressure and limits the cabin pressuresignal to a maximum of 7.9 psi differential, at a flight altitude of 28,000’or higher. Cabin altitude in excess of 14,625’, a circuit in the controllerdrives the outflow valve to full closed.

Manual Mode Flight Profile

• The manual pressurization mode provides the means of manually con-trolling the cabin pressure by the position of the outflow valve. Manualcontrol of the valve is accomplished by the AC or DC actuator locatedon either end of the outflow valve unit. By selecting either AC or DCpositions on the pressurization mode selector, the respective AC or DCcontrol circuits will be armed. The source of electrical power for ACmode of operation comes from the No.2 AC transfer bus. Power sourcefor the DC mode of operation is from the DC standby bus. The illumina-tion of the green MANUAL light indicates the controller is in the manualmode of operation.

• Caution should be used when selecting manual AC or DC on the modeselector. The outflow valve will have an immediate response by thepositioning of this switch. Valve gate rotation (full movement) is approxi-mately four (4) seconds using the AC actuator and is approximatelyeight (8) seconds using the DC actuator. Full rotation is 85 degrees ofmovement from stop-to-stop in either mode.

Pressurization System Description

• (Oral Topic) The airplane is pressurized by bleed air, supplied to anddistributed by the air conditioning system. Pressurization and ventilationare controlled by varying the opening of the outflow valves. A propor-tional relationship is maintained between ambient and cabin pressure inthe climb or the descent, and maximum differential is maintained duringcruise.



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Pressure Relief Valves

• Two pressure relief valves provide maximum safety pressure relief. Thepressure relief valves limit the differential pressure to a maximum of8.65 psi. The pressure relief valves are located one on each side of theaft outflow valve, near the tail of the aircraft. The valves are actuatedby air pressure.

Negative Relief Door

• A negative relief (vacuum relief) door provides negative pressure relief.The door modulates open to outside air pressure differential of -1.0psi. The negative relief valve is located in the fuselage skin on the rightside below and forward of the aft service door (below the cabin floorlevel).

Main Outflow Valve

• The main outflow valve is located in the aft section of the pressurebulkhead. The valve consists of a aluminum rotating gate within arectangular frame. The AC/DC actuators are mounted on each end ofthe frame with only one actuator operating at any one time. The main



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outflow valve has a heating element installed that will maintain a tem-perature above 320F when the outflow temperature is below 400F anddeactivate when the outflow temperatures is above 700F. Overheatprotection has been provided to protect the outflow valve should tem-peratures exceed 1650F.

• When operating in AUTO or STANDBY, should the cabin pressure ex-ceed 14,625’, an aneroid switch in the pressurization controller will closethus energizing the K3007 relay that drives the DC actuator of theoutflow valve to the full closed position.

• The main outflow valve’s AC or DC actuator may be inoperative fordispatch provided the aircraft is used for pressurized cargo operationsonly, and that the aircraft must be depressurized before landing. TheAC and DC actuators may be inoperative for unpressurized flights pro-vided the outflow valve is selected to the open position. Extendedoverwater operations are prohibited. Further limitations and proceduresare listed in your MEL for operations with an inoperative main outflowvalve. Refer to your MEL.

Forward Outflow Valve

• The forward outflow valve is located on the left side of the E/E compart-ment. Air is drawn from the equipment cooling system and is exhaustedoverboard. The valve is driven by an AC motor to either the open orclosed positions. The forward outflow valve does not modulate. Whenthe main outflow valve is more than 30 opened, the forward outflowvalve will open. When the main outflow valve is within half degree ofbeing closed, the forward valve will close.

Flow Control Valve

• The flow control valve controls the exhaust vented from the electronicequipment compartment during ground operations, pressurized flight,and pressurized flight below a cabin differential pressure of 2.5 psi.When the valve is closed, the air is then routed forward to the forwardcargo compartment liner for inflight heating. The flow control valve islocated toward the forward bottom of the FWD Cargo compartment.



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Pressurization System Tests

• The following pressurization system tests are designed to check theoperation of the auto, standby, and manual modes of the pressurizationsystem. All system tests are performed with the air condition packswitches in the OFF position.

Auto Mode Test

• The objective of the auto mode test is to simulate a high rate of pres-sure change that provides an opportunity for the pressure rate detectorto detect an abnormal condition. This simulation will cause the pressur-ization controller to automatically switch to the standby mode. To ac-complish this procedure, the pressurization mode selector is placed tothe AUTO position. Move the FLT/GRD switch to GRD position. En-sure the cabin altitude indicator is selected to 500’. above field elevationand the cabin rate selector is on the INDEX mark. Mode lights areextinguished and the outflow valve is in the open position. Set the FLT /GRD switch to the FLT position and observe the outflow valve indicatorhas displayed the closure of the valve following a 10 second movementdelay.Observe the following:1. The AUTO FAIL light illumination.2. The standby light illumination.3. Check the outflow valve has moved towards the closed position.

Standby Mode Test

• The objective of the standby mode test is to ensure the opening andclosing of the outflow valve when selecting cabin altitudes above andbelow the current field altitude. To accomplish this procedure, set thecabin altitude indicator to 500’ below field elevation. Observe the out-flow valve’s movement towards the closed position. Select the FLT/GRD switch to the GRD position and observe the AUTO FAIL &STANDBY lights are extinguished and observe the outflow’s valvemovement towards the open position. Select the FLT/GRD switch tothe FLT position and ensure the system stays in the AUTO mode. Theoutflow valve position indicator will gradually move towards the closedposition.



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Manual Mode Test

• The objective of the manual mode test is to check, by slewing, themovement of the outflow valve via the manual process. This test mustbe run immediately after the auto and standby checks have been com-pleted. If the initial check inputs have cleared the test circuit (after 30seconds) the AUTO FAIL and STANDBY lights will not illuminate. Toinitiate the system check, place the pressurization mode selector toAUTO and observe the illumination of the AUTO FAIL & STANDBYlights. Place the pressurization mode selector to MAN AC and observethe AUTO FAIL & STANDBY lights are extinguished. After the illumina-tion of the MANUAL light, check the movement of the outflow valve inboth directions by using the outflow valve switch. Place the pressuriza-tion mode selector to MAN DC and observe the AUTO FAIL &STANDBY lights are extinguished. After the illumination of theMANUAL light, check the movement of the outflow valve in both direc-tions by using the outflow valve switch. To complete the test, select theFLT/GRD Switch to GRD and the pressurization mode selector toAUTO; observe if the MANUAL light has extinguished and the outflowvalve has moved to the OPEN position.

Pressurization Limitations

• Maximum differential pressure: 8.65 psi

• Operating differential pressure: 7.5 +/- .1 psi (35,000' controller)7.8 +/- .1 psi (37,000' controller)

• Maximum cabin differential pressurefor takeoff and landing: .125 psi

Air Conditioning & Pneumatics

• Uses 8th/13th stage air.• Bleed valve electrical on/off.• 13th stage modulation/shutoff.• RH side 8th/13th pre-cooled.• Air cleaners within system.• Purge valve within system.• Pack valve two rates of flow.• Compressor overheat 3650F.



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• Separator water dump overbd.• Gasper fan system.• Hyd/water pressure - engine.

Air Conditioning & Pneumatic Sensors Locations

Sensor 200Duct limit sensor. 1400FWater separator. 350FRam air temp. 2300FPass. cabin temp. 35-2000FSupply duct temp. 35-2000FBleed valve overheat 4900FTurbine inlet. 2100FCompressor discharge. 3650FDuct overheat. 1900FDuct overheat. 2500F

Pneumatic Ground Cart

• The external pneumatic ground cart can be directly connected to thebleed manifold. The connection point is located on the right side of theisolation valve. Normal operating procedures requires the battery switchto be selected ON. This provides power to various protective sensorsand circuits within the pneumatic system. The pneumatic ground cartcan be used to operate both air conditioning packs, providing that theground cart can maintain a minimum 20-25 psi. For extra cooling,operate the left pack using the APU bleed source and the right packusing the external pneumatic source. The isolation valve should beplaced to CLOSED. This technique can also be used when a limitedamount of external pressure is available.



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Air Temperature Source Selector

• The two position air temperature source selector is located on the topcenter of the temperature control panel. By selecting the supply ductposition, the temperature sensed from the main distribution supply ductwill be displayed on the temperature indicator. Selecting the passengercabin position, temperature sensed in the forward cabin ceiling area willbe displayed on the temperature indicator. The supply duct indicatormay be inoperative for dispatch provided both duct overheat warningsystems are functional. Refer to your MEL.

Air Mix Valve Indicator

• The air mix valve indicators are located on the temperature controlpanel. The indicators provide a visual reference of the position of theair mix valves. The design purpose of the air mix valve is to control thepack output temperature by directing airflow through the pack andaround the pack to the air mixing chamber.



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Air Mix Valves

• Each pack has two air mix valves that controls the hot and cold airwhich is used to meet the selected temperature requirements. Bothvalves are connected by a common shaft, with each valve plate offsetby 900 degrees. This 900 degrees offset causes each valve opening andclosing to be the opposite of each other. The air mix valves can becontrolled either automatically or manually by using the passenger cabintemperature selector in the AUTO mode or in the MANUAL mode.

• The physical location of the air mix valves are inboard of the secondaryheat exchanger. The air mix valves may be inoperative for the respec-tive inoperative pack providing operational limitations for the inoperativepack are followed. Refer to your MEL for further details.

• (Oral Topic) There are three ways the air mix valves drive to the fullcold position:1. Duct overheat.2. Pack trip-off.3. Selection of the pack switch to the OFF position.

Duct Overheat Light

• (Oral Topic) The illumination of the amber DUCT OVERHEAT lightindicates an overheat condition exists within the passenger cabin duct.Upon sensing the overheat condition, the temperature mix valves willdrive to full cold.

• Two thermal sensors are used to provide warning and detection of theseduct overheats. The 880C cabin duct sensor is used in the temperaturecontrol system and the 1210C cabin duct sensor is used in the pack trimsystem. Should the duct temperature exceed 880C, air mix valves willmove to the full cold position. Moving the temperature selector to acooler temperature will prevent the mixing valves from programmingback to an overheat condition. Use the TRIP RESET switch to reset thesystem as soon as the duct cools.

• The 1210C has been provided as a backup sensor, and will trip the packshould the cabin duct temperatures exceed the preset limits. The ductoverheat warning lights may be inoperative for dispatch provided thesupply duct temperature indicators operate normally.



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Passenger Cabin Temperature Selector

• The passenger cabin temperature selector can be selected for use inthe automatic or manual mode of operation. When placed to AUTO, theautomatic temperature regulator controls passenger cabin temperatureas selected. The temperature sensor is located in the cabin ceiling andthe associated regulator is located in the electronic equipment bay.

• The selection of the MANUAL mode, provides direct control of the airmix valves by the flight crew. The temperature control system has atopping circuit of 600C. The purpose of this topping circuit is to preventpossible pack trip-offs.

• The MEL describes procedures to check the operation of the manualand automatic modes of the temperature control system. To check theproper operation of the automatic mode is accomplished by selecting awarmer then a colder temperature, and observe that the air mix valveindicator moves towards HOT and then COLD.

• Checking the proper operation of the manual mode is accomplished bymomentarily selecting WARM and then COLD and observing mix valveindicator move toward HOT and then COLD. The indicator should stayin the last position when the selector is released when performing themanual check. These checks can be used for both the passenger andflight deck systems.

• The manual or automatic mode of the passenger cabin control tempera-ture system or the flight deck temperature control system may be inop-erative for dispatch if the respective pack is operational. If both modesare inoperative, the respective pack should not be used. Observe packinoperative limitations. Refer to your MEL.

Gasper Fan Switch

• The gasper fan switch controls the gasper fan which is used to increasethe airflow to the individual gasper air outlets. The gasper air systemuses cold air tapped off the supply line from the right pack that leads tothe mixing chamber. The gasper fan is a 115V AC, 3-phase, motor-operated fan that is contained in the right side of the air conditioningdistribution bay. Overheat protection is provided by the circuit breaker



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system and internal thermal protectors contained within the fan motor.The gasper fan may be inoperative for dispatch with no limitations ap-plied. Refer to your MEL.

• (Oral Topic) The gasper fan system normally is supplied air from theright air conditioning pack. With an inoperative right pack, conditionedair from the supply duct can still flow through the gasper air system.

Wing-Body Overheat Test Switch

• A wing-body overheat condition is normally caused by a bleed air ductleak. The wing-body overheat test switch has been provided to test thewing-body overheat detector circuits. Pressing the test switch willcause both WING-BODY OVERHEAT lights to illuminate, indicating thesystem is operating normally. There are 6 primary sensors locatedthroughout the wing and body segments. Each sensor is preset for2550F limits. Depending on aircraft model, the test may take as long as10 seconds to accomplish a complete self test.

• With the illumination of the wing-body overheat light, non-normalprocedures calls for the closure of the isolation valve switch. This willprevent the isolation valve from opening when the affected side enginebleed switch is selected OFF. Depending on which light is illuminated(left or right), the checklist will guide you to isolate the bleed sourcesupplying the hot air.

• For a left wing-body overheat light illumination, select the APU bleed airswitch to the OFF position. This will stop the APU bleed air flow fromentering the left side of the pneumatic ducting. Should the light stillremain illuminated, the leak is in the APU compartment, therefore,select the APU switch to the OFF position. Refer to your non-normalchecklist.

• A bleed air leak or rupture is a serious condition. Bleed air is routedthrough the aircraft under high pressure and temperatures. This air isdirected through stainless steel manifolds in the strut to the duct in theleading edge of the wing. A titanium duct runs from the APU along theleft side of the aft cargo compartment, then inside the keel beamthrough the wheel well and air conditioning bay where it joins the cross-over duct to the left side of the isolation valve. Therefore, treat this non-normal with respect.



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• The aircraft may be dispatched with an inoperative left wing-body over-heat detector provided the right pack and engine bleed is used forpressurization only. The use of the APU is prohibited except for enginestarting. This includes using the APU for electrical power source. Theisolation valve and left engine bleed valve must remain closed for alloperations except engine starting. The aircraft must not be operated inknown or forecasted icing conditions. With the left pack not operational,flights must remain below FL 250.

• The aircraft may be dispatched with an inoperative right wing-body over-heat detector provided the left pack, left engine bleed, and APU bleedair is used for pressurization only. The isolation valve and right enginebleed valve must remain closed for all operations except engine starting.The aircraft must not be operated in known or forecasted icing condi-tions. With the left pack not operational, flights must remain below FL250. Refer to your MEL for operational details.

Dual Bleed Light

• The purpose of the dual bleed warning light is to alert the crew of thepotential that one of the engine bleed switches and the APU bleedswitch are in positions that could result in excessive bleed pressurizingthe duct simultaneously. This condition could cause back pressure tothe APU. The illumination of this amber light indicates one of thefollowing condition may exist:1. APU bleed air valve is OPEN and the No.1 engine bleed switch is

ON.2. The No.2 engine bleed switch is ON, the APU bleed air valve and

isolation valve are OPEN.

• (Oral Topic) The illumination of this light pertains to the position of thebleed switches and the APU bleed switch with reference to excessivebleed pressure. This excessive bleed pressure may also exists if exter-nal bleed air is being used and if the isolation valve is selected open.With this configuration, the dual bleed light will not illuminate. The dualbleed light is like a circuit, to complete the circuit various switches mustbe certain positions. This is in reference to the engine bleed switchesand the opening of the APU bleed valve. Electrical power for this circuitis provided by the 28V DC battery bus.



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• The dual bleed light system may be inoperative for dispatch providedthe APU bleed air is not used during flight operations and the bleedvalve is closed before each departure. Refer to your MEL for proce-dures that are used to verify the closure of the APU bleed valve.

Ram Door Full Open Light

• The illumination of the blue ram door full open light indicates the ramdoor is in the full open position. The purpose of the ram air system is toautomatically control the outside airflow to the heat exchangers so thatthe cooling packs maintain a constant operating temperature. A 2300Fsensor located in the ACM compressor discharge duct signals the ramair controller to control the airflow through the heat exchanger systemby modulating the mechanically linked ram door and exit louvers. Thisoperation is automatic. During ground operations or during flight withthe flaps extended, the ram air door will position itself to the full openposition for maximum cooling. During flight operations (flaps retracted),the ram air door will modulate between normal open and normal closedpositions. During high ACM temperature conditions (temperatures ex-ceeding 2300F), the ram air door will automatically open.

• The pack ram air system may be inoperative for dispatch providedoperations are not conducted on gravel runways or runways coveredwith slush and/or standing water. The respective pack must also beselected to OFF during takeoffs and landings on wet runways and theisolation valve should be selected to the closed position. Refer to yourMEL.



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Turbofan System

• The turbofan system has been provided to help augment the flow of airthrough the heat exchangers. This augmented airflow occurs duringground operations and when the flaps are not in the up position. Thisoperation is automatic and is controlled by the turbofan control valve.The valve is activated electrically when the respective pack is on, andwhen the air ground safety sensor/flap limit switch is activated. Pneu-matic air is used to operate this air turbine. This augmented airflowcooling is provided to the heat exchangers by drawing outside air intothe ram air ducts by the turbofan.

• The turbofan system may be inoperative for dispatch provided the re-spective pack is operated only in flight with the flaps retracted. With theflaps extended, the respective pack switch should be placed to the OFFposition and the isolation valve should be selected to the closed posi-tion. Refer to your MEL.

F Outflow Closed Light

• The illumination of the blue forward outflow closed light indicates theforward outflow valve is closed. With the valve closed, the main outflowvalve is within a half degree of being closed. With the valve open, themain outflow valve is 30 degrees or more open. The forward outflowvalve closure light may be inoperative for dispatch with no special limita-tions.



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Pneumatic Duct Pressure Indicator

• The pneumatic system has been provided with two pressure transmit-ters that monitors pressure in the left and right pneumatic ducts. Theduct pressure gauge is located on the air conditioning control panel andis calibrated for 0-100 psi indications. The power source for this indica-tor is 28V AC from the No.1 transfer bus. The sensor is located justafter the engine bleed valve and slightly before the pack valves.

• The pneumatic duct pressure indicator may be inoperative for dispatchprovided the flight crew uses the start valve open lights to verify theclosure of the engine start valve during engine starting. For aircraft thatdo not have these lights, the APU may be used to determine startercutout by monitoring APU EGT levels. The closure of the engine startvalve is indicated on the APU EGT indicator as a distinct drop in APUEGT at starter cutout. Refer to your MEL.



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Air Conditioning Pack Switch

• The respective air conditioning pack valve is controlled by a two positionswitch. Selecting the pack switch to the OFF position, closes the packvalve and terminates airflow into the air conditioning pack. Selecting thepack switch to ON, allows metered airflow into the pack at a rate of 80lbs/min. If the APU is the source of this airflow, the rate is increased to90 lbs/min. The valve is electrically controlled (battery bus powered)and pneumatically operated. The pack shutoff valves may be inopera-tive for the respective inoperative pack provided the associated limita-tions are observed. With one pack inoperative, you must limit youraltitude to 25,000’. Refer to your MEL.

Isolation Valve Switch

• The bleed air isolation valve has been provided so that the respectivepneumatic system (left and right side) can be separated or connectedas desired. The valve is located in the crossover duct within the keelbeam area, just right of the APU duct junction. The valve switch is athree position switch, powered by the No.1 transfer bus. Selecting theswitch to the CLOSE position, closes the isolation valve. Selecting theswitch to the OPEN position, opens the circuit for the isolation valve.The AUTO position, will close the isolation valve if “ALL” engine bleedand air conditioning pack switches are ON. The same valve action willautomatically open the isolation valve if “any one” of the engine bleed orair conditioning pack switches are selected to the OFF position.

• (Oral Topic) During flight operations, the isolation valve switch is nor-mally placed in the AUTO position. The actual valve position is con-trolled by the position of the bleed switches and pack switches and NOTsolely by the position of the isolation valve switch.

• The isolation valve switch is used in the wing-body overheat checklist.This non-normal procedure requires the isolation valve switch to beplaced in the CLOSED position. This action prevents unwanted openingof the isolation valve when procedural checks calls for the associatedbleed air switch to be selected OFF. The isolation valve switch mayalso be used with an external pneumatic bleed air cart. The isolationvalve switch should be selected to the OPEN position. This will allowexternal bleed air to be supplied to the left side of the pneumatic mani



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fold, thus allowing operation of the left pack (if desired). The operationof two packs at the same time from one bleed air source is permittedprovided the external air cart can maintain 20-25 psi with both packsoperating. Please observe the maximum external bleed air pressurelimitation of 60 psig and/or 2320C. Another situation that requires theuse of the isolation valve switch is the selection of the switch to theCLOSED position for an unpressurized takeoff and landing. Refer toyour airline supplementary procedures or AFM for details.

• The bleed isolation valve may be inoperative for dispatch provided flightoperations into known or forecasted icing conditions is prohibited. Referto your MEL.

• (Oral Topic) The APU bleed valve has no affect on the operation of theisolation valve.

Pack Trip Off Light

• The illumination of the amber pack trip off light indicates the respectiveair conditioning pack has tripped to OFF due to an overheat. Thisoverheat condition may have occurred in the compressor outlet duct(1850C), in the turbine inlet duct (990C), or in the supply duct (1210C).Upon the detection of the overheat condition, the pack valve will auto-matically close and the air mix valve will drive to the full cold position.Non-normal procedures calls for the temperature selector to be placedto a warmer temperature setting. This will help reduce the work load ofthe air conditioning system by reducing the demand for colder air. Thetrip reset switch has been provided to reset the system once the packhas cooled below the preset limits. This procedure is for compressor orturbine malfunctions only.

• The respective pack trip warning system may be inoperative for dis-patch for an associated inoperative pack provided the MEL limitationsare followed concerning single pack operations and altitude restrictions.Refer to your MEL with reference to those procedures concerning theplacement of the isolation valve switch to the CLOSED position andpack switch positions.



B737-200

Wing Anti-ice Schematic Decal

• (Oral Topic) The wing anti-ice schematic decal is located on the airconditioning control panel, just right of the pack trip off warning lights.The purpose of the wing anti-ice schematic decal is to provide a quickvisual reference of the schematic relationship between the wing anti-icesystem and the pneumatic system. The key point to remember is thatthe wing anti-ice system taps into the pneumatic duct lines downstream of the bleed switches and upstream from the pack switches.Therefore, the position of the pack switches have no effect on theoperation of the wing anti-ice system.

Wing-Body Overheat Light

• The illumination of the amber wing-body overheat light indicates a pneu-matic duct leak. There are 6 primary sensors located throughout thewing and body segments that help identify this hot bleed air leak. Tem-peratures exceeding 2550F are normally associated with a warning light.

• (Oral Topic) The location for these overheat sensors is a common oralquestion for instructors and examiners. The illumination of the left wing-body overheat light indicates a “bleed air” leak (overheat) in the follow-ing areas:1. Bleed air duct in the left engine strut.2. Left wing leading edge.3. Left air conditioning bay.4. Keel beam area.5. APU bleed air duct.

• (Oral Topic) The illumination of the right wing-body overheat light indi-cates a leak (overheat) in the following areas:1. Bleed air duct in the right engine strut.2. Right wing leading edge.3. Right air conditioning bay.

• Refer to the wing-body test switch for detailed information concerningthe wing-body overheat warning system.



B737-200

Bleed Trip Off Light

• The illumination of the amber bleed trip off light indicates excessivetemperatures exist within the engine bleed air system. The overheatswitch is located downstream of the precooler. With temperatures ex-ceeding 2540C, the respective bleed air valve will automatically close.The bleed trip off light will remain illuminated until the manual resetbutton is used to reset the system. The system temperature must havecooled approximately 10% before the reset function is used.

• During bleed trip-off conditions, the bleed air valve will automaticallyclose, thus causing the loss of bleed air to the respective system. Withthe loss of bleed air, the respective pack valve will close and the mixvalve will move to the full cold position. Non-normal procedures directsyou to use the reset button once the system has cooled. Should thebleed trip off light remain illuminated and if wing anti-ice is required, theselection of the pack switch to the OFF position will cause the isolationvalve to open. This will provide bleed air from the opposite side for winganti-ice protection.

• (Oral Topic) Operating two packs from the same bleed source (oneengine) is not recommended.

• The engine bleed trip off lights may be inoperative for dispatch providedthe respective engine bleed is not used except for engine starting andthe airplane is not operated in known or forecasted icing conditions.Refer to your MEL for one engine bleed unusable and two enginebleeds unusable procedures.

Bleed Trip Off Light

• The illumination of the amber bleed trip off light indicates excessivetemperatures and pressures exist within the engine bleed air system.The overheat switch is located downstream of the precooler. The over-pressure sensor is a separate switch and is located within the pressureregulator. With temperatures exceeding 2540C and pressures exceed-ing 180 psi, the respective bleed air valve will automatically close. Thebleed trip off light will remain illuminated until the manual reset button isused to reset the system. The system temperature must have cooledapproximately 10% before the reset function is used.



B737-200

• During bleed trip-off conditions, the bleed air valve will automaticallyclose, thus causing the loss of bleed air to the respective system. Withthe loss of bleed air, the respective pack valve will close and the mixvalve will move to the full cold position. Non-normal procedures directsyou to use the reset button once the system has cooled. Should thebleed trip off light remain illuminated and if wing anti-ice is required, theselection of the pack switch to the OFF position will cause the isolationvalve to open. This will provide bleed air from the opposite side for winganti-ice protection. Flight operations with two packs operating from oneengine bleed source is not recommended.

• (Operational Tip) Following a bleeds-off takeoff, should the bleed tripoff light illuminate and normal reset is not possible, the following tech-nique may be used to reset the system. This technique should beaccomplished at altitudes above 1500’ AGL and TAT at/or below 380C.Select the respective engine anti-ice switch to ON, select the resetswitch, reconfigure the pressurization system, and then select the en-gine anti-ice switch to OFF. These actions will normally reset the bleedtrip-off circuit and reconfigure the aircraft. The reason for this type ofbleed trip-off occurring is that the relief valve upstream of the bleedvalve does not have enough flow capacity to limit the pressure in theduct below the overpressure switch. Follow your airline and themanufacturer’s approved procedures.

• The engine bleed trip off lights may be inoperative for dispatch providedthe respective engine bleed is not used except for engine starting andthe aircraft is not operated in known or forecasted icing conditions.Refer to your MEL for one engine bleed unusable and two enginebleeds unusable procedures.

Engine Bleed Air Switch

• The two position engine bleed air switch controls the respective enginebleed valve. The bleed air switch provides the means to terminate theflow of bleed air from the engines to the pneumatic manifold. The valveis driven by an AC electric motor and circuit. The position and control ofthe valve can also be initiated by the engine fire switch and by anoverheat sensor located downstream of the precooler. Pulling the en-gine fire switch will automatically close the engine bleed valve. Bleedair temperatures exceeding 4900F will also close the bleed air valve.



B737-200

• The pneumatic system has been designed to accept bleed air from therespective source with pressures up to 106 psi and temperatures of4500F. The actual air volume with two packs operating is 160 poundsper minute. Engine starting is approximately 110 pounds per minute.For wing anti-ice applications, approximately 120 pounds per minute.

• The engine bleed air shutoff valves may be inoperative for dispatchprovided the valve is secured closed after starting the engine. Theaircraft must not be dispatched into known or forecasted icing condi-tions.

Modulating and Shutoff Valve

• The control and design functions of the modulating and shutoff valve areautomatic. The B737-200 utilizes bleed air obtained from the 8th and13th stage engine bleed ports. Should 8th stage bleed air be insufficientfor aircraft demands or requirements, the 13th stage Modulating andShutoff valve modulates open to supply increased air flow. A tempera-ture sensor is used to modulate the 13th stage bleed air to prevent thebleed air from exceeding 4500F. Do not confuse this valve with theengine bleed valve, pressure relief valve, or the pressure regulator.Each unit has a different function and location within the pneumaticsystem. The pressure relief valve has been designed to provide protec-tion against pressures above 106 psi.

• The bleed air modulating and shutoff valve may be inoperative for dis-patch provided the aircraft is not operated in known or forecasted icingconditions. Check system operation, should the valve modulate out ofthe closed position, the respective bleed air switch must be selectedOFF after engine starting. Refer to your MEL for further information andlimitations.



B737-200

No Bleed Takeoffs - C Flow

• When configuring the bleed air control panel for a no engine bleedtakeoff, the following technique may be helpful in remembering theproper switch and valve position required for this departure configura-tion. The C-flow begins by placing the respective switches in the follow-ing order:1. Right pack switch ON.2. Isolation valve switch CLOSE.3. Left pack switch ON.4. Left bleed switch OFF.5. APU bleed switch ON.6. Right bleed switch OFF.

• Upon completion of the takeoff profile, the bleed switches can bereconfigured by reversing the C-Flow. This is normally accomplishedafter reaching at least 1500’ or until any obstacle clearance altitude hasbeen attained. Additional note, when operating in icing conditions, taxiwith the engine bleed air switches ON and the APU bleed switchesOFF. This will ensure adequate bleed air for anti-ice requirements.

Trip Reset Switch

• The trip reset switch is used to reset a bleed trip-off, pack trip-off, or aduct overheat. The fault that has cause the trip-off or overheat condi-tion has to be corrected or removed before attempting a reset. Therespective condition warning light will remain illuminated until the resethas been made.

APU Bleed Air Switch

• The two-position APU bleed air switch controls the APU bleed valve.The valve is of a modulating valve type used to control airflow from theAPU to the aircraft’s pneumatic system. With APU operations at orabove 95%, the bleed air switch can be selected to ON, thus providingbleed air as required to the aircraft’s system.

• APU bleed air may be used during flight operations and/or duringground operations. The APU bleed air valve is DC controlled and pneu-matically operated. During shutdown, the valve will automatically close.



B737-200

• (Oral Topic) APU bleed air has the following limitations: APU bleed airusage is limited to a maximum altitude of 17,000’. APU bleed andelectrical load is limited to a maximum altitude of 10,000’.

• The APU should be operated for a minimum of two minutes prior toselecting the APU bleed air switch. This allows time for the APUtemperatures and operations to stabilize prior to high bleed air demandsand high internal temperatures. The APU bleed valve should be closedwhen:1. Ground air is connected, with the isolation valve opened.2. Left engine bleed valve is in the opened position.3. The right engine bleed valve and the isolation valve are in the

opened positions, the APU bleed valve may be open during enginestart, but avoid power applications above the idle position.

• The APU bleed valve should be closed during ground operations requir-ing engine anti-icing. This will ensure that sufficient engine bleed air isavailable for cowl anti-icing. The reason for this procedure is to preventthe possibilities of APU bleed air back-pressures causing the 9th stagevalve to close.

• The APU bleed air valve may be inoperative in the closed position fordispatch purposes providing the APU is only used to provide electricalpower. Refer to your MEL.



B737-200

Air Conditioning Pack

Primary & Secondary Heat Exchangers



B737-200

Outboard Landing Lights

• The outboard landing lights are controlled by two three-positionswitches, located just below the overhead panel. The RETRACT posi-tion when selected, retracts the outboard landing lights and the lightswill be extinguished. The EXTEND position will extend the outboardlanding lights, but the lights will remain extinguished until the outboardlanding light switch has been selected to the ON position. Then all theoutboard lights will illuminate.

• The actual outboard landing lights are of the retractable type, located inthe outboard flap track fairing. These lights may be extended at anyaircraft speed. They shine forward approximately parallel to the water-line of the aircraft regardless of flap position. With the outboard landinglights extended throughout the entire flight, expect to see approximatelya 1% increase in flight plan fuel usage (reference information from theMEL).

• The power source for the left outboard light is from the 115V AC busNo.1. The power is then reduced by transformers to 16.5V AC. Thepower source for the right outboard light is from the 115V AC bus No.2.A thermostatic switch has been provided to prevent the possibilities of amotor overheat during retraction/extension process.



B737-200

• One landing light may be inoperative on each side for dispatch providedone of the two operating lights is in the inboard position and the tworemaining lights operate normally. The outboard extend/retract motormay be inoperative for dispatch provided the light is in the extendedposition and the light operates normally. Refer to your MEL.

Inboard Landing Lights

• The inboard landing lights are controlled by two ON/OFF switches,located just below the overhead panel. These lights are located in theleading edge (near the fuselage), adjacent to the runway turnoff lights.The lights shine forward and down towards the ground and ahead of theaircraft’s line of travel. The power source for the left inboard light isfrom the 115V AC bus No.2. The power is then reduced by transform-ers to 16.5V AC. The power source for the right inboard light is fromthe 115V AC bus No.1. The inboard light’s buses are electrically cross-connected.

Runway Turnoff Lights

• The runway turnoff lights are controlled by two ON/OFF switches, lo-cated just below the overhead panel. These lights are located in theleading edge (near the fuselage), adjacent to the inboard landing lights.



B737-200

The lights shine outboard 300 and have a beam width of 500. The powersource for the left runway turnoff light is from the 28V AC bus No.1.The power source for the right runway turnoff light is from the 28V ACbus No.2. Both lights may be inoperative for dispatch. Refer to yourMEL.

Taxi Light

• The taxi light switch controls the illumination of the single nose-wheel,strut-mounted light. This light is mounted on the nose-strut, thereforethe taxi light will point in the same direction as the nose-wheel. Powersource for the taxi light is from the 28V AC bus No.1. Many airlinesrecommend the taxi light should not be used for takeoff and landing.This will help to increase the service life of the light and reduce mainte-nance costs. The taxi light may be inoperative for dispatch. Refer toyour MEL.



B737-200

Position Lights (as installed)

• The position lights switch is a three-position toggle switch that controls theillumination of the navigation lights and the selection of the respectivepower source. Selecting the ON BAT position will illuminate the navigationlights. Powered for this position is from the 28V DC battery bus (if no otherpower is available). The battery switch must be ON to provide this function.The ON position of the position lights switch illuminates the red-green wingtip navigation lights and the white trailing edge tip lights. Power for thisposition is from the 28V AC No.2 transfer bus.

• The navigation lights consist of two fixed green lights facing right out-board and forward, two fixed red lights facing left outboard and forward,and one fixed white light facing aft on each wing-tip trailing edge.

• Various navigation lights may be inoperative for dispatch provided thefollowing combination exists. One stationary red wing tip bulb, onestationary green wing tip bulb, and one stationary white tail light at eachwing position operate normally. All navigation lights may be inoperativeduring day flight operations only. Refer to your MEL.



B737-200

Position Lights (Configuration Two)

• The position lights switch is a three-position toggle switch that controlsthe illumination of the navigation lights and the selection of the type oflight desired. Selecting the STROBE/STEADY position will illuminatethe red/green navigation lights, the white trailing edge wing-tip lights,and the wing-tip/tail strobe lights. Selecting the STEADY position,illuminates the red/green wing-tip navigation lights and the white trailingedge wing-tip lights only. With an interruption of normal power (loss ofall generators), the navigation and the white position lights will operateonly if the switch is in the steady position. All navigation lights may beinoperative during day flight operations only. Refer to your MEL.

• The navigation lights for this configuration consist of one fixed greenand one high intensity white strobe light facing right outboard and for-ward. One fixed red and one high intensity white strobe light facing leftoutboard and forward. One fixed white light facing aft on each wing tiptraining edge. One strobe light located on the tail cone above the APUexhaust.



B737-200

Anti-Collision Light Switch

• The two-position anti-collision light switch controls the illumination of thetwo red rotating beacon lights. These lights are located on the upperand lower fuselage. Each light consists of a xenon arc flashtube lampand solid state circuits for power storing, timing and triggering the flash-tube. The timing circuit controls the flash illumination at a rate of oneflash every second. Caution should be used when completing a pre-flight of the light unit. The light unit contains high energy voltage. Thepower source for the upper light is from the No.1 115V AC transfer bus.The power source for the lower light is from the No.2 115V AC transferbus.

• The anti-collision beacons may be inoperative for night dispatch pro-vided the wing tip and white tail strobe lights are installed and areoperating normally. Refer to your maintenance department for refer-ence to the type of beacon installed on your aircraft. Different beaconshave different limitations with regards to the number of inoperative bulbsfor FAR requirements. Refer to your MEL.

Wing Illumination Switch

• The two-position wing illumination switch controls the illumination of thetwo white leading edge lights (one on each wing). These lights are flushmounted on the fuselage, forward of the wing. The beam is directed toilluminate the leading edge of the wing with a beam width of 130. As asecondary function, the wing inspection lights may assist in operationand servicing the aircraft by ground crews. The power source for thewing inspection lights is from the 28V AC ground service bus.

• The wing inspection lights may be inoperative for night dispatch pro-vided a portable light is available when dispatching into known or fore-casted icing conditions. This portable light must be of suitable capacityfor proper wing inspection. For flight operations during day hours, thelights may be inoperative for dispatch. Refer to your MEL.



B737-200

Wheel Well Lights

• The two-position wheel well switch controls the illumination of the wheelwell dome light and the landing gear inspection floodlights. A secondcontrol switch is located on the external power receptacle panel thatcontrols the nose wheel well lights only. The power source for wheelwell lights is from the 28V AC ground service bus.

• The inspection flood lights for the main gear may be inoperative fordispatch during day operations only. The flood lights may also beinoperative provided an alternate landing gear indicating system (otherthan viewer type and center panel system) is installed and operatesnormally. The dome lights may be inoperative for dispatch. Refer toyour MEL.

Logo Light Switch

• The two-position Logo Light Switch controls the illumination of the twovertical fin inspection and insignia identification lights (one on eachside). The logo light system may be inoperative for dispatch withoutany limitations. Refer to your MEL.



B737-200

Emergency Cockpit Lighting

• (Oral Topic) The cockpit lighting has been designed so that a powerfailure of either bus will result in only a partial failure of the cockpit and/or passenger lighting system. With the total loss of all AC power, thefollowing lights will be powered by the battery bus.1. Standby compass light.2. White dome light.3. Emergency instrument flood light.4. Selected system/warning lights.5. Lavatory dome light.6. Emergency exit lights (internal battery).

• (Oral Topic) The emergency instrument flood lights will automaticallyilluminate upon the failure of the No.2 AC transfer bus.

• (Oral Topic) The dim entry lights and the fluorescent mirror lights in thelavatories are powered from the hot battery bus. Therefore, with thebattery switch in the OFF position, these are the only lights that can beilluminated.



B737-200

Engine Start Panel (as installed)

• The GRD position of the engine start panel provides high energy ignitionto both igniters when the engine start levers are placed to the idleposition. This pulsating high energy power is applied to plugs in boththe No.4 and No.7 burner cans. Power source for the No.4 plug is fromthe battery bus. Power source for the No.7 plug is from the AC transferbus. The solenoid held-spring loaded switch also positions the startervalve to the open position. Power for the starter valve is from thebattery bus.

• The OFF position of the engine start panel provides no ignition. TheLOW IGN position of the engine start panel provides low energy con-tinuous ignition to one igniter only in the No.7 burner can when theengine start levers are in the idle position. Power source for the No.7plug is from the AC transfer bus. This position is used to improve igniterservice life and to minimize the possibilities of engine flameouts duringtakeoff, landing, turbulence, and during flight operations in icing condi-tions.

• The FLT position of the engine start panel provides high energy ignitionto both igniters when the engine start levers are in the idle position.This pulsating high energy power is applied to plugs in both the No.4and No.7 burner cans. Power source for the No.4 plug is from thebattery bus. Power source for the No.7 plug is from the AC transferbus.



B737-200

• The left ignitor (No.7) of the high energy system may be inoperative oneach engine for dispatch purposes with no special limitations applied.The low energy system may be inoperative for dispatch provided switch-ing is available to permit selection of the operative high energy systemfor continuous ignition. Refer to your MEL.


• The GRD position of the engine start panel provides high energy ignitionto both igniters when the engine start levers are placed to the idleposition. This pulsating high energy power is applied to plugs in boththe No.4 and No.7 burner cans. Power source for the No.4 plug is fromthe AC standby bus. Power source for the No.7 plug is from the ACtransfer bus. The solenoid held-spring loaded switch also positions thestarter valve to the open position. Power for the starter valve is fromthe battery bus.

• The OFF position of the engine start panel provides no ignition. The LIGN position of the engine start panel provides high energy continuousignition to one igniter only in the No.7 burner can when the engine startlevers are in the idle position. Power source for the No.7 plug is fromthe AC transfer bus. This position is used to improve igniter service lifeand to minimize the possibilities of engine flameouts during takeoff,landing, turbulence, and during flight operations in icing conditions.

• The R IGN position of the engine start panel provides high energycontinuous ignition to one igniter only in the No.4 burner can when theengine start levers are in the idle position. Power source for the No.4plug is from the AC standby bus. This position is used to improveigniter service life and to minimize the possibilities of engine flameoutsduring takeoff, landing, turbulence, and during flight operations in icingconditions.

• The FLT position of the engine start panel provides high energy ignitionto both igniters when the engine start levers are placed to the idleposition. This pulsating high energy power is applied to plugs in boththe No.4 and No.7 burner cans. Power source for the No.4 plug is fromthe AC standby bus. Power source for the No.7 plug is from the ACtransfer bus.



B737-200


• The GRD position of the engine start panel provides high energy ignitionto both igniters when the engine start levers are placed to the idleposition. This pulsating high energy power is applied to plugs in boththe No.4 and No.7 burner cans. Power source for the No.4 plug is fromthe AC standby bus. Power source for the No.7 plug is from the ACtransfer bus. The solenoid held-spring loaded switch also positions thestarter valve to the open position. Power for the starter valve is fromthe battery bus.

• The OFF position of the engine start panel provides no ignition. The Aposition of the engine start panel provides high energy continuous igni-tion to one igniter only in the No.7 burner can when the engine startlevers are in the idle position. Power source for the No.7 plug is fromthe AC transfer bus. This position is used to improve igniter service lifeand to minimize the possibilities of engine flameouts during takeoff,landing, turbulence, and during flight operations in icing conditions.

• The B position of the engine start panel provides high energy continu-ous ignition to one igniter only in the No.4 burner can when the enginestart levers are in the idle position. Power source for the No.4 plug isfrom the AC standby bus. This position is used to improve igniterservice life and to minimize the possibilities of engine flameouts duringtakeoff, landing, turbulence, and during flight operations in icing condi-tions.

• The OVRD position of the engine start panel provides high energyignition to both igniters when the engine start levers are placed to theidle position. This pulsating high energy power is applied to plugs inboth the No.4 and No.7 burner cans. Power source for the No.4 plug isfrom the AC standby bus. Power source for the No.7 plug is from theAC transfer bus.



B737-200

Leading Edge Devices Indicators

• Shown above are the three types of leading edge device indicator pan-els installed on the various models of the B737. The purpose of theleading edge device indicators (amber & green lights) is to provide avisual indication of the position of the individual leading edge flaps and/or slats. The system consists of two panels of annunciator indicatorlights, one located on the forward flight instrument panel and the otheron the aft overhead panel. The aft overhead indicator panel consists ofone amber and one green light for each leading edge slat and/or flap.



B737-200

• Only one light can illuminate at any one time for each slat and/or flap.The lights are controlled by two extend proximity sensors for the exten-sion function and one reed switch for the retraction indications for eachslat and/or flap. The leading edge annunciator control unit is located inthe E/E compartment. The power source for the control unit is from theNo.1 28V DC bus and all lights are dimmable.

• The individual amber annunciator light for each leading edge device(LED) will illuminate anytime the LED leaves the retract position. Whenthe LED reaches the extend position, the green annunciator light willilluminate and the amber light will be extinguished.

• The forward panel lights or the aft overhead leading edge annunciatorpanel may be inoperative for dispatch. If the forward panel lights areinoperative, the aft overhead panel annunciator must be used to verifythe proper position of the leading edge devices. MEL procedures re-quire a special placard to be installed indicating the proper device posi-tions for flap configurations used. Refer to your MEL.

• Indications for one leading edge slat may be inoperative on both theoverhead panel and the forward annunciator panel for dispatch providednormal LED operations are verified by the flight crew before each take-off/landing, maximum speed is limited to 300 kts. below FL200 or .65Mach above FL200, the overhead panel operates normally.

Leading Edge Annunciator Panel (100)

• Refer to panel 1071, for flap positions 1 thru 25, slats 1,2,5,6, willindicate in the EXT position, all other devices will be extended. For flappositions 30 and 40, slats 1,2,5,6 will indicate FULL EXT position, allothers devices will be extended. When the leading edge devices are inthe full retracted position, the annunciator lights will be extinguished.

Leading Edge Annunciator Panel (200 Basic)

• Refer to panel 1072, for flap positions 1 thru 25, slats 1 and 6 willindicate in the EXT position, all other devices will be extended. For flappositions 30 and 40, slats 1 and 6 will indicate FULL EXT position, allother devices will be extended. When the leading edge devices are inthe full retracted position, the annunciator lights will be extinguished.



B737-200

Leading Edge Annunciator Panel (200A/300/400)

• Refer to panel 1073, for flap positions 1 thru 5, all slats will indicate inthe EXT position and flaps will indicate extended. For flap positions 10thru 40, all slats will indicate FULL EXT position, and flaps will indicateextended. When the leading edge devices are in the full retractedposition, the annunciator lights will be extinguished.

L.E.D. Amber Transit Lights

• The amber leading edge devices transit lights are located on the aftoverhead LED annunciator panel. The illumination of the amber lightindicates the corresponding leading edge device is in transit and/or is indisagreement between the position of any leading edge flap/slat and thetrailing edge flap position. This may also indicate an asymmetrical and/or no leading edge device condition.

L.E.D. Green EXT/FULL EXT Lights

• The green leading edge device EXT/FULL EXT are located on the aftoverhead LED annunciator panel. The illumination of the green lightindicates the corresponding leading edge device are in the EXT positionor the FULL EXT position.



B737-200

L.E.D. Annunciator Panel Test Switch

• The L.E.D. annunciator panel test switch is located on the overheadL.E.D. annunciator panel. By pressing the test button, a system test ofall annunciator panel lights is completed.

• (Simulator Hint) During simulator training, anytime you have a L.E.D.malfunction, always perform a L.E.D. panel test. This is a very fast wayto determine if the simulator instructor has failed a slat and/or flap.Most simulators are programmed to show the No.1 slat failure, sobesure to check that light first.

(Forward Center Instrument Panel)

LE FLAPS TRANSIT Light

• The amber LE FLAPS TRANSIT Light is located on the forward centerinstrument panel. The illumination of this light indicates one or all of theLEDs are in transit. It may also indicate the LEDs are not in theprogrammed position with respect to the trailing edge flaps. With theillumination of the LE flaps transit light, one of the following non-normalconditions may exist:1. Asymmetrical condition.2. No leading edge device condition.3. L.E.D. extended with flaps up.



B737-200

• (Oral Topic) Many simulator instructors will ask the pilot trainee thegeneral procedures for asymmetrical, no leading edge, and/or L.E.D.extended with flaps up conditions. The reason for this inquiry is that themalfunctions and the checklist titles are very confusing and misleading.For asymmetrical or no leading edge malfunctions, call for the ASYM-METRICAL OR NO LEADING EDGE DEVICES checklist. This condi-tion is normally seen during the approach phase of flight in the simula-tor. This checklist requires the planning for a flaps 15 landing withairspeeds at Vref 15+5. Also, limit bank angles to 150 below 210 kts.For leading edge devices extended with the flaps in the up position, callfor the LEADING EDGE FLAPS TRANSIT checklist. This condition isnormally seen during the departure phase of flight in the simulator. Thischecklist requires airspeed to be limited to a maximum of 230 kts withmultiple leading edge devices extended with the trailing edge flaps in theup position. With only one leading edge device extended, limit airspeedto a maximum of 300 kts or .65M (whichever is lower). During flapextension, accomplish the ASYMMETRICAL OR NO LEADING EDGEDEVICES checklist.

LE FLAPS EXT Light (200 Basic)

• The green LE FLAPS EXT light is located on the forward center instru-ment panel. With the flap lever in the positions 1 through 15, the illumi-nation of this green light indicates all leading edge flaps are extendedand all leading edge slats except 1 and 6 are extended. Slats 1 and 6will remain in the intermediate position until flap position 25 has beenselected, at which time, slats 1 and 6 will extend fully.

LE FLAPS EXT Light (200A)

• The green LE FLAPS EXT light is located on the forward center instru-ment panel. With the flap lever in the positions 1,2, or 5, the illuminationof this green light indicates all leading edge flaps are extended and allleading edge slats are in the intermediate position. With the flap lever inposition 10 through 40, all leading devices will be fully extended.



B737-200

Leading Edge Devices

• The leading edge of the wing consists of high lift devices known asKrueger flaps and leading edge slats. Two Krueger flaps are installedinboard of each engine. These flaps are driven by a two-position hy-draulic actuator. Visual displays are provided to the flight crew, indicat-ing either the retract or extend positions. Located outside of the en-gines are the leading edge slats. These slats are driven by a three-position hydraulic actuator. The hydraulic pressure for both types ofactuators is provided by A System hydraulics. A System hydraulics ispowered by both engine driven hydraulic pumps. Visual displays areprovided to the flight crew, indicating the retract, intermediate, and thefull extend positions.



B737-200

• The L.E.D.s are controlled by the leading edge control valve, which ispositioned by the trailing edge drive unit so that the L.E.D.s operatetogether with the trailing edge flaps. When the trailing edge flaps leavethe UP position, the leading edge flaps extend fully while the leadingedge slats extend to an intermediate position. As the trailing edge flapsextend past the Flaps 5 position, the leading edge slats will move to thefull extended position. The sequence is reversed when the flaps areretracted. Pressure to each hydraulic actuator is provided by a commonhydraulic bus. This has been designed to help prevent any asymmetri-cal problems that might occur.

• In the event of a System A hydraulic failure, the leading edge flaps andslats are extended by the standby hydraulic system using alternatehydraulic lines to each drive unit. Upon activation of the alternate flapsmaster switch (extend flaps), all of the leading edge devices will extend.The extension process may take as long as one minute, therefore ap-proach planning is important. The leading edge devices cannot beretracted by the standby hydraulic system.

Altitude Alert Speaker

• Located on the aft overhead panel is the altitude alert speaker. Thisspeaker provides aural altitude alert tones and works in conjunction withthe altitude alert light and the altitude alert control.



B737-200

Thrust Reverser Isolation Valve Panel

• The thrust reverser isolation valve panel is located on the aft overheadpanel. The panel consists of two override switches, one for each re-spective engine system, and one amber isolation valve warning light forthe entire thrust reverser system.

• (Oral Topic) The override switch is a two position switch, labeled NOR-MAL and OVERRIDE. The override switch is normally placed in theNORMAL position. This allows normal thrust reverser operations whencertain conditions are satisfied. Those conditions are:1. Engine oil pressure is more than 35 psi.2. The fire switch is down.3. Air/ground safety sensor is operating in the ground mode.4. Hydraulic pressure is available (System A or standby hydraulics).

• The OVERRIDE position of the override switch provides the means ofbypassing the engine oil pressure switch and the air/ground safety sen-sor. This opens the isolation valve, directing available hydraulic pres-sure to the thrust reverser selector valve. The override position shouldnot be used during normal ground or flight operations.

• The thrust reverser override switches may be inoperative for dispatchfor the respective reverser. Refer to your MEL.



B737-200

Isolation Valve Light

• The illumination of the amber ISOLATION VALVE light indicates thethrust reverser system comparator has sensed a disagreement betweenthe system electrical condition to either isolation valve and the availablehydraulic pressure. During ground operations, the illumination may indi-cate hydraulic pressure is not available to one or both thrust reverserselector valves. The isolation valve will be in the closed position andreverse thrust may not be available. During flight operations, the illumi-nation of the light indicates hydraulic pressure is available to either orboth thrust reverser selector valves. The isolation valve will be in theopen position and protection against thrust reversal may have been lost.Do not actuate the thrust reversers inflight.

Thrust Reverser

• (Oral Topic) Thrust reverse action is accomplished by hydraulic, electri-cal, and mechanical controls. Hydraulic pressure is provided by SystemA hydraulics via the landing gear down-line. The standby hydraulicsystem may also be used as an alternate source of hydraulic pressurein case of System A pressure failure. The standby system has incorpo-rated within the pressure lines hydraulic fuses. This will prevent thetotal loss of the standby hydraulic system due to a rupture in the thrustreverser pressure lines. Limitations prohibit the intentional use of re-verse thrust during in-flight operations.



B737-200

• Electrical control of the thrust reverser system is directed at controllingthe isolation valve. The isolation valve will be energized opened allow-ing hydraulic pressure to extend or retract the reverser when threeconditions are satisfied. Those conditions are the fire handle switchmust be down, engine oil pressure above 35 psi, and the aircraft is inthe ground mode. The failure of any one item will cause the spring-loaded isolation valve to close. This failure will be indicated by theillumination of the ISOLATION VALVE light. Electrical power source forthe thrust reverser system is supplied by the battery bus.



B737-200

• The selection of the reverse thrust levers allows hydraulic pressure, viathe selector valve, to the lock actuators and then to the door actuators.Return hydraulic fluid flows from the actuators through the selectorvalve. A manual lockout plunger has been provided as part of the thrustreverser isolation valve assembly to mechanically prevent any operationof the reverser.

• The one thrust reverser may be inoperative for dispatch provided it issecured closed and if the override system is used, is armed only afterlanding. Your maintenance department will secure the reverser bylockwiring the thrust reverser handle, installing a ground lock on theisolation valve, and collaring the circuit breaker. Refer to your MEL.

• (Opertaional Note) There is a potential for thrust reverser contactingthe ground during landing. Information from Boeing indicates that areverser in transit will contact the ground at a pitch attitude of 6 de-grees. A normal Flap 30 landing touchdown angle is between 4 and 6degrees. Therefore, use caution and avoid “holding it off” in an attemptto get a smooth touchdown. High deck angles during landing increasesthe risk of a reverser strike.

Alternate Reverser Hydraulic Pressure

• Some early aircraft are equipped with an alternate hydraulic pressureaccumulator that will operate the reversers in case of main systemfailure. These accumulators are located in the wing-body fairing andprovides pressure through the thrust reverser shuttle valves.



B737-200

Service Interphone System

• The service interphone system provides communication between theflight deck, cabin attendants, and ground personnel. The flight crewmay communicate by using the handset or the audio selector panelcontrols. The cabin attendants may communicate by using either theforward or aft cabin handsets. Ground personnel may communicate atany one of the seven interphone stations. Those stations are located atthe external power panel, APU access panel, wing refueling station,FWD right wheel well, FWD left wheel well, aft passenger cabin ceiling,and at the electronic equipment rack area.

• The flight deck to cabin, to flight deck, or the cabin to cabin serviceinterphone system may be inoperative for dispatch providing alternatecommunications procedures are established. These alternate proce-dures must include communication procedures for normal and emer-gency situations. In addition, the PA system must operate normally.The visual alerting portion of the interphone system may also be inop-erative for dispatch provided the PA system operates normally.



B737-200

Service Interphone Switch (as installed)

• The service interphone switch is located on the aft overhead panel. Theselection of the two-position switch to the ON position, connects theexternal jacks to the service interphone system. When selected to theOFF position, the external jacks are deactivated. The OFF position hasno effect on internal communications on the flight deck or cabin areas.

External Power Interphone Station

• Connects the ground crew to the flight interphone system. The serviceinterphone jack connects the ground crew to the service interphonesystem if the service interphone switch has been selected ON.

Flight Interphone System

• The flight interphone system provides private and independent commu-nications between cockpit crewmembers without intrusion from the ser-vice interphone system.



B737-200

White Dome Light Switch (as installed)

• The three-position white dome switch is located on the aft overheadpanel. The switch controls the two overhead white lights. With the lossof all AC power, the white dome lights will be powered by the batterybus.

• (Oral Topic) With the loss of all AC power, the following lights will beavailable (powered by the battery bus):1. Standby compass light.2. White dome lights.3. Emergency flood lights.4. Various system-warning lights.

• (Oral Topic)) With the battery switch selected OFF and external powerconnected, the following lights will be available (powered by the hotbattery bus):1. Dim entry lights.2. Fluorescent mirror lights in the lavatories.



B737-200

White/Red Dome Light Switches (as installed)

• The two-position white dome switch is located on the aft overheadpanel. The ON/OFF switch controls the two overhead white lights. Thered dome light controller controls the variable intensity red dome lightsoverhead on the sidewalls.

Audio Selector Panel #1Audio Selector Panel (overview)

• The audio selector panels (ASP) installed on the B737 consists of sev-eral different designs. We have chosen two of the most prevalent typesinstalled on the B737-200. Basic functions of each type of ASP are thesame, they serve as independent communication control panels for indi-vidual crewmembers or for the flight deck observer. Each crewmemberhas the capability to select and control required radio, navigation,interphones, and PA functions. A transmitter selector is located oneach ASP for individual crewmember use. Transmissions can be madeby using the boom microphone or the oxygen mask. The ASP has noaffect or control on the functions of the GPWS system, altitude alertsystem, and/or the windshear alert system. The audio selector panelmay be inoperative for dispatch provided only the inoperative panels arein excess of those required for flight deck crewmembers.



B737-200

ASP - Receiver Switch (Panel #1))

• The individual receiver switch allows the reception of the respectivecommunication system or navigation receiver. Multiple switches may beselected to the UP position at any time.

ASP - Transmit Light (Panel #1))

• The green transmit light is located just above the receiver switch. Theillumination of this light indicates the transmitter selector has been se-lected to this position, thus allowing transmission to the selected radioor PA.

ASP - Transmitter Selector (Panel #1))

• The transmitter selector is located directly under the receiver switches.The selector may be rotated to the respective communication systemfor subsequent transmission. For PA selection, selector should bepulled up. Reception is also possible over the selected system regard-less of whether the associated receiver switch is ON. Located withinthe center of the selector, is the volume control. Rotating the controlknob adjusts volume of all receivers.

ASP - Filter Switch (Panel #1))

• The three-position filter switch is located to the right of the transmitterselector. The switch controls the audio reception from the VHF, NAVand ADF radios. Selected to the VOICE position, the ASP will receivevoice audio only. Selected to the RANGE position, the ASP receivesstation identifier audio only. Selected to the BOTH position, the ASPreceives both voice and range audio.

ASP - PTT Switch (Panel #1))

• The PTT switch (Push To Talk) is located to the left side of the transmit-ter selector. Selecting the PTT switch, keys the oxygen mask or boommicrophone for transmission.



B737-200

ASP - Oxygen/Boom Switch (Panel #1)

• The two-position oxygen/boom switch is located to the left of the PTTswitch. Selection of the OXYGEN position provides oxygen mask trans-missions capabilities. The selection of the BOOM position providesboom microphone transmission capabilities. The PTT switch on thepilot’s control wheel can also be used to transmit when using the oxygenor boom microphone.

Audio Selector Panel #2ASP - Receiver Switch (Panel #2)

• Pressing DOWN the individual receiver light-switch allows the receptionof the respective communication system or navigation receiver and theillumination of the internal light. Pressing the switch a second time,deselects the receiver and the switch returns to the UP extinguishedposition. The switch may be rotated to control the volume of the se-lected receiver. Multiple switches may be selected at any one time.



B737-200

ASP - Transmitter Selector (Panel #2)

• The transmitter selector is located on top of the receiver switches.Pressing DOWN with the individual transmitter light-selector allows thetransmission of the respective communication system or navigation re-ceiver and the illumination of the internal light. Only one selector switchmay be depressed at any one time. Depressing any other switch,deselects the first switch. The switch-selector labeled INT is used withthe CAB/SERV INT receiver switch and/or the I/C receiver switch.Interphone transmission requires both INT and the receiver switches tobe selected to the ON position.

ASP - Amplifier Switch (Panel #2)

• The amplifier switch is located on the bottom left corner of the ASPpanel. It is used to select the desired amplifier for the ASP.

ASP - ALT/NORM Switch (Panel #2)

• The two-position ALT/NORM switch is located on the bottom left cornerof the ASP panel. When selected to the ALT position, the ASP willoperate in a degraded mode. Selecting the NORM position, the ASPwill operate normally.

ASP - PTT Switch (Panel #2)

• The three-position PTT switch (Push To Talk) is spring loaded to theneutral position. Selecting the PTT switch to the R/T position, keys theoxygen mask or boom microphone for transmission as selected by thetransmitter selector. The selection of the I/C position, keys the oxygenmask or boom microphone for transmission over the flight interphonesystem. This position bypasses the transmitter selector controls.

ASP - MASK/BOOM Switch (Panel #2)

• The two-position mask-boom switch selects the oxygen mask or theboom microphone for communications.



B737-200

Flight Crew Oxygen Indicator

• The flight crew oxygen indicator displays pressure values as sensed atthe crew oxygen cylinder. The flight crew oxygen system is completelyindependent from the passenger oxygen system. The single-bottle sys-tem is normally charged to a pressure value of 1850 psi.

• Minimum dispatch pressures are 1100 psi, but may vary depending onthe size of the cylinder installed and the duration of the flight. Refer toyour operations manual for minimum pressures. The crew indicatormay be inoperative for dispatch provided an alternate procedure is usedto verify that the oxygen supply is above the minimum dispatch levels.Refer to your MEL.

• A common oral question asked by many examiners is referenced to therequirement of having the battery switch either ON or OFF when read-ing the indicator during preflight. The battery switch must be in the ONposition to read the flight crew oxygen indicator.



B737-200

Passenger Oxygen Indicator

• The passenger oxygen indicator displays pressure values as sensed atthe passenger oxygen cylinder. The passenger oxygen system is com-pletely independent from the flight crew oxygen system and is of acontinuous-flow design. The dual-bottle system is normally charged toa pressure value of 1850 psi. Minimum dispatch pressures are 1100psi, but may vary depending on the size of the cylinders installed andthe duration of the flight.

• A common oral question asked by many examiners is referenced to thethree ways of activating the passenger oxygen system. The system canbe activated by:1. Automatic activation of the system when the cabin altitude reaches

14,000’ (via the pneumatic continuous flow control unit).2. Activation of the passenger oxygen switch on the aft overhead

panel (electro-pneumatic continuous flow unit).3. Manual activation of the cockpit floor-mounted manual-reset activa-

tion handle.

• Passenger oxygen is provided by oxygen masks that will drop from thepassenger service unit (PSU). Each PSU may contain 3 or more mask,depending upon configuration installed. The flow of oxygen to the pas-senger mask is initiated when the mask is pulled down, causing theactuator pin to be withdrawn from the unit.

• A common oral question refers to the flow rate and the oxygen dilutionlevels. Oxygen flow is at a constant rate and is diluted by cabin air invariation with cabin altitude. Because of this diluted oxygen supply, DONOT use the passenger oxygen system when smoke or fire is present.When the cabin altitude is below 14,000’, the entire passenger systemmay be shut-off by using the manual-reset handle located in the cockpit.An individual oxygen mask flow can be terminated at the PSU by rein-stalling the actuator pin.



B737-200

• The passenger indicator may be inoperative for dispatch provided analternate procedure is used to verify the oxygen supply is above theminimum dispatch levels. The entire passenger system may be inop-erative for dispatch provided the flight is not conducted where the mini-mum enroute altitude is above 14,000’ MSL, both A/C packs are operat-ing normally, the pressurization system is operating normally, the flightis conducted at or below FL250, portable oxygen units are provided for10% of the passengers, and all passengers are appropriately briefed bythe crew. Refer to your MEL.

Passenger Oxygen Indicator Light

• The passenger oxygen indicator light is located within the oxygen con-trol panel on the aft overhead panel. The illumination of this ambercaution light indicates oxygen pressure is being sensed within the pas-senger oxygen system following system activation.

Passenger Oxygen Switch

• The two-position passenger oxygen switch is located under the passen-ger and crew oxygen indicators. The passenger oxygen switch is alwaysin the guarded NORM position unless the switch has been manuallyselected to the ON position. The NORMAL position provides for theautomatic activation of the passenger oxygen system should the cabinaltitude exceed 14,000’. This automatic activation will cause the oxygenmask to drop from all cabin PSUs. Manually selecting the ON positionof the passenger oxygen switch will also activate the oxygen systemand drop all mask. This function should be used if the automatic func-tion fails to lower mask.

• The automatic presentation or function of the passenger oxygen deploy-ment system, as listed under the PSU classification, may be inoperativefor dispatch provided the manual deployment function of the system isoperating normally and the flight remains at or below FL 300. Refer toyour MEL.



B737-200

Manual Actuation & Reset Handle

• The manual actuation & reset handle is located on the floor of the flightdeck. Opening the floor door allows access to the handle. The three-position reset handle allows the manual activation or manual resetting ofthe passenger oxygen system. Moving the handle to the PULL ONposition will activate the oxygen system. Pushing the handle in for 5seconds, closes the oxygen flow control valves and resets the systemwhen cabin altitude is below 14,000’.



B737-200

Flight Crew OxygenMasks (as installed)

•The flight crew oxygenmask has been designed toincorporate a regulator-typesystem that provides oxygeneither in a diluted or 100%O

2 format. The bottom por-

tion of the mask contains theregulator control functions.The crew member has theoption of selecting NORMALor 100% oxygen. An EMER-GENCY control knob hasalso been provided thatchanges the flow of the regu-lator from diluter demand tosteady flow when the knobhas been rotated.

• The mask also contains a self-test feature. When pressing together theRESET-TEST lever and PRESS TO TEST knob, oxygen flow is allowedinto the mask. Flow can also be checked by the flow indicator. Theentire oxygen mask/regulator is stored in a metal-box which is mountedon the sidewall next to the pilot seat. The oxygen mask container hasbeen designed with a shutoff valve mechanism that prevents unwantedoxygen flow inside the box. This mechanism is designed with the doorclosing latch.

Flight Crew Oxygen Shutoff Valve

• The flight crew oxygen shutoff valve is located on the right cockpitbulkhead, behind the first officer’s seat. Turning the knob counterclock-wise allows oxygen to flow to each flight deck station. Turning the knobclockwise shuts off oxygen flow.



B737-200

Flight Crewmember Oxygen Panel

• The flight crewmember oxygen panel (FCOP) is located at each pilotand observer station. The panel consists of emergency lever, flow indi-cator, supply lever, and oxygen diluter lever.

(FCOP) Emergency Lever

• The two-position emergency level is located on the left side of the flightcrewmember oxygen panel. The ON position supplies 100% oxygenunder positive pressure to the crewmember’s oxygen mask The OFFposition selects off the emergency lever control, the oxygen and airmixture is now controlled by the oxygen diluter lever.

(FCOP) Oxygen Diluter Lever

• The two-position oxygen diluter lever is located on the center section ofthe flight crewmember oxygen panel. The 100% position provides pureoxygen on demand to the crewmember. The NORMAL position pro-vides a variable mixture of oxygen and cabin air to the crewmemberbased on cabin altitude.



B737-200

(FCOP) Oxygen Supply Lever

• The two-position OFF/ON oxygen supply lever controls the oxygen sup-ply to the regulator.

(FCOP) Oxygen Flow Indicator

• The oxygen flow indicator indicates oxygen flow through the regulator tothe mask.

Portable Passenger Oxygen (as installed)

• There are 4 cabin oxygen cylinders normally installed on the B737. Thecylinders are fitted with a pressure gauge, pressure regulator, and anON/OFF control valve. The cylinders are pressurized to 1800 psi, witha capacity of 4.25 cubic feet (120 liters) of oxygen. Each bottle has twocontinuous flow outlets. One outlet provides a flow rate of two liters perminute, and the other outlet provides a flow rate of four liters per minute.The four liters per minute outlet has been designated for first aid pur-poses. Duration can be determined by dividing 120 liters by 4 liter/mn,therefore, the bottle provides 30 minutes of normal use. Portable oxy-gen bottles (POB) can be used for therapeutic purposes or as walk-around units. Do not allow the POBs pressure drop below 500 psi. Thereason for the 500 limitation, is to meet FAA 15 minute flight attendant“walk-thru” requirements.



B737-200

Portable Crew Oxygen (as installed)

• The flight crew portable oxygen system consists of a oxygen cylinder,pressure regulator, ON/OFF valve, and a pressure gauge. The unitinstalled has been designed as a portable full face and respiratory pro-tection system. The cylinder is charged to 1800 psi, and contains 311liters of oxygen. Duration of this cylinder is approximately 103 minutesusing the 3 liter constant flow outlet. The portable system provides thecapabilities of offering both demand and constant flow oxygen. Theregulator is of a demand design, with a full-face mask attached to it. Asecond fitting provides constant flow oxygen.

• Four portable oxygen dispensing units (mask and bottle) are normallyinstalled, two units are required for dispatch. Any in excess of thoserequired by FAR regulations may be missing or unserviceable providedthe proper distribution of the remaining bottles is maintained throughoutthe aircraft. The bottles that are not properly serviced must be replacedor serviced at the next available maintenance facility. Refer to yourMEL.

Personal Breathing Equipment (PBE) Smoke Hoods (as installed)

• Five PBEs are normally installed throughout the aircraft. Four units arerequired for dispatch. Any in excess of those required by FAR regula-tions may be inoperative.



B737-200

(FDR) Flight Recorder Test Switch

• The flight recorder test switch is a two-position guarded switch. TheNORMAL position of the switch has two functions of operations. Duringground operations, the FDR will only operate when power is being sup-plied and one of the two 35 psi oil switches has closed. During flightoperations, the FDR will operate anytime power is being supplied to theunit (regardless of the condition of the oil pressure switch).

• The TEST position of the flight recorder test switch has been providedto enable the unit to bypass the engine oil pressure switches and the airground switch to power the flight recorder on the ground. Electricalpower must be available to the flight data recorder for this TEST func-tion to occur. A valid test is indicated by the flight recorder OFF Lightextinguishing and the illumination of the trip and date light.



B737-200

Flight Data Recorder (as installed)

• There are several different types of Flight Data Recorders (FDR) avail-able for the B737. Only the two most common types will be discussedin this review. The flight recorder has been designed to provide apermanent tape record of various system and operational technicaldata. The unit is sealed in a fire-resistant container to safeguard thisdata. Technical data that is recorded includes airspeed, altitude, head-ing, vertical acceleration, and elapsed time. The flight recorder is nor-mally located behind an access door in the aft cabin ceiling, just forwardof the aft pressure bulkhead.

• The flight data recorder provides a continuous 25 hour record of theaircraft’s parameters. Recording occurs anytime the unit is being pow-ered. Electrical power for the unit is provided by the No.1 transfer busand the battery bus. During ground operations, the recording processbegins when engine No.1 or engine No.2 35 psi oil pressure switchcloses. This function of the oil switch is bypassed during flight opera-tions. The flight data recorder will remain operating and recording aslong as power is being provided to the unit. This includes the casescenario of a dual engine flameout.

• The flight data recorder may be inoperative for dispatch provided thecockpit recorder is operating normally and repairs to the unit is madewithin three flight days. Refer to your MEL.

• FAR regulations states the flight data will not be used in determiningany certificate action or civil penalty, arising out of an accident or occur-rence. The flight data recorder will only be used in determining thecauses of accidents and occurrences under investigation by the NTSB.



B737-200

(FDR) Trip and Date Selectors

• The trip and date selectors are located to the left center of the FDRpanel. By rotating the selector, the trip’s date and number can beinserted and recorded on the tape.

(FDR) Recording Time Remaining Indicator

• The recording time remaining indicator is located to the right center ofthe FDR panel. This portion of the recorder is not normally used sincethe FDR provides a continuous 25 hour record of the aircraft’s param-eters. Any indication noted on the indicator, will be shown as valvesabove zero.

(FDR) Event Switch

• The event switch-button is located to the lower right side of the flightdata recorder panel. Pressing the switch-button will transcribe a markon the tape to identify the time of an event. Do not press the switch-button until after the trip and date light is extinguished.



B737-200

(FDR) Trip and Date Light

• The amber trip and date light is located directly below the recording timeremaining indicator. The illumination of this light indicates trip and dateinformation is being recorded. The 15 minute transcribing cycle doesnot interfere with the recording of other information.

(FDR) Repeat Switch

• The repeat switch-button is located to the left side of the amber trip anddate light. Pressing the button-switch will initiate the transcribing pro-cess of the trip and date data.

(FDR) OFF Light

• The illumination of the amber OFF Light indicates the recorder is notoperating or the test is invalid. The illumination may also indicate powerfailure, loss of input data, or a electronics malfunction.

(FDR) Documentary Data Thumbwheel Switches

• The documentary data thumbwheel switches are located on the topcenter of the FDR panel. The thumbwheel switches are rotated to enterdate and flight identification data. The information concerning day,month, flight number, and leg identification can be inserted for record-ing.



B737-200

Mach Airspeed Warning Test Switch

• The mach airspeed warning test switch is located on the aft overheadpanel. Pressing the switch-button performs a test of the wire continuityof the aural clacker warning system. A valid test is indicated by thesounding of the aural clacker warning.

• The mach airspeed warning system is independent from other flightinstrument systems, it provides an aural warning anytime the maximumoperating speed is exceeded. The aural warning clacker can only besilenced by reducing airspeed. Inputs are received from the No.1 auxil-iary pitot-static system. See pitot-static system schematic for reference.

• The mach airspeed warning system may be inoperative for dispatchprovided both mach indicators are operating normally. If the overspeedclacker occurs earlier than programed, aircraft speed must remain be-low the point at which the clacker sounds. If the overspeed clackeroccurs below M.78, the system must be deactivated by pulling therespective circuit breaker and observe all speed limitations. If the sys-tem is completely inoperative, the following speed limitations apply:

1. Mmo - .78 above FL 230.2. Vmo - 340 LIAS below FL 230. (Refer to your MEL).



B737-200

Stall Warning Off Light

• (Oral Topic) The illumination of the amber stall warning off light indi-cates one of the following failures has occurred: heater failure of theangle airflow sensor, system signal failure, and/or electrical power fail-ure to the entire system. To determine which component of the stallwarning system is inoperative, simply perform the TEST. If the lightremains illuminated, then the heating component has failed. If the lightextinguishes itself and the indicator (spinner) fails to rotate, then thestall warning system continuity is inoperative.

Stall Warning Switch

• The three-position stall warning switch is located to the right side of thestall warning panel. The switch is normally selected to the NORMALposition. With the switch in the NORMAL position, electrical power isavailable for the internal heater of the angle airflow sensor.

• (Oral Topic) The internal heating feature is only available when theNo.1 engine is operating and/or when the air ground safety sensor is inthe air mode.

• (Oral Topic) The TEST position of the stall warning switch has beendesigned to test the system with the No.1 engine operating or notoperating. The visual test indications for each condition is a commonarea of interest of examiners during orals. With the No.1 engine notoperating, movement of the test switch to the TEST position will causethe OFF light to extinguish, the rotation of the test indicator, and thevibration of the control columns. The OFF light will illuminate againafter the test has been completed. Performing the test with the No.1engine operating, will cause the rotation of the test indicator and thevibration of the control columns. The OFF light will remain extinguishedduring and after the test.

• The movement of the stall warning switch to the HTR OFF locked toggleposition, removes electrical power from the angle airflow sensor. Thisposition of the stall warning switch is normally used for various types ofmaintenance checks and is not normally used by flight crews.



B737-200

Stall Warning Test Indicator

• The black and white test indicator disc has been provided to give avisual indication of the electrical continuity through the angle airflowsensor and flap position transmitter during system testing.

Stall Warning System

• The stall warning system has been designed to provide the flight crewadvance notice of an impending stall. Design regulations requires thisadvance warning to occur at a minimum of seven percent above theactual stall speed. This artificial warning is provided by the controlcolumn shaker (stick shaker). At seven percent above stall, the stickshaker alerts the crew of the impending stall.

• The design of the B737 also provides a “natural buffet” warning to occurjust prior to the actual stall. The stall warning system componentsconsist of a control column shaker, heated angle of airflow sensor, flapposition sensor, stall warning amplifier, air/ground safety sensor, andthe stall warning panel. The stall warning system is deactivated on theground by the air/ground safety sensor. The stall warning system mustbe operative for dispatch. Refer to your MEL.



B737-200

Master Fire Warning Lights

• The red master fire warning switch-light is located on either side of thecenter lightshield panel. The illumination of this warning light indicates ared warning light on the fire protection panel (center console) has illumi-nated. This may be caused by the testing of the fire protection systemor an actual fire has been detected. In addition to this visual warning ofa possible fire, the aural indication of the fire alarm bell sounding willoccur. Should the aircraft be on the ground when this occurs, theremote APU horn will also sound. These lights will remain illuminatedas long as the situation is present.

• The master fire warning switch-light also functions as a switch. Press-ing the light unit, extinguishes both fire warning lights, silences bothalarm bells (fire and APU), and resets the system for additional warn-ings.

• (Oral Topic) The significance of the red warning light is to stress theimportance of the situation and that the situation requires immediatecorrective action by the flight crew. Many examiners may ask theapplicant to describe the major systems of the aircraft that have redwarning lights associated with them. Those systems include the engine,APU, landing gear, wheel well, and the autopilot disconnect.



B737-200

Master Caution Lights

• The amber master caution light is located on either side of the centerlightshield panel. The illumination of this caution light provides a visualcue that a “system annunciator” light has also illuminated somewhereoutside the normal field of vision of the pilot. Both caution lights willremain illuminated as long as the non-normal situation remains. Press-ing either light/switch, extinguishes both master caution lights, extin-guishes the respective system annunciator light, and resets the mastercaution system for further non-normal conditions.

• (Oral Topic) A commonly asked subject by examiners concerning themaster caution system, refers to single system failures and the illumina-tion of the amber master caution light. Most single system failures donot illuminate the amber master caution lights, but are stored within thecaution alerting system for pilot recall. Pressing either system annun-ciator panel, recalls the single failure (and any other faults that mayexist), and displays the respective fault on the annunciator panel.

• The master caution warning system must be operative for dispatch.The push-to-test function of the light/switch may be inoperative providedthe intended function of the caution warning system has been verifiedoperational prior to dispatch. Refer to your MEL.



B737-200

System Annunciator Panel

• The amber system annunciator panel is located on either side of thecenter lightshield panel. The illumination of any amber caution light onthis panel provides a visual cue that a “non-normal condition” existssomewhere outside the normal field of vision of the pilot. The systemannunciator provides warning only for aircraft systems that are locatedon the overhead, aft overhead, and on the lower fire protection panel.To extinguish any system light, press either annunciator panel. Press-ing the system annunciator panel a second time, recalls the respectivefault on the annunciator panel.

• (Oral Topic) Some examiners have inquired during orals the itemsassociated with the illumination of a system annunciator. Therefore, wehave provided the following section to assist the pilot applicant on recall-ing those items.

• The illumination of the amber FLT CONT system annunciator light indi-cates one of the following flight control caution lights has illuminated.

Low Quantity Low PressureYaw Damper Feel Diff PressMach Trim

• The illumination of the amber FUEL system annunciator light indicatesone of the following fuel system caution lights has illuminated.

Low PressureFilter Bypass

• The illumination of the amber ELEC system annunciator light indicatesone of the following electrical system caution lights has illuminated.

Low Oil Pressure High Oil TemperatureStandby Power OFF Transfer Bus OFFBus OFF



B737-200

• The illumination of the amber APU system annunciator light indicatesone of the following auxiliary power unit system caution lights has illumi-nated.

Low Oil PressureHigh Oil Temperature/FaultOverspeed

• The illumination of the amber OVHT/DET system annunciator light indi-cates one of the following overheat/fire detection system caution lightshas illuminated.

Engine No.1 Overheat APU DET InoperativeEngine No.2 Overheat

• The illumination of the amber ANTI-ICE system annunciator light indi-cates one of the following anti-Ice system caution lights has illuminated.

Window Overheat Pitot Heat OFFCowl Anti-Ice

• The illumination of the amber HYD system annunciator light indicatesone of the following hydraulic system caution lights has illuminated.

OverheatLow Pressure

• The illumination of the amber DOORS system annunciator light indi-cates one of the following aircraft’s doors caution lights has illuminated.

FWD/AFT Entry Tire Screen (as installed)Equip. Compartment FWD/AFT CargoFWD/AFT Service Airstair (as installed)

• The illumination of the amber ENG system annunciator light indicatesone of the following engine system caution lights has illuminated.

Reverser

• The illumination of the amber OVERHEAD system annunciator lightindicates one of the following aircraft’s overhead caution lights has illu-minated.

Equip. Cooling OFF Emer. Exits Not ArmedFlight Recorder OFF Passenger Oxygen ON



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• The illumination of the amber AIR COND system annunciator light indi-cates one of the following air conditioning system caution lights hasilluminated.

Duct Overheat Dual BleedPack Trip OFF Wing-Body OverheatBleed Trip OFF Auto FailOFF Scheduled Descent

Flight Director Switch (SP177)

• The flight director switch is located on either side of the centerlightshield panel. The selection of the switch to the ON position, en-ables command bar display on the respective ADI. Upon initial selectionof the F/D switch, the command bars will not appear unless commandpitch and roll modes are engaged. FCC A provides data to theCaptain’s command bars and FCC B provides data to the First Officer’scommand bars. With both switches ON, logic for both sets of commandbars are controlled by the master FCC. This is indicated by the illumi-nation of the master flight director indicator light.



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• (Oral Topic) A commonly asked subject by examiners, with reference toflight director operation, is the general description of the flight directortakeoff mode. The F/D takeoff mode requires both F/D switches to beselected ON. The F/D takeoff mode is initiated by the selection of theTO/GA button during the initial takeoff roll. The command bars areinitially 100 nose-down and wings level. At approximately 60 kts, thecommand bars moves to 150 nose-up. After wheels up, the commandbars provide guidance to maintain pitch that will achieve MCP speedplus 20 kts. During the takeoff roll and the climb, F/D commands wingslevel. Normally at 400’, takeoff mode is terminated and LVL CHG isselected.

• (Oral Topic) Many examiners are known to ask the opeartion of the F/Dduring an engine failure at takeoff. The F/D has been designed toprovide three modes of pitch guidance during takeoff with engine fail-ures. The first mode: provide pitch guidance that will maintain V2speed. This speed mode is provided should the engine fail prior toreaching V2 speed. The second mode: provide pitch guidance thatmaintains the reference speed at the time at which the failure occurred.This speed will be somewhere between V2 and V2+20. The third mode:provide pitch guidance that maintains V2+20 kts. This speed modeoccurs after obtaining V2+20 or higher speeds.

• Both flight director systems may be inoperative for dispatch providedapproach minimums do not require their use. Many airlines have speci-fied within their flight operations manuals further limitations concerninginoperative flight directors and autopilots. For example, with both flightdirectors and the autopilot(s) inoperative, forecasted visibility must bebetter than 3/4 mile (4000’ RVR). With forecasted visibility of 3/4 mile(4000’ RVR) or less, at least one flight director and one approach cou-pler must be operable in order to be dispatched. One additional note,should both flight directors and/or autopilots become inoperable enroute, the approach may be executed to published minimums. Refer toyour MEL.



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Master Flight Director Indicator Lights (SP177)

• The illumination of the respective master flight director light indicates therespective FCC the controlling the F/D modes and provides the altitudealert reference mode. At least one F/D switch must be selected to theON position, before the MA light can illuminate. With neither A/Pengaged in the CMD mode, the first F/D selected ON, will be the con-trolling master FCC. With one or both A/Ps engaged in CMD, the FCCfor the A/P in CMD is the master FCC (regardless of which F/D switch isturned ON first). The illumination of both MA lights, indicates the re-spective FCC is controlling the F/D modes for the respective flight direc-tor, thus providing independent F/D operation.

Autothrottle Arm Switch (SP177)

• (SP177) The two-position autothrottle arm switch automatically controlsthrust through all phases of flight. The PDC provides thrust lever valuesto the autothrottle system. Selecting the A/T switch to the ARM posi-tion, arms the A/T system for engagement with EPR, SPEED, or PDCSPD modes of operation. The actual switch is held in the ARM positionmagnetically. Thrust lever movement is accomplished by theautothrottle servo motors.



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• (Oral Topic) Many examiners are known to have asked the differentways of autothrottle disengagement. The following are examples ofways to disconnect the A/T.1. A/T Arm switch selection to the OFF position.2. Pressing the A/T disengage switch.3. Automatic disengagement, 2 seconds after landing.4. Asymmetrical thrust lever position (more than 100).5. A/T system fault.

• The autothrottle system provides various modes of operation. Thosemodes of operation are: A/T Takeoff Mode, N1 Mode, Speed Mode, N1Equalization Mode, Arm Mode, Descent Retard Mode, Go-aroundMode, and A/T Mode of Engagement and Transfer. Only those modesof operation that are normally covered in orals will be discussed in thisreview.

• (SP177) The A/T Takeoff Mode is engaged when the A/T arm switch isselected to the ARM position and a PDC takeoff page is engaged.Normally, this procedure is accomplished when the aircraft is cleared onthe active runway. The confirmation of the arming of the A/T system, isthe annunciation of ARM on the flight mode annunciator panel. Theselection of the TO/GA button, initiates thrust lever movement. Theannunciator panels will reflect a change of A/T status from ARM to EPR.The A/T system will automatically set takeoff thrust by 60 KIAS. Theannunciator panel will indicate A/T status of THR HOLD by 64 KIAS.Once THR HOLD has been obtained, only manual changes by the flightcrew can be made to the thrust levers. The THR HOLD function will bemaintained until 400’ RA (approximately 18 seconds after takeoff). At400’ RA, the flight crew may select LVL CHG on the AFDS modecontrol panel. The selection of MCP EPR switch will initiate the climbthrust reduction phase of flight. Refer to your operations manuals fortakeoff procedures with regards to climb reductions and flight profiles.EPR values for the various flight modes are obtained from the PDC.



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• (SP177) The Go-around Mode of the autothrottle system is automati-cally armed when descending below 2000’ RA (autothrottles engaged).The selection of the TO/GA button, initiates the go-around process ofthe autothrottles. The thrust levers will advance to a “REDUCED” go-around thrust setting. This provides a climb rate of approximately 1000’to 2000’ FPM. The flight mode annunciator panel will display GA for A/T. If full go-around thrust is required by the flight crew, pressing theTO/GA button a second time, will advance the thrust levers to a “FULL”go-around thrust setting.

IAS/MACH Display (SP177)

• The IAS/MACH display indicator is located above the speed selector onthe mode control panel. Selected speed is displayed in 1-knot incre-ments beginning at 110 kts. The speed selector sets the desired speedon the display indicator and on the airspeed indicator.

• The IAS/MACH display indicator provides a “blank” display wheneverPDC SPD mode is engaged. The indicator will also be blank during atwo-engine AFDS go-around or when the A/T is engaged in PDC SPDmode.

• (SP177) The IAS/MACH display indicator will also display commandspeed limiting modes. The AFCS provides thrust and speed pitch com-mands that avoids exceeding any preset limit speeds. Limits includesVmo/Mmo limitations, landing gear placard speeds, wing flap placardspeeds, and minimum speeds (1.3 Vs flap configuration).

Change-Over Switch (SP177)

• The change-over switch is located to the lower left of the IAS/MACHdisplay. Selecting the switch, changes the display between IAS andMACH. The switch has also been designed with an automatic change-over function that occurs at approximately FL235.



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Mode Selector Switches (SP177)

• The mode selector switch consists of 4 switch-lights that controls theengagement of EPR, SPEED, LVL CHG, and PDC. The momentaryselection of the switch-light activates the selected mode and illuminatesthe ON light. Pressing the switch-light a second time, deactivates theselected mode.

• An important note concerning the operation of the AFDS, with regardsto the switch-lights selection, is the operational status of the selectedmode. The illumination of the ON light does not verify the operation ofthe selected mode. Only the display of mode condition on the flightmode annunciator panel verifies the status of the selected mode.

Heading/Bank Angle Selector (SP177)

• The heading and bank angle selector is located at the center of theAFDS panel. Rotating the selector, sets the desired heading marker onboth HSIs (EHSIs). The heading selected is displayed on the headingdisplay indicator. The bank angle portion of the selector, sets bankangles of 10, 15, 20, 25, and 30 as desired. These angles are usedduring HDG SEL and VOR modes of operation.



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Altitude Display Indicator (SP177)

• The altitude display indicator is located to the upper-right side of theheading/bank angle selector. The indicator displays altitudes, as se-lected by the flight crew, from 0 to 50,000’ in 100 foot increments. Theprimary purpose of the display indicator is for altitude reference, altitudealerting, and for automatic level-offs. Depending on the unit installed,during first power-up of the aircraft, unmodified display units will showan altitude of 10,000’. Modified units will display previous selectedaltitudes following the initial power-up of the aircraft.

Altitude Selector (SP177)

• The altitude selector is located directly below the altitude display. Therotation of the selector sets the desired altitude in the display indicator.

Vertical Speed Display (SP177)

• The vertical speed display indicator is located to the right side of thealtitude display indicator. The indicator can display vertical speeds from-7900 to 6000 fpm. The indicator will be blank when the vertical speedmode is not active. The selection of the vertical speed thumbwheel setsvertical speed as displayed on the indicator.



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Autopilot Engage Paddles (SP177)

• The autopilot engage paddles are located to the right of the verticalspeed thumbwhell. The paddles are labeled with three modes of en-gagement. With either paddle selected to the OFF position, the respec-tive autopilot is disengaged. Movement of either paddle to the CWSposition, engages autopilot pitch and roll as controlled by pilot inducedmovement of the control wheel. The selection of the paddles to theCMD position, enables all command modes of operation for the AFDS inaddition to the CWS modes. The first A/P paddle selected to the CMDposition, will be the master FCC, regardless of which F/D switch isselected first.

• CWS functions of the autopilot are designed to limit the pilot controlforce if attitude limit is exceeded. For example, attitude input is exces-sive, the autopilot will return to the attitude limits when the control forceis released. When operating in the CWS mode, if roll control force of 60

of bank or less occurs, the autopilot will roll wings level and holds theexisting heading.



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Test Position #1

Test Position #2

Flight Mode Annunciator Panels (as installed)

• The flight mode annunciator panels are located on the forward instru-ment panel (one on either side). The panels provide a visual display ofthe current status of the AFDS, autothrottles, and the PDC. The me-chanical annunciator displays are three-sided prisms and have externalillumination.



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Test Position #1

Test Position #2

Flight Mode Annunciator Panels (as installed)

• The flight mode annunciator panels are located on the forward instru-ment panel (one on either side). The panels provide a visual display ofthe current status of the AFDS, autothrottles, and the PDC. The me-chanical annunciator displays are three-sided prisms and have externalillumination.



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• Panel 1113, reflects the testing of the FMA by TEST mode #1. Thelarge A/P and A/T display lights will be illuminated a steady amber.These lights are internally illuminated light caps. The A/P status displaywill indicate CWS ROLL and CWS PITCH.

• Panel 1114, reflects the testing of the FMA by TEST mode #2. Thelarge A/P and A/T display lights will be illuminated a steady red. The A/P status display will indicate SINGLE CH and A/P OFF.

• Panel 1115 and 1116, displays the available ENGAGED mode items.Each item will be annunicated as black letters on a green background.Note, the A/P STATUS, A/P, and A/T displays are blank.

• Panel 1117, displays the available ARMED mode items. Each item willbe annunicated as white letters on a black background. Note, the A/PSTATUS, A/P, and A/T displays are blank.

• Only one of the two SP-177 flight mode annunciator panels, may beinoperative for dispatch provided the engage system is at the pilot posi-tion with the operating annunciator. Observe any approach minimumslimitations that may apply. There are no other restrictions concerningthe use of other systems (autothrottles, F/D, etc.) as long as the pilotwith the operating annunciator is operating the controls.



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Forward Instrument Panel



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Mach/Airspeed Indicator (as installed)

• Airspeed information is provided by the respective air data computer tothe associated electric mach/airspeed indicator. The airspeed indicatordisplays indicated airspeed and Mach/Vmo information. The airspeedcursor control knob is located in the lower left corner of the mach/airspeed indicator. Pushing IN on the control knob, engages the automode. While in the auto mode, the airspeed cursor will be automaticallypositioned by AFDS/FCC commands. Pulling OUT on the control knob,the manual mode is selected. The airspeed cursor can be positioned bymanually rotating the control knob.

• Either one of the two airspeed indicators may be inoperative for dis-patch provided the remaining indicator operates normally. Refer to yourMEL.

Marker Beacons

• The marker beacon indicator lights are located on both the Captain'sand First Officer's forward flight instrument panel. These lights indicatebeacon passage for airways, outer, and middle approach markers. Themarker beacon receivers are designed to receive modulation frequen-cies from various electronic navigation facilities, when transmitting a 75Mhz vertical fan (boneshape) pattern. Located next to the display lights,is the HIGH/LOW switch. This switch is used to adjust the light sensitiv-ity of the receiver unit as displayed by the marker lights.



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• The marker beacon receivers may be inoperative for dispatch duringday VFR operations. For IFR operations, the ADF-LF navigation re-ceiver must be operative and weather conditions at the destination mustbe above approach minimums. The compass locator may be used as asubstitute without affecting the approach minimums. Refer to yourMEL.

Instrument Comparator System

• The instrument comparator system has been provided to give cockpitwarning to the flight crew of any significant deviations between theCaptain's and First Officer's compass headings, pitch/roll indicators,localizer, and glide slope deviations output from the No.1 and/or theNo.2 VHF navigation units. Comparator warnings are also provided fordeviations of the radio altimeter output signals. The instrument com-parator system may be inoperative for dispatch provided approach mini-mums do not require its use. Refer to your MEL.

Instrument Comparator Test Switch

• The instrument comparator test switch is located to the right side of theinstrument comparator lights. Selecting the test switch to position oneor two, illuminates all instrument comparator lights except the MONPWR light. This is a very common oral subject asked by examiners.



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Instrument Comparator Lights

• The illumination of a amber instrument comparator light indicates therespective instrument has exceeded the established tolerances. Referto diagram 1126 (page 175) for tolerances of the respective instru-ments.

Attitude Director Indicator

• (Oral Topic) Many examiners review the warning flags that may appearwithin the ADI. Those flags include the attitude warning flag, glide slopewarning flag, computer warning flag, runway flag, and the speed flag.The following discussion reviews each of those warning flag displays.

• The display of the red attitude warning flag indicates the display indica-tor may be unreliable for due to instrument power failure and/or thetesting of the ADI test switch. Depending on the type of failure, thedisplay may also indicate a 900 left bank. Selection of the vertical gyrotransfer switch may be used to provide reliable information from theoperable system. Depending on equipment installed, the transfer mayreceive information from the alternate vertical gyro. Refer to AFM forequipment installed in your aircraft.



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• The display of the red glide slope warning flag indicates the glide slopeinformation is unreliable with the ILS frequency tuned. This warning flagprovides parallel indications with the glide slope warning flag of the HSI.

• The display of the red runway flag indicates the localizer frequency istuned and the localizer signal is not valid. The loss of the radio altim-eter, may also cause the runway symbol to be displayed. In this case,the localizer function will not be impaired. The runway symbol willremain out of view with VOR frequencies selected.

• The display of the red computer warning flag indicates the flight directoris inoperative due to electrical power loss. This power loss will causethe flight director command bars to retract.

• The display of the red speed flag indicates the autothrottle system isinoperative.

Altimeter (as installed)

• Information presented on the associated electric altimeter is derivedfrom the respective air data computer. The apperance of the red warn-ing OFF flag over the digital counter box, indicates the ADC signal hasbeen lost. Internal system malfunctions can also cause the OFF flag tobe displayed. The apperance of the NEG flag in the two left-handwindows indicates the altitude is below zero feet.



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• On early aircraft, the Captain’s altimeter was electrically operated. If afailure occurred within the system, automatic switching would occurrfrom electric to barometric control. The “new” barometric data is sup-plied to the Captain’s atimeter from the First Officer’s system.

• The various altimeters are equipped with instrument vibrators. Thesevibrators can be either servo pneumatic or pneumatic. A combination ofboth types can also be installed. With the servo pneumatic version, onemay be inoperative for dispatch provided the associated air data com-puter operates normally. For the pneumatic type, one may be inopera-tive for dispatch provided VMC conditions exist for departure and arrival.Refer to your MEL for further information.

Standby Altimeter/Airspeed Indicator (as installed)

• The standby altimeter and airspeed indicator is a two-function displayindicator. The altimeter function provides standby reference informationobtained from the Captain’s static system. A green flag appears in theleft window when the altitude is below 10,000’. A striped flag appears inthe left window when the altitude is below zero feet. The standardbarometric correction display is set by the barometric setting controller.The standby altimeter has a range of -1000’ to 50,000’. The secondfunction of this indicator is standby airspeed drum. The standby air-speed indicator displays airspeed in knots. The Captain’s pitot-staticsystem is the source of pneumatic pressure used for the operation ofthe standby airspeed indicator.



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Radio Altimeter (as installed)

• Two low-range radio altimeters are provided to give reference of aircraftheight above the ground. The indicator provides visual reference ofaltitude up to 2500’ above the ground. When the Captain's radio altim-eter is inoperative, all modes of the GPWS are inoperative. A redwarning flag has be provided to warn the flight crew that a possiblemalfunction within the system has occurred. The following failures maybe at fault: instrument/system power failure, loss of return signal below2500’, incorrect altitude tracking, and/or radio altimeter testing has oc-curred.

• The radio altimeter test switch has been provided that performs a sys-tem check. The following items occur when performing the test: thealtitude pointer drives to 100’, the warning flag appears, and the DHlights illuminate at or below the altitude indicated by the DH cursor. Thedecision height light is located at the upper left-hand corner of the radioaltimeter indicator. The illumination of this amber light indicates thealtitude pointer is below the DH cursor.

• Both radio altimeters may be inoperative for dispatch provided approachminimums do not require its use. Refer to your MEL.



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Clock (as installed)

• Time is normally presented as either Greenwich Mean Time (GMT), in theupper digital display, or as chronograph (elapsed time), in the lower digitaldisplay.

• GMT time is displayed in the classic 24-hour format. The controls for GMTtime are located at the lower right corner of the indicator. Selecting thecontrol bar to the RUN position, starts the time display. The HLD mode,stops the time display and sets the seconds display to zero. The SS (slowslew) advances the time display as minutes movements only. The FS (fastslew) advances the time display as hours only. Local time can also beinserted, if desired. But, this technique of time display, is not normally usedby most flight crews.

• Chronograph and elapsed time controls are located at the upper left andlower left corners of the clock indicator. Pressing the chronograph controlknob (top left corner), controls the start, stop, and reset functions. Whenselecting the chronograph control knob, any existing elapsed time dis-played will be overridden. Elapsed time controls (lower left corner) hasthree functions. The control lever is spring loaded to the HLD position.Selecting RESET, returns the elapsed time digital display to zero time.Moving the lever to the HLD position, stops the elapsed time display at thecurrent indicated time. Selecting the RUN position, starts the elapsedtime function of the clock. The chronograph display reflects elapsed timerange as zero to 99 hrs 59 min and chronograph time range as zero to 99min.

• Either clock can be inoperative for dispatch, provided the other clock isoperating normally. Refer to your MEL.



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Radio Magnetic Indicator Compass - RMI

• The radio magnetic indicator compass has been provided to displayADF and VOR bearing information. The Captain’s RMI receives com-pass inputs from the #2 compass system’s directional gyros. The FirstOfficer’s RMI receives inputs from the #1 compass system’s directionalgyros.

• Warning flags have been provided to advise the flight crew of possiblepower failures and malfunctions within the system. The display of theVOR/ADF No.1 or No.2 warning flag, indicates either a power failureor an unreliable VHF NAV signal. The display of the heading warningflag, indicates the selected compass signal is invalid.

• The DME indicator located above the RMI indicator, has a 300 nauticalmiles maximum search for all DME stations. With the DME warning flagin view, indicates electrical power has been lost and/or an invalid DMEreceiver.



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Horizontal Situation Indicator (as installed)

• The Captain’s and First Officer’s HSI provides compass and VHF navi-gation information. Compass information is sent to the respective indi-cator directly from the respective compass controller. The associatedcompass controller computes information provided by the flux valvesand directional gyros.

• Power to the Captain’s HSI major components is provided by the 115VAC standby bus. This is the power source for the No.1 vertical gyro,No.1 instrument transformer, and the No.1 compass system. The DCstandby bus provides power for the No.1 VHF NAV and the No.1 glideslope. Power to the First Officer’s HSI major components is providedby the No.2 DC bus. This power supplies the 115V AC (unswitched)No.2 radio bus. Items on the unswitched radio bus are the No.2 com-pass system, No.2 instrument transformer, and the No.2 vertical gyro.The 28V AC (switched) radio bus powers the No.2 VHF NAV and theNo.2 glide slope.

• (Oral Topic) Several warning flags have been provided to warn thepilots of system component failures. These warning flags are commonsubjects for oral examinations. The display of the HDG warning flagindicates the selected compass is invalid and/or electrical power failure.The display of the VOR LOC warning flag indicates the navigation signal



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is below the acceptable level. Possible areas of malfunctions mayinclude NAV receiver failure and/or electrical power failure. The displayof the GS warning flag indicates the glide slope signal is below accept-able levels. Possible areas of malfunctions may include glide slopereceiver failure and/or electrical power failure.

• (Oral Topic) Occasionally during oral examinations, questions concern-ing the course deviation bar are asked. VOR deviation of 1 dot equals50. LOC deviation of 1 dot equals 10. NAV deviation of 1 dot equals 2nautical miles cross track deviation.

Vertical Speed Indicator (as installed)

• The Captain’s and First Officer’s electric vertical speed indicators dis-plays instantaneous VSI flight information, derived from the respectiveair data computer. Located on the electric vertical speed indicator is theOFF FLAG. This flag, when in view, indicates the air data computeraltitude rate signal has been lost, and/or a electrical malfunction hasoccurred.

Vertical Speed Indicator (100/200 - as installed)

• The Captain’s and First Officer’s pneumatic vertical speed indicatorsdisplays VSI flight information, derived from the respective pitot staticsystem or as supplied from the alternate system (when selected).



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Speed Brake Do Not Arm Light

• The amber speed brake do not arm light is located on the right side ofthe forward instrument panel. The illumination of this warning lightindicates abnormal conditions and/or internal system test inputs to theautomatic speedbrake system. An electrical fault within the system mayalso illuminate this amber warning light when conditions warrants. Thislight is deactivated when the speed brake lever is in the DOWN position.

Speed Brake Armed Light

• The green SPEED BRAKE ARMED light is located directly below theamber speed brake do not armed light. The illumination of this lightindicates valid automatic speedbrake system inputs. This light is deacti-vated when the speed lever is in the DOWN position. The valid inputindicates no faults are detected within the ground speed brake electricalsystem.



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Speed Brake Test Switches

• The speed brake test switches test the fault detection circuits of theautomatic speedbrake system. This test is a maintenance function only.

Speedbrake and Spolier System

• The primary purpose of the speedbrake and spoiler system is to supple-ment the ailerons for lateral control and to provide increased drag andlift reduction when the spoilers are used as speedbrakes. Speedbrakesconsists of 4 flight spoilers and 4 ground spoilers. Speedbrake opera-tion is controlled manually by the speedbrake lever or automatically byan electric actuator.

• The flight spoilers supplement the ailerons for lateral control. At 90 ofcontrol wheel movement, the spoiler system activates the flight spoilerup movement on the up aileron wing.

• Ground spoilers operate only on the ground to increased drag and toreduce lift. The automatic extension of all flight and ground spoilersoccurs if the speed brake lever is in the ARMED position and both thrustlevers are in idle. 60 kts of wheel spin-up on any two main wheels,causes the speed brake lever to automatically move to the UP position.Should the system not receive the wheel spin-up signal, the panels will



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still extend after the air-ground safety sensor changes to ground mode.This action occurs when the right main strut compresses. After touch-down, the panels will retract automatically if either thrust lever is ad-vanced. On a rejected takeoff, the panels will extend after wheel spin-up and after thrust reverser application.

• System A hydraulics provide actuator power for the ground spoilers andflight spoiler No.3 and No.6. System B hydraulics provide actuatorpower for flight spoilers No.2 and No.7.

• The auto spoiler system may be inoperative provided the system isdeactivated and that all operations are conducted in accordance withthe AFM. Pilots must verify manual spoiler operation and advise dis-patch of increases in landing field length. Consider landing technique ofmanual deployment of the spoilers prior to application of reverse thrust.



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Fuel Quantity Indication System

• The fuel quantity indication system consists of two methods of determin-ing fuel quantity. The primary method of determining fuel quantity is viathe hollow coaxial cylindrical capacitors connected in parallel with com-pensator units. This electrical indicating method uses the capacitorswith fuel acting as a dielectric. Changes in fuel levels alters the currentthrough the capacitors, thus causing a change in the fuel indicator read-ings in the cockpit and at the fueling station. The compensators areused to for density changes. Each tank has 12 capacitors and 1 com-pensator.

• (Oral Topic) The secondary method of determining fuel quantity con-sists of five manual measuring sticks for each main tank. This mechani-cal method uses dripsticks that sense fuel height and requires MELconversion charts to indicate the actual fuel quantities. This method ofdetermining fuel quantity is used only for tanks No.1 and No.2. Thereare no dripsticks for the center tank. The procedure for dripstick mea-suring commences with the unlocking of the stick head and the loweringof the fiberglass tube. When the fuel enters into the hollow tube andstarts flowing out of the bottom drip hole, the lowering of the stick isterminated. A marked reading on the outside of the calibrated tube istaken and compared to fuel charts located in the MEL book. The dripstick may be graduated in inches, gallons, or kilograms. Besure to note



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which markings are installed and that the dripstick reading point is takenon the inside wing plane. Use caution to ensure the red arrow on thebottom of the dripstick head is pointing away from all ground personnel.Beware of fueling with the No.2 dripstick extended to the desire fuellevel mark. Fuel may begin flowing from the bottom drip hole prior tothe fuel reaching the desired fuel tank level. For correct fuel readings,allow five minutes for fuel leveling before extending the dripsticks formeasurements.

• (Oral Topic) The fuel quantity indicator indicates usable fuel in therespective tank. Accuracy of the fuel quantity indicator is within +/- 3%of full scale indication. The source of power for this indicator comesfrom the standby AC bus.

• One main tank fuel indicator may be inoperative for dispatch providedthe respective tank is emptied and refilled with a known quantity of fuel.If this procedure can not be followed, then the dripstick method may beused following each refueling. Also, all boost pumps in the respectivetank must be operational, the center tank indicator must operate nor-mally, and the flight crew must periodically compute the fuel remainingvia a precomputed flight plan or chart.

• The center tank fuel indicator may be inoperative for dispatch providedthe center tank boost pumps operate normally.

Fuel Quantity Test Switch (Analog Indicators - as installed)

• The fuel quantity test switch provides a system test of the fuel indicatorsand quantity indication system. Selecting the PRESS switch, drives theindicators downwards towards the zero mark. The total fuel indicationwill also change during this test procedure. Located on the fuelingcontrol panel is a second test switch. When testing this switch, the fuelindicator pointers will move upscale. As a general pilot technique, donot perform the system test while the aircraft is being fueled. This willprevent a premature termination of the automatic fueling process.



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Fuel Quantity Test Switch (Digital Indicators - as installed)

• The fuel quantity test switch provides a system test of the fuel indicatorsand quantity indication system. Selecting the PRESS switch, illumi-nates all numbers and arcs, followed by the illumination of the maximumquantity for each tank. The system returns to normal fuel quantityindications. Whenever an error code of 2, 3, or 4 is displayed in thelower right corner of each gauge during the system testing. Should anyother code be displayed, indicates a malfunction. Contact your mainte-nance control.

• Digital Error Codes:Error 0 Incapacitating error has occurred and the indicator can not

compute fuel weight.Error 1 A compensator line has been shorted to a ground.Error 2 Plates of the compensator has been shorted or an exces-

sive leak has developed at the compensator.Error 3 Compensator leakage.Error 4 Shorted line or the line has completely opened.Error 5 Plates on the tank has shorted or excessive leak has

occured.Error 6 Tank leakage. Conductive contamination around tank

probes.Error 7 DCTU data out of limits.Error 8 DCTU mechanical error.Error 9 Indicator failure.Error 10 Signal line failure.



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Engine Pressure Ratio (EPR)

• The engine pressure ratio (EPR) indicator displays the ratio of turbinedischarge pressure (Pt7) to compressor inlet pressure for any thrustlever setting, except for engine idle. The EPR indicator is used as theprimary thrust setting reference. The engine exhaust pressure (Pt7) issensed by six probes projecting into the exhaust system that providesan average exhaust pressure. The engine inlet pressure (Pt2) is sensedby a pitot probe mounted through the center of the nose dome. Theprobe is heated for anti-icing by engine bleed air that is present in thenose dome. System maintenance test connections are located on thebottom section of the forward engine area. The power source for theEPR indicators is from the respective 115V AC transfer bus.

• (Oral Topic) Three different “warning flags” have been provided to alertthe crew of various system failures. The display of the warning flagover the digital readout indicates the loss of electrical power and/orinstrument failure. The display of the warning flag over the EPR windowindicates the failure of the transmitter. And, the display of the warningflag covering the lower digital window indicates failure of the PDC.



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• EPR reference selector (bottom right corner of the indicator), com-mands automatic or manual selection of the EPR reference cursor. Thereference cursor displays input signals from the PDC or manual inputsfrom the crew. The IN position of the selector commands automaticdisplay from the PDC and the OUT position displays desired EPR val-ues for reference only. The display of the letter "M" indicates manualoperations.

• Both EPR indicators are required for dispatch, except one EPR refer-ence selector cursor may be inoperative. Both digital counters may beinoperative for dispatch. Refer to your MEL.

N1 RPM Indicator

• The N1 RPM display indicates low pressure compressor speed in per-cent of RPM for monitoring engine performance. This instrument is selfpowered, only the integral lighting of the tachometer requires systempower. Some indicators require AC power from the standby bus foroperations, refer to your operations manual for type installed. Theactual location of the low pressure tachometer (N1) is on the frontaccessory drive pad behind the nose dome.

• The indicator dial has graduated display readings between zero and110% RPM, with small readings graduated in 10 units for each 10percent of change in speed indications.

• Only one N1 indicator may be inoperative for dispatch provided therespective engine’s N2 and fuel flow indicators are operating normally.The digital display portion of the indicator may be inoperative for dis-patch. Refer to your MEL.



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Exhaust Gas Temperature

• The exhaust gas temperature indicator has been provided for crewmonitoring of the engine’s mechanical integrity of the engine’s turbines.The EGT indicator displays turbine gas temperature in degrees 0C, assensed by eight thermocouples. The indicator requires 115V AC power,as provided by the standby bus. The eight thermocouples are arrangedin a circular pattern in the engine exhaust. This provides an averageexhaust gas temperature for cockpit display.

• The indicator dial has graduated display readings between zero and8500C, with expanded readings between 5000 and 7000 for more accu-rate display indications.

• Both EGT indicators are required for dispatch. Except for EIS equippedaircraft, only the digital portion of the indicator may be inoperative fordispatch. Refer to your MEL.



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EGT Limitations

Condition JT8D-9 JT8D-9A JT8D-15

Takeoff (5min) 580C 590C 620C

Max. Cont 540C 545C 580C

Gnd Start (+15C) 420C 420C 550C

Gnd Start (-15C) 350C 350C n/a

Gnd Start n/a n/a 550C

Flt Start n/a n/a 620C

N2 RPM Indicator

• The N2 display indicates high pressure compressor speed in percent ofRPM. This instrument is self powered, only the integral lighting of thetachometer requires system power. The actual location of the N2tachometer is on the right side of the accessory drive case.

• The indicator dial has graduated display readings between zero and110% RPM, with small readings graduated in 10 units for each 10percent of change in speed indications.

• One N2 indicator may be inoperative for dispatch provided the respec-tive engine’s N1 and fuel flow indicators are operating normally. Analternate starting procedure must also be used for starting the engine.This alternate starting procedure begins by starting the engine with theoperative N2 first. Starting times and N1 values are noted and used torepresent the inoperative indicator’s N2 RPM values for the movementof the start lever to the idle position. In addition, N1 RPM is noted forstarter cutout purposes. Refer to your MEL for detailed proceduresconcerning starting and aborted start procedures with an inoperative N2indicator.



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Fuel Flow Indicator

• The fuel flow indicator displays fuel consumption rate in pounds perhour. The digital readout on the respective engine’s indicator reflectstotal fuel consumed for that engine. The electrical power for the indica-tor is provided by the 115V AC transfer bus.

• One fuel flow indicator may be inoperative for dispatch provided theassociated N1, N2, and fuel quantity indicator is operating normally.Refer to your MEL.

Oil Pressure Indicator

• The oil pressure Indicator displays engine oil pressure in PSI as mea-sured within the engine oil distribution system. The oil system ispressuirzed by the engine driven pump oil pumps located within theaccessory gear drive case. The oil pressure indicating system has beendesigned to sense oil pump output pressure on one side of an internaldiaphragm. The other side of the diaphragm has ambient pressures.

• Electric power for the oil pressure indicator is received from the 28V ACtransfer bus. Any power interruption will cause the indicator to displaythe last pressure sensed.



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Oil Temperature Indicator

• The oil temperature indicator displays oil temperature as monitoredwithin the oil distribution system. The temperature sensor is located atthe outlet from the oil cooler. The temperature sensor is of a resistancetype temperature bulb design that has direct contact with the oil. The oiltemperature indicator is of a resistance ratiometer design. Electricpower for the oil temperature indicator is received from the 28V ACtransfer bus.

Oil Quantity Indicator

• The oil quantity indicator provides a visual indication of usable oil con-tained in each oil tank. The indicator dial has graduated display read-ings that indicates amounts in gallons. When the indicator displays anamount indicating zero, approximately 1.3 gallons remain within the oillubrication system. Minimum oil quantity for dispatch is 2.5 gallons or50% (refer to your company operations manual). Oil quantity indicationmay be inaccurate if the engine has been shut down more than 30minutes.

• (Oral Topic) The No. 2 oil tank capacity is approximately .5 gallon lessthan the No. 1 oil tank. This is due to the dihedral of the wing andinterchangeable engines. The No. 1 oil filter port is higher than the No.2 oil filter port.



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• The oil quantity sensing unit is of a capacitance type design and islocated within the oil tank. Electric power for the oil pressure indicatorsis received from the respective 115V AC transfer bus.

Oil Test Switch

• The oil test switch has been provided to test the oil quantity indicatorpointer and indicator. Selecting the test switch, drives the oil quantitypointers towards the zero level marking. This action is accomplished byconnecting the area of the oil quantity sensing unit to a ground, thussimulating an empty tank. Releasing the switch, the indicator pointerreturns to the previous display.

Reverser Unlock Light

• The illumination of the amber REVERSER UNLOCK light indicates thethrust reverser door is not stowed and/or in the locked position. Referto thrust reverser section of this guide for further information.

• One reverser unlock light may be inoperative for dispatch provided thethrust reverser has been “visually verified” to be in the closed position.“Visually verified” refers to the inspecting of the overcenter links andguide carriage for actual position. Refer to your MEL .

Start Valve Open Light

• The illumination of the amber START VALVE OPEN light indicates thestarter valve is open and air is being supplied to the starter motor. If theengine starter does not cutout by 40% N2, or if the START VALVEOPEN light illuminates ground operations, crew action is to place theengine start switch to the OFF position. If the light remains illuminated,then the start lever should be selected to CUTOFF, the isolation valveshould be positioned to CLOSE, and the respective bleed switch shouldbe placed to the OFF position. For inflight operations, the illumination ofthe start valve open light may require the engine to be secured. Theseactions will isolate bleed pressure from the engine start valve and pre-vent possible damage to the starter motor. Refer to your airline com-pany procedures for details.



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• The START VALVE OPEN light may be inoperative for dispatch pro-vided the start valve arming system operates normally. Normal startingprocedures should be followed. Verification of the opening and closingof the start valve is observed using the duct pressure indicator. Refer toyour MEL for procedures.

Low Oil Pressure Light

• The illumination of the amber LOW OIL PRESSURE light indicatesengine oil pressure is below 35 psi. It is essential that a “cross-check”of other engine instruments be accomplished prior to crew action. Withengine oil pressure in the yellow band, the thrust lever should be re-tarded. The engine can be operated at a reduced power setting. Withengine oil pressure at or below the red radial markings, securing theengine should be considered. Refer to your operations manual fordetails.

• The sensor for the LOW OIL PRESSURE light is down stream of the oilpressure transmitter on the left side of the accessory drive case. Thelow oil pressure sensor senses oil supply pressure. The electricalpower source for the oil pressure indicating circuit is supplied from therespective 28V DC transfer bus.

• One LOW OIL PRESSURE light may be inoperative for dispatch pro-vided the respective engine’s oil quantity, oil temperature, and oil pres-sure indicators operate normally. Refer to your MEL with reference tothrust reverser operating notes.

Oil Filter Bypass Light

• The illumination of the amber OIL FILTER BYPASS light indicates an“impending bypass” of the main oil filter. The “impending bypass” refersto the design of the oil filter differential pressure switch that senses filterinlet and outlet pressures. When the filter begins to block, the differen-tial pressures increases, the switch produces a ground and the lightilluminates.



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• During cold weather operations, the OIL FILTER BYPASS light mayremain illuminated after starting the engines. As the oil begins to warm,the bypass light extinguishes itself. Normal oil warming occurs withinfive minutes of engine starting. If the light remains illuminated after thistime, consider securing the engine.

• The illumination of the OIL FILTER BYPASS light during noncriticalflight operations may require the engine to be operated at a reducedthrust setting. Just enough thrust should be used to keep the lightextinguished. If the light remains illuminated, the engine may require tobe secured.

Yaw Damper Indicator

• The yaw damper Indicator is located on the center instrument panel, justabove the engine instruments. The indicator displays yaw dampermovement of the rudder. Pilot rudder pedal movements are not dis-played.

• The yaw damper system has been designed to prevent unwanted dutchroll. The yaw damper system receives yaw signals from the rate gyrosand sends those signals to the yaw damper coupler. All of this informa-tion is then sent to the rudder power control unit that moves the rudderfor yaw control. No cockpit rudder pedal movement can be felt due toyaw damper motion.



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• (Oral Topic) During yaw damper operations, airspeed signals from theair data computer will decrease the amount of yaw damper movement(deflection) as the aircraft airspeed increases.

• The yaw damper may be inoperative for dispatch provided the yawdamper switch remains OFF. Aircraft with the SP-77 autopilot, the flightmanual limitation that restricts the use of the autopilot (except for pitchmode) to 30,000 feet or below applies. Refer to your MEL.

Flap Position Indicator

• The flap position indicator is located on the forward center instrumentpanel. The flap indicator displays the angular position of the left andright trailing edge flaps. The flap indicator receives position informationfrom two transmitters mounted on the outboard flap torque tube in eachwing. Electrical power source for the indicating system is supplied fromthe No.2 28V AC transfer bus.

• (Oral Topic) The secondary function of the flap position indicationsystem is to provide trailing edge flaps asymmetry protection. Thebasic concept of asymmetric protection is to stop hydraulic operation ofthe trailing edge flaps when a significant difference exists between theposition of the left and right trailing edge flaps. Refer to asymmetric flapprotection section for additional information.



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Anti-skid System

• The anti-skid system is designed to provide maximum and effectivebraking for any runway condition without skidding. Each main wheelhas been provided with individual transducers which signal wheel speedinformation to the anti-skid control unit. The anti-skid control Unit thenelectronically regulates the anti-skid valves to control braking pressurewith regards to wheel deceleration speed.

• (Oral Topic) The anti-skid system provides anti-skid protection, lockedwheel protection, touchdown protection, and hydroplane protection. Theanti-skid system controls the amount of hydraulic pressure that is ap-plied for manual braking and/or autobraking.

• Power source for the outboard anti-skid system is supplied from theNo.1 transfer bus and the power source for the inboard anti-skid systemis supplied from the No.1 transfer bus. Thence, when operating onstandby power only, the anti-skid systems will be inoperative.

• The air-ground sensors (located on the right main landing gear), sup-plies control logic for the anti-skid system. This control logic providesbrake release for touchdown protection and allows normal anti-skidbrake pressure after wheel spinup.



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• The anti-skid system continuously monitors itself and checks for:1. Transducer: open or short.2. Normal valve: open circuit.3. Power loss or control switch position.4. Failure in the control unit (normal system).5. Differences between parking brake valve and/or switch inequity.6. Park brake set in air.

Anti-skid Inop Lights

• The illumination of the respective amber anti-skid inop light indicates asystem fault has been detected within the associated anti-skid monitor-ing system. The illumination of both anti-skid lights at the same timeindicates a disagreement between the parking brake lever position andthe parking brake shutoff valve position.

Anti-skid Control Switch

• The anti-skid control panel has incorporated two anti-skid controlswitches. The left switch controls the inboard anti-skid and the rightswitch controls the outboard anti-skid. The two position switch, ON/OFF, controls the electrical power supplied to the respective anti-skidcontrol unit. Selection of the OFF position, illuminates the associatedanti-skid inop light.

• The anti-skid system(s) may be inoperative for dispatch provided alloperations are conducted in compliance with the AFM. If only onesystem is inoperative, the operative system should be selected ON toprovide antiskid protection. The anti-skid control switch(es) must beselected OFF to ensure full manual braking capability on the inoperativesystem. With anti-skid inoperative, payload considerations should bereviewed, since takeoff & landing runway length limitations may be afactor. In addition, speedbrakes must be manually extended sinceautomatic extension of the speedbrakes may not occur. Refer to yourMEL for details



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Autobrake System

• The autobrake system is designed to provide smooth, consistent, andimmediate brake application on touchdown. This automatic brake actionrequires both antiskid systems to be ON and operational. Theautobrake system applies hydraulic pressure to all brakes to slow theairplane at the rate selected. The antiskid system maintains priorityover the autobrake at all times to protect against skid or locked wheels.

• Autobraking is initiated when both thrust levers are retarded; at leastone wheel speed on each side of the airplane is greater than 60 kts, andthe average wheel speed is greater than 70 kts. The first stage of initialbrake pressure of 200 psi, is followed by a positive pressure rate of 100psi/second for 15 seconds. A second rate proportional to the decelera-tion selected, achieves the selected deceleration within 3 seconds.

• (Oral Topic) Arming of the autobrake system occurs when:1. Air-ground safety sensor is in the flight mode.2. Anti-skid switches are selected ON.3. Autobrake selector positioned to rate level.



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• (Oral Topic) Activation of autobrake pressure occurs when:1. Thrust levers are retarded to idle.2. Main wheel spin-up.

• (Oral Topic) Termination of autobraking occurs when:1. Movement of the spoiler handle to down detent.2. Application of both brake pedals.3. Selection of the autobrake select switch to OFF.

Autobrake Select Switch

• The autobrake select switch is used to select the level of desired brak-ing. The knob-switch must be pulled out to select the MAX position ofdeceleration.

• The autobrake system may be inoperative for dispatch provided thesystem is deactivated. Refer to your MEL.

Autobrake Inop Light

• The illumination of the amber autobrake INOP light indicates a malfunc-tion exists in the autobrake system.

Autobrake Selection Criteria

MIN Autobrake position MIN, provides a nominal deceleration rate thatprovides light manual braking at 80 kts.

MED Autobrake position MED, provides moderate deceleration rates.This is normally used during wet/icy runways, slippery runways,and/or limited landing runways distances.

MAX Autobrake position MAX, provides the maximum deceleration ratefor minimum stop distance.



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Braking Action Reports

• Braking action reports may be given in two formats as listed below. TheMU report equals the standard report as shown.

BRAKING ACTIONS MU REPORTSGood .35 or greaterFair .26 to .34Poor .18 to .25NIL .15 or less

Landing Gear Indicator Light

• (Oral Topic) The landing gear indication system has been designedwith three red indicator lights, three green indicator lights, and an auralwarning horn. The system provides the flight crew with visual and auralwarnings of various landing gear conditions. The illumination of any redindicator light indicates the landing gear is in transit and/or the landinggear lever and the landing gear do not agree. The red lights alsoprovide visual warnings when the aircraft is in a possible landing con-figuration and the landing gear is not extended and locked. In conjunc-tion with the red visual warning, an aural warning can be heard when theaircraft is in the landing condition and any gear is not extended andlocked.



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• The illumination of any green indicator light indicates the respective gearis extended and locked. The landing gear warning horn is deactivatedwhen all landing gears are extended and locked. Electrical power issupplied from the 28V DC battery bus.

• The landing gear indication system has been designed with proximityswitch-type sensors. There are two downlock sensors for each maingear, that provides gear downlock indications. The sensors are knownas the primary and secondary. They are located on the outboard side ofeach main gear strut. The primary sensor provides signals for the gearindicating light system and the secondary provides signals for the gearaural warning system. The uplock sensor signals are supplied by oneproximity switch-type sensor. The uplock sensor is located on the sideof the uplock hook of each main gear. The nose gear uses separatedown and lock sensors. The two nose gear sensors are located on thelock brace unit.

• Either of the two systems (visual or aural), may be inoperative fordispatch, provided the center panel visual indications operate normally.Refer to your MEL.

Landing Gear Warning Horn

• (Oral Topic) The landing gear warning horn has been provided to giveaural warnings when the aircraft is in the landing configuration and anygear is not extended and locked. The horn is activated by thrust leverand flap positions. The landing gear warning horn will sound steadywhenever the following conditions exists.

1. Flaps are located at positions 1 through 10, the horn will soundwhen either or both thrust levers are retarded to the idle position.The horn can be silenced with the landing gear horn reset switch.

2. Flaps are located at positions 15 or 25, the horn will sound wheneither, but NOT both thrust levers are retarded to the idle position.The horn can be silenced with the landing gear horn reset switch.The horn CANNOT be silenced when BOTH engines operating lessthan 1.6 EPR.



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3. Flaps are located at positions 30 or 40, the horn will sound regard-less of thrust lever position or engine EPR settings. The horncannot be silenced by any means.

• The aural gear warning horn is located forward of the control stand,below the first’s officers instrument panel. The horn will not sound whenthe trailing edge flaps are in the up retraced position (up position.)Electrical power is provided by the 28VDC battery bus system.

Takeoff Configuration Warning Horn

• (Oral Topic) The takeoff warning horn has been provided to give theflight crew an aural warning of possible unsafe configurations that mayexist. The system is armed when the aircraft is on the ground andeither engine has accelerated towards the takeoff power levels. Thewarning horn provides an intermittent warning and can only be cancelledwhen the unsafe configuration has been rectified. The warning horn islocated forward of the control stand, below the first’s officers instrumentpanel. The following items will sound the intermittent takeoff warninghorn.

• Speedbrake is not in the down position.• Trailing edge flaps are not in positions 1 through 25.• Stabilizer trim is not in the green band area.• Leading edge devices are not in the proper position for takeoff.

Note: A simple way to remember these items, is the use of a mental recallchecklist of “spoilers, flaps, and trim”. These are the configurationitems that can ruin your day if they are not properly set.

Landing Gear Lever

• Landing gear operation is accomplished by using hydraulic pressureprovided by System A. The landing gear lever controls the a selectorvalve, that allows System A pressure to the landing gear actuators forextension and/or retraction. The selection of the gear handle to the UPposition, mechanically operates the selector valve (via a cable system),allowing hydraulic pressure to enter the “up lines” for retraction. Theselection of the gear handle to the DOWN position, mechanically oper-ates the selector valve (via a cable system), allowing hydraulic pressureto enter the “down lines” for extension. The OFF position, blocks hy



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draulic pressure at the selector valve. The OFF position is the normalcruise position with the gear retracted.

• (Oral/Simulator Topic) To prevent possible retraction of the landinggear while on the ground, a solenoid operated latch has been provided.The solenoid latch operates with the air-sensing proximity sensor.When the aircraft is on the ground, the air sensor de-energizes thesolenoid and moves a latch into a position that prevents the physicalmovement of the landing gear handle to the UP position. The failure ofthe air/ground sensor, is a common simulator fault that may occur dur-ing simulator training. This problem may occur shortly after departureand is indicated by the failure of the landing gear handle movement tothe UP position. The checklist refers to two situations for gear handlemovement failure.

Situation No.1: Landing Gear Solenoid FailureWith the landing gear down, flaps retracted, and the takeoff warninghorn is silent: condition indicates the landing gear solenoid had failed.Procedures call for the selection of the landing gear override trigger andthe movement of the gear handle to the UP and OFF positions. Refer-ence the landing gear override trigger description section. Refer to yourAFM for details.

Situation No.2: Air/ground Sensor FailureWith the landing gear down, flaps retracted, and the takeoff warninghorn is sounding: condition indicates the air/ground sensor has failed.Procedures call for the pulling of the fanding gear AIR-GRD relay circuitbreaker and landing at the nearest suitable airport. The reason for anearly landing, is that other aircraft systems may also be affected by thisfailure. Those systems include pressurization, electrical standby powertransfer, and standby hydraulic power activation to name a few. Referto your AFM for details.



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Landing Gear Override Trigger

• The purpose of the override trigger is to allow the landing gear to beraised, bypassing the solenoid lock. The solenoid lock has been pro-vided prevent possible retraction of the landing gear while on theground. The solenoid latch operates with the air-sensing proximity sen-sor. When the aircraft is on the ground, the air sensor de-energizes thesolenoid and moves a latch into a position that prevents the physicalmovement of the landing gear handle to the UP position.

Manual Gear Extension Handles

• The manual gear extension system has been designed to provide thecapability to lower the landing gear when hydraulic system pressure isnot available. These manual procedures should also be used wheneverthe landing gear lever is placed to the DOWN position and the greengear light(s) do not illuminate. The system comprises of three manualcontrol handles (one for each landing gear), that operates a series ofcables and drums that will release the respective gear from the up andlocked position when pulled. The handles are located on the centerflight deck floor, under a small access door.



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• Prior to operating the manual extension system, the landing gear controllever should be selected to the OFF position to prevent possible hydrau-lic locks during the manual extension process. The OFF position, alsoremoves hydraulic pressure from the landing gear actuators. A singlepull of the control handle, approximately 18 inches in length (45 lbs ofhand pressure) for the main gear and 8 inches in length (25 lbs of handpressure) for the nose gear, will release the associated landing gear.The gear will free-fall into the locked position, and thus illuminating therespective green landing gear indicator light. After the illumination of allindicator lights, the landing gear lever should then be selected to theDOWN position. Should any green indicator light fail to illuminate, theuse of the gear down lock viewers will be required to verify the properalignment of the mechanical downlock indicator markings. Warning, donot hold the manual extension handles during normal hydraulic opera-tion of the landing gear.

Main Gear Downlock Viewer

• The main gear downlock viewer provides the means for inflight visualinspection of the main gear downlock indicators when the normal lightindicating system is inoperative. The downlock viewer is located in thefloor near aisleway of the main cabin, at approximately the 3rd passen-ger window aft of the overwing emergency exit door. Mirrors are



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aligned with cutouts in the viewer tube, and are arranged so that themain landing gear downlock indicators are centered in the field of visionof each mirror. When a main gear is down and locked, the red paintstripes on the lower side strut will align with the red paint stripe on thelower downlock link.

• The wheel well light switch (inspection flood lights - main gear), must beon to illuminate the area for inspection. The inspection flood lights maybe inoperative for dispatch during day operations only. For other typesof operations, the lights may be inoperative provided a landing gearindicating system other than the viewer system and independent of thecenter panel indicating system has been installed. Refer to your MEL.

Nose Gear Downlock Viewer

• The nose gear downlock viewer provides the means for inflight visualinspection of the nose gear downlock indicators when the normal lightindicating system is inoperative. The nose gear viewer window andcover are located in the flight deck floor. Two red arrows are painted onthe lock strut, one on the lock link and one on the lock brace. Down andlocked indications are shown when the red stripes are aligned. Thewheel well light must be on to illuminate the area for inspection.



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Total Air Temperature Indicator (as installed)

• The digital Total Air Temperature indicator displays temperature datareceived from the TAT computer via a single exterior TAT probe. TheTAT probe is located on the left forward fuselage. Temperature valuesfrom -60 C0 to +60 C0 can be displayed from the indicator. TAT tem-perature values are valid only during inflight operations.



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• (Oral Topic) The definition of Total Air Temperature (TAT), is outsideair temperature PLUS all the ram rise. During ground operations, TATvalues are approximately the outside air temperature when the pitotheat switch is selected OFF.

• The total air temperature indicator may be inoperative for dispatch pro-vided an “alternate means” of determining temperature is available.“Alternate means” of temperature may include temperature as displayedby the PDCS, SAT, or RAT. Refer to your MEL.

Total Air Temperature Indicator

• The total air temperature indicator displays temperature received di-rectly from the single exterior TAT probe. The TAT probe is located onthe left forward fuselage. Temperature values from -70 C0 to +50 C0

can be displayed from the indicator. TAT temperature values are validonly during inflight operations.

• The total air temperature indicator may be inoperative for dispatch pro-vided an “alternate means” of determining temperature is available.“Alternate means” of temperature may include temperature as displayedby the PDCS, SAT, or RAT. Refer to your MEL.



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Air Temperature/True Airspeed Indicator (as installed)

• The temperature and airspeed indicator provides selectable displays oftotal air temperature (TAT), static air temperature (SAT), and true air-speed (TAS). TAT temperature values are valid only during inflightoperations, however, outside air temperatures (OAT) can be indicatedduring ground operations with the pitot heat switches OFF. All tempera-ture and airspeed values are received from the No.1 ADC.

• A selector push button has been provided to select in sequence TAT,SAT, or TAS displays. As the value is presented in the center of theindicator, an annunciation of the TAT, SAT, or TAS is illuminated aboveit. The digital displays are presented in 0C for TAT and SAT. Trueairspeed values are presented in knots.

• The total air temperature system uses a single temperature probe,which is located on the left forward fuselage. The TAT probe has beendesigned with three internal sensing elements, providing temperaturedata to each ADC. TAT data from the No.1 ADC is provided to theFMC, both IRSs, FCC “A”, and the autothrottle system. TAT data fromthe No.2 ADC is provided to the FMC, both IRSs, FCC “B”, and theautothrottle system

• The total air temperature indicator may be inoperative for dispatch pro-vided an “alternate means” of determining temperature is available. Re-fer to your MEL.

TAT-MAX EPR Indicator (as installed)

• The true air temperature and maximum EPR indicator displays TATtemperatures, maximum EPR values for selected flight modes, and an-nunciates the flight mode selected. TAT temperatures displayed re-flects ambient air temperature that has been corrected for compressionheating (ram rise). TAT temperature values are valid only during inflightoperations. A warning flag has been provided to give visual warnings ofTAT signal failures.



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• The maximum EPR portion of the indicator provides a continuous dis-play of maximum allowable EPR for the flight mode selected. Data fromthe TAT probe and from the ADC are used to determined the correctEPR value displayed. The values are also automatically adjusted forthe use of engine bleed air. A warning flag has been provided to givevisual warnings of internal system and/or electronic failures.

• The EPR flight mode selector is located on the lower right corner of theindicator and provides the means to select the appropraite maximumEPR for the current flight conditions. Four modes of EPR display areavailable for selection. They consist of GA (maximum go-around),CONT (maximum continuous), CLIMB (maximum climb) and CRZ(maximum cruise). Pressing the selector knob, provides the means totest the system and drives the TAT and MAX EPR displays to a presetvalue.

Hydraulic Brake Pressure Indicator

• The hydraulic brake pressure indicator is located on the First Officer’sflight instrument panel. A secondary brake pressure indicator is alsolocated in the wheel well area. The indicator displays System A andSystem B brake pressures. Pressure indications are sensed from theprecharge side of the brake accumulator. Nominal pressures of 3000psi are normally displayed, maximum limits of 3500 psi are indicated bythe shaded band. Accumulator precharge pressures are shown as 1000psi.



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• (Oral Topic) Displays of zero pressure, indicates the precharge hasbeen lost. Loosing the nitrogen precharge will have no affect on normalbraking, as long as normal hydraulic pressures are maintained.

• The brake system receives hydraulic pressure from two independenthydraulic sources. System A pressure is provided to the inboard brakesand System B pressure to the outboard brakes. A brake accumulatorhas beem provided for each brake system. The brake accumulator hasseveral functions. The accumulator stores hydraulic pressure for brakesoperations, maintains instantaneous flow of fluid to the brakes, anddampens pressure fluctuations.

• (Oral Topic) With the lost of normal system pressures, accumulatorbraking provides approximately 5 to 6 applications of “emergency” brak-ing. The accumulator is precharged with nitrogen or compressed dry airto 1000 psi.



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• Brake pressure is controlled by the anti-skid system. Automatic brakingis provided during gear retraction to the main gear wheels. Nose wheelstopping (during gear retraction), is accomplished by brake snubberslocated in the top of the nose wheel area.

• Either of the two brake pressure sensing units (A or B), may be inopera-tive for dispatch provided that both wheel well brake pressure indicatorsoperate normally. These wheel well indicators must also be visuallychecked prior to each departure. Refer to MEL.

Hydraulic System Pressure Indicator

• The hydraulic system pressure indicator is located on the First Officer’sflight instrument panel, just below the brake pressure indicator. Thepressure indicator displays System A and System B hydraulic pres-sures. Each hydraulic system is incorporated with two separate sensingunits. This design provides backup pressure indications should eithersensor and/or indicator failure. The system pressure indicator sensoris located downstream of the pumps and check valves, prior to the usingunits. The low pressure caution light’s sensors are located in the pumpoutput lines. Nominal pressures of 3000 psi are normally displayed,maximum limits of 3500 psi are indicated by the shaded band. Selec-tion of the associated system pumps to the OFF positions, will causethe respective pointer to display zero pressure.



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• The hydraulic pressure indicator (A and B), may be inoperative fordispatch provided system pressure can be verified from the brake pres-sure indicator prior to each flight and all hydraulic low pressure lightsmust operate normally. Refer to your MEL.

Hydraulic System B Low Quantity Light

• (Oral Topic) The amber hydraulic system B LOW QUANTITY light islocated on the First Officer’s flight instrument panel. The illumination ofthe caution light indicates hydraulic fluid level is low.

• (Oral Topic) System A and System B reservoirs are located in the mainwheel well and are interconnected by fluid balance lines. To demon-strate the design concept of the hydraulic balance line, imagine theillumination of the B LOW QUANTITY light. This indicates System Bhydraulic fluid has been reduced to approximately .65 gallons. Now,observe System A quantity indicator. System A quantity level displaysapproximately 1.84 gallons, indicating the fluid level inside System Areservoir has also been reduced. This reduction of fluid terminates atthe top of the fluid balance line standpipe, thus preventing any furtherreduction in fluid from System A to System B reservoir.



B737-200

Hydraulic System A Quantity Indicator

• The System quantity indicator is located on the First Officer’s flightinstrument panel. A remote quantity indicator is located at the hydraulicreservoir within the wheel well. The indicator displays hydraulic quantitywithin the reservoir. With System A full, the indicator displays 3.5 USgallons. With the indicator pointer at the RF (refill) mark, indicates thereservoir contains 2.4 US gallons.

• System A quantity indicator may be inoperative for dispatch providedhydraulic quantities are checked prior to departure and System A pres-sure indicator operates normally. System B and the standby hydraulicsystem low quantity lights must also operate normally. Refer to yourMEL.

• (Oral Topic) There are no hydraulic quantity indicators for System B.System A and System B reservoirs are located in the main wheel welland are interconnected by fluid balance lines. By the design of thehydraulic system, System B reservoir should always be full. With adecrease of hydraulic fluid in System B, System A quantity indicatorshould show the decrease. Refer to the hydraulic system low quantitylights section for further information.



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Nose Wheel Steering

• Nose wheel steering is provided for aircraft directional control duringground operations. Nose wheel steering uses Hydraulic System A pres-sure through the landing gear down line. Normal steering is accom-plished by either the steering control wheel or the rudder pedals. Bothare mechanically connected to the nose gear steering valve by cables.The steering valves directs 3000 psi of hydraulic fluid to the nose gearsteering cylinders. This action turns the steerable portion of the landinggear. The steering control wheel (Captain’s side panel), can turn thenose wheel 780 from center and the rudder pedals at full deflection canturn the nose wheel 70 from center. The wheel turn of 950 providescontrol for the maximum steering angle of 780. Rudder pedal steeringbecomes inactive as the nose gear strut extends.



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GPWS Inoperative Light

• The illumination of the amber GPWS INOP light indicates the loss ofelectrical power and/or invalid inputs from the VHF NAV, ADC, or radioaltimeter are being received. In case of GPWS malfunctions, it isrecommended not to deactivate the GPWS by pulling the respectivecircuit breaker or the use of the flap/gear inhibit switch. Only use thosemethods of deactivation when approved procedures required it.

• The GPWS system may be inoperative for dispatch provided alternateprocedures are followed and repairs to the system are made within 3flight days. Pilots should maintain the MEA along all published airwaysand operate above the minimum IFR altitudes when off the airways.Refer to your MEL.

GPWS System Test Switch

• The GPWS system has been designed with an internal self-test featurethat performs the self-test when the SYS TEST switch-button is se-lected. Upon pressing the button, the GPWS, BELOW G/S, and theINOP lights will illuminate. The aural warnings of the GPWS system(“GLIDE SLOPE - WHOOP WHOOP PULL UP”) will also sound, indicat-ing proper operation. The GPWS test system has been designed so thetest feature is deactivated between 50 and 1000 feet radio altitude,anytime the landing flaps are selected, and when the aircraft is airborne.



B737-200

GPWS Flap/Gear Inhibit Switch

• The guarded flap/gear inhibit switch has two modes of operation avail-able, NORMAL and INHIBIT. During daily operations, the switch isplaced in the NORMAL guarded-position, thus providing flap and landinggear position logic for GPWS Modes 2, 3 and 4. The lifting of theprotective guard and the selection of the switch to the INHIBIT position,cancels warnings and alerts caused by the flaps not being in a landingposition. The INHIBIT feature also cancels warnings as they relate tothe position of the landing gear. This INHIBIT feature for the landinggear warnings is only used during partial and/or gear up landing proce-dures.

Ground Proximity Warning System (as installed)

• (Oral Topic) There are several different models of GPWS systemsavailable for the B737. The following discussion reviews the Mark IIGPWS system and has been provided as an introduction to the conceptof the GPWS system. Some examiners may ask general questions withregards to the basic modes of operation and the respective warningsprovided. Refer to your AFM for detailed information concerning themodel installed in your aircraft.

• The Mark II ground proximity warning system has been designed toprovide the flight deck crew visual and aural warnings of unsafe flightconditions and/or configurations. The GPWS system has five modes ofoperations available between radio altitudes of 50’ and 2450’. TheGPWS computer receives data inputs from the radio altimeter, baromet-ric altitude, No.1 ADC, Captain’s glide slope, and gear/flaps positions.The loss of any data input to the GPWS computer, will only affect themode of warning using that information.

• The five modes of operations of the GPWS system are:Mode 1 Excessive descent rate.Mode 2 Excessive terrain closure rate.Mode 3 Altitude loss after takeoff or go-around.Mode 4 Unsafe terrain clearance - not in landing configuration.Mode 5 Excessive deviation below glide slope.



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• Mode 1: Excessive Descent Rate: Provides aural and visual warn-ings for excessive descent rate to terrain (aircraft configuration not afactor). The two aural warnings are announced based on the ratio ofdescent rate and radio altitude. The initial warning of “SINK RATE” isgiven, followed by “WHOOP WHOOP PULL UP” if the condition wors-ens. Visual alert warnings are provided by the illumination of the redPULL-UP lights (located on the forward flight instrument panel).

• Mode 2: Excessive Terrain Closure Rate: Provides aural and visualwarnings for excessive terrain closure rate. The initial warning of “TER-RAIN - TERRAIN” is given, followed by “WHOOP WHOOP PULL UP” ifthe condition worsens. Visual alert warnings are provided by the illumi-nation of the red PULL-UP light (located on the forward flight instrumentpanel).

• Mode 3 Altitude Loss after Takeoff or Go-around: Provides auraland visual warnings for excessive altitude loss after takeoff or go-around. This mode is automatically activated between 50’ - 700’ radioaltitude and when the flaps or landing gear are retracted. The auralwarning of “DON’T SINK” is given and the visual alert warnings areprovided by the illumination of the red PULL-UP lights (located on theforward flight instrument panel).

• Mode 4 Unsafe Terrain Clearance - Not in Landing Configuration:Provides aural and visual warnings for unsafe terrain clearance withreference to airspeed, landing gear, and/or flap configurations. Thewarning of “TOO LOW GEAR - TOO LOW GEAR” is given when theaircraft is below 500’ radio altitude and 181 kts and the landing gear isnot down. The warning of “TOO LOW TERRAIN” is given when theaircraft is below 1000’ radio altitude and 235 kts, with the landing gearand/or flaps are not in the landing configuration. The warning of “TOOLOW FLAPS” is given when the aircraft is below 200’ radio altitude and149 kts, with the landing gear and flaps are not in the landing configura-tion. Visual alert warnings are provided by the illumination of the redPULL-UP lights (located on the forward flight instrument panel).

• Mode 5 Excessive Deviation below Glide Slope: Provides aural andvisual warnings for excessive deviation of 1.3 dots below the glideslope. The initial “soft” warning of “GLIDE SLOPE - GLIDE SLOPE ” isgiven when the aircraft is below 1000’ radio altitude and 1.3 dots ofdeviation. This “soft” warning increases in amplitude to a “loud” warning



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should the aircraft reach 2 dots of deviation and a radio altitude of lessthan 300’. This mode is automatically armed when valid signals arereceived from the Captain’s G/S receiver and with radio altitudes below1000’. Mode 5 can be cancelled or inhibited by pressing the BELOW G/S warning light only when in the “soft” range of alert. Visual alertwarnings are provided by the illumination of the red BELOW G/S lights(located on the forward flight instrument panel).

Center Console

Horizontal Stabilizer Trim Control System

• The controls for the horizontal stabilizer trim control system are locatedon the center console and on the control column wheel. The purpose ofthe stabilizer trim system is to provide longitudinal trim of the aircraft byvarying the angle of attack of the horizontal stabilizer. The trim systemconsists of the stabilizer connected to a jackscrew mechanical device.The jackscrew is controlled by two electric 115V AC actuators andmanual control cables. The maximum travel limit of the stabilizer isapproximately 17 units of trim.



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• The control of the stabilizer trim system is by the main electric actuator,via the thumb trim switches located on the control wheels. The autopilotactuator is controlled by the pitch channels and the speed trim system.As a backup source of control , manual cables have been provided, viathe manual trim wheel (located on the center console control stand).The system has been designed with trim cutout switches, that providesthe means of removing all electrical power from either control actuator.

• The stabilizer’s main electric trim actuator is a two speed motor. Trimspeed control is automatically determined by the position of the trailingedge flaps. The motor operates at low speed whenever the flaps areretracted and at high speed whenever the flaps are extended to anyposition. There is a maintenance limitation of actuator operation of 2minutes ON and 13 minutes OFF for possible overheating problems.This information is important to remember during runaway trim condi-tions. Over-trimming by pilots in the opposite direction may occur dur-ing this non-normal condition. Over-trimming has the potential to causeadditional overheat and/or stalled problems to occur. Thence, compli-cating the problems already existing.

Stabilizer Trim Band Range

• The stabilizer trim band range has been designed to provide a visualreference of stabilizer position in trim units. The green takeoff referenceband displays units of stabilizer trim that can be used for takeoffs. Trimpositions outside this green band range during takeoffs, will cause theaural takeoff warning horn to sound.



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Stabilizer Trim Wheel

• The stabilizer trim wheel is located on either side of the center consolecontrol stand. The wheel rotates anytime the stabilizer is in motion.Manual trimming of the stabilizer is accomplished by rotating the trimwheel in the desired direction, causing a chain assembly to move cablesconnecting the aft assembly of the stabilizer’s jackscrew gearbox cabledrum. Manual operation of the trim wheel, will cause the disengage-ment of both electric actuators.

Stabilizer Trim Cutout Switches

• The stabilizer trim cutout switches provides the means of removing allelectrical power from stabilizer’s electric actuators. Electrical power forthe stabilizer trim actuators is supplied from the No.2 115V AC transferbus and the No.2 28V DC bus.



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Speed Brake Lever

• The speed brake lever is located on the left side of the center consolecontrol stand. The speed brake lever has four positions available. TheDOWN detent position, positions all flight and ground spoiler panels tothe flared position. The ARMED position, arms the automaticspeedbrake system. Upon touchdown, the speedbrake handle movesto the UP position, and all flight and ground spoilers extend. TheFLIGHT DETENT position, extends all flight spoilers to their maximumposition for inflight use. The UP position, extends all flight and groundspoilers to their maximum position for ground use.

• Movement of the speedbrake lever, actuates the control assembly to allthe spoiler panels as speedbrakes. Cables run aft to the spoiler mixerand ratio changer in the right main gear wheel well. The spoiler panelsare positioned to any position between 00 - 400. The speedbrakesystem has been designed with a device called the "speedbrake leverno-back brake". This device prevents the speedbrake lever from beingrepositioned by vibration or from cable feedback movement.

Parking Brake Lever

• The park brake lever provides the means of setting the parking brake.This can be accomplished from either the Captain's or First Officer'srudder pedals by depressing the brake pedals fully and pulling the brake



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lever on the control stand. To release the parking brake, the Captain'sor First Officer's brake pedals need only be applied, causing the brakebell cranks to disengage.

• The parking brake shutoff valve is installed in the return line brake line,between the four anti-skid control valves. The purpose of the brakeshutoff valve is to prevent pressure bleed-off after the initial applicationof brake pressure. The battery switch must be selected ON prior tosetting the parking brake. Electric power for the brake valve operationis provided by 28V DC power from the battery bus.

Autothrottle Disengage Switches

• The autothrottle disengage switches-buttons are located on the outsideportion of both thrust lever handles. The selection of either switch, willautomatically disengage the autothrottle. The disengagement of the A/Twill cause the illumination of the red disengagement lights (light flashesintermediately) and the automatic movement A/T arm switch to the OFFposition. Selecting the disengage switch-button a second time, extin-guishes both A/T disengagement caution lights.

Takeoff/Go-Around Switches

• The F/D takeoff mode and autothrottle system activation are initiated bythe selection of the TO/GA button. The F/D mode directs the commandbars for 100 nose-down and wings level. At approximately 60 kts, thecommand bars moves to 150 nose-up. After wheels up, the commandbars provide guidance to maintain pitch that will achieve MCP speedplus 20 kts. During the takeoff roll and initial climb, F/D commandswings level. Normally at 400’, takeoff mode is terminated and LVL CHGis selected.

• Another function of the TO/GA button is the activation of autothrottlesystem. The selection of the TO/GA button during takeoff, initiatesthrust lever movement. The annunciator panels will reflect a change ofA/T status from ARM to EPR. The A/T system will automatically settakeoff thrust by 60 kts. The annunciator panel will indicate A/T statusof THR HOLD by 64 kts. Once THR HOLD has been obtained, onlymanual changes by the flight crew can be made to the thrust levers.The THR HOLD function will be maintained until 400’ RA (approximately18 seconds after takeoff).



B737-200

• During the execution of a go-around, the selection of the TO/GA buttonwill initiate the go-around mode of the autothrottles. The thrust leverswill advance to a “REDUCED” go-around thrust setting. This provides aclimb rate of approximately 1000’ to 2000’ fpm. The flight mode annun-ciator panel will display GA for A/T. If maximum go-around thrust isrequired by the flight crew, pressing the TO/GA button a second time,will advance the thrust levers to a “MAX” go-around thrust setting. Thego-around mode of the autothrottle system automatically arms whendescending below 2000’ RA.

Thrust Levers

• The thrust levers are connected to the fuel control unit by various con-trol cables and linkage. They are designed with a lockout mechanismthat prevents simultaneous actuation of the forward and reverse thrustlevers. The reverse thrust levers are attached to the top section of theforward thrust levers. As with the forward thrust levers, the reversethrust levers provides the means to control the variable fuel supply forreverse thrust operations.

• The reverse thrust levers have incorporated a detent position thatserves as a warning of approaching temperature limits range. Move-ment past this detent position, may cause an engine over-temperaturecondition to occur.



B737-200

Flap Lever

• The flap lever controls the operation of the flap drive system, throughthe use of a cable drum system. The major components of the flapdrive system are the trailing edge control valve, leading edge flap con-trol valve, hydraulic drive motor, flap load limiter, and the alternate drivemotor. The TE and LE control valves regulates hydraulic pressure tovarious hydraulic motors and actuators for flight control operation. Incase of flap lever cable breakage, the system has been designed toallow the flap bypass valve to prevent hydraulic pressure from operatingthe flap power unit.

• A flap load limiter has been designed to provide protection for the trail-ing edge flaps against excessive airspeeds. The system is activatedwhen the flap control lever is moved to the 40 position. When theinternal airspeed switches close at excessive airspeed, the trailing edgeflap control valve is positioned to retract the flaps to 30 units. The flapswill automatically return to the 40 position when the airspeed decreasesto a range of 147 to 157 kts.

• (Oral Topic) Flap gates have been provided to help prevent inadvertentflap lever movement past predetermined reference points. The flapgate at position 1 provides a “reference check point” for single enginego-arounds. The flap gate at position 15 provides the “reference checkpoint” for normal two engine go-arounds. The flap lever is also de-signed with a spring-loaded lock system which locks the handle in eachdetent position, thus providing another means to help prevent inadvert-ent flap lever movement.

Start Levers

• The start levers are located on the front of the center control stand. Thetwo-position levers, provide control for the fuel flow system and theignition circuits. Movement of the lever to the IDLE position, allows theopening of the main fuel shutoff valve in the fuel control unit. Theignition circuits are also energized with this selection. Movement of thelever to the CUTOFF position, closes the main fuel shutoff valve andde-energizes the ignition circuit.



B737-200

• During engine starting, the start levers are positioned to IDLE detent atapproximately 20% N2 or when maximum motoring has been achieved.The definition of maximum motoring is when N2 acceleration is lessthan 1% in 5 seconds. The minimum N2 starting value for start levermovement is 15% N2. Refer to your AFM.

Stabilizer Brake Release Knob

• The flight control system has been incorporated with a stabilizer brake.The purpose of the stabilizer brake is to stop unwanted trim motion.This is accomplished by the pilot moving the control columns oppositeto the trim motion, thus engaging the stabilizer brake. The stabilizerbrake release knob has been added to the system to help release thisbraking action. The stabilizer brake can also be released by reversingthe trim direction.

Rudder Trim Wheel

• The single rudder trim control wheel is located on the aft center console.Rotating the trim wheel in either direction repositions the rudder feel andcentering mechanism, which causes a shift in the rudder neutral controlposition. The rudder pedals are displaced proportionately.



B737-200

Aileron Trim Wheel

• The single aileron trim control wheel is located on the aft center con-sole. Rotating the trim wheel in either direction repositions the aileronfeel and centering mechanism, which causes a shift in the aileron neu-tral control position.

Service Interphone Handsets

• The service interphone handsets are located on the center controlstand. The handset is primarily used for communications with the cabincrew. The selection of the service interphone switch to the ON position,provides the connection of the handset for communications with anyexternal jack.

• The service interphone system may be inoperative for dispatch providednormal, alternate, and emergency communication procedures are estab-lished. The PA system must also be operational with an inoperative.Many airlines have procedures outlined in their flight operations manualspertaining to the use of a chime code system. This chime code systemmay be used as an alternate procedure, as mentioned above. Refer toyour MEL.

PA Hand Microphone

• The PA hand microphone is located next to the service interphonehandset. The primary purpose of the PA hand-held microphone is toprovide the capability of making direct PA announcements to the cabin,thus bypassing the audio selector panels.

• The passenger address system may be inoperative for dispatch pro-vided normal, alternate, and emergency communication procedures areestablished. The flight deck and cabin interphone system must beoperational. Refer to your MEL.



B737-200

Overheat Detector Switch

• The overheat detector switches are located on the upper left corner andon the upper right center of the fire protection panel. Each switch hastwo positions, NORM or FIRE. The selection to the NORM positionconnects the associated engine overheat detector system to the respec-tive amber ENG 1 or ENG 2 OVERHEAT light. The selection to theFIRE position connects the associated engine overheat detector systemto the fire warning lights and bell.

• The overheat detector switch is normally used when the fire detectioncircuit is inoperative. The MEL directs the selection of the OVT DETswitch from the NORM position to the FIRE position, thus providing firewarning detection using the overheat circuit. With inoperative overheatdetection circuit inoperative, a fire test must be accomplished prior toeach takeoff. There is no circuit switching for overheat circuit malfunc-tions. One overheat detection system or fire detection system perengine may be inoperative for dispatch provided the operable system istested prior to each departure. Refer to your MEL.



B737-200

Overheat/Inop and Fire Test Switch

• The overheat inoperative and fire test switch is located directly belowengine’s No.1 overheat detector switch (left side of the fire protectionpanel). The spring loaded two-position switch is designed to perform asystem test of the overheat detection circuit, fire detection circuit, andthe APU fire detection circuit.

• The selection of the switch to the OVHT INOP position, initiates systemtesting of the overheat detector loops and causes the illumination of themaster caution light, OVHT/DET annunciator, ENG No.1/ENG No.2overheat lights, and the APU DET INOP lights.

• The selection of the test switch to the FIRE position, initiates systemtesting of the fire detector loops on both engines, the fire detector on theAPU, and the fire detector in the wheel well. The testing of the firesystem will cause the illumination of the master fire warning lights, ENGNo.1/ENG No.2 fire warning lights, APU warning lights and the wheelwell lights. In addition, the warning alarm bell sounds within the flightdeck, the APU horn sounds in the wheel well, and the APU fire warninglight in the wheel well illuminates by flashing.

Wheel Well Fire Warning Light

• (Oral Topic) The illumination of the red WHEEL WELL fire warning lightindicates a fire in the main gear wheel well may exist. The fire alarmbell will sound and the master fire warning light will illuminate. Manychecklists require immediate pilot action of lowering the landing gear.Examiners may ask: “at what speeds do you lower the landing gearduring wheel well fire conditions”. Extending the landing should alwaysoccur at airspeeds below 270 kts and/or .82M. The reason for extend-ing the landing gear is to induce additional airflow through the wheelwell, thus directing flames away from sensitive areas within the wheelwell. As a precaution, do not retract the landing until at least 20 minutesafter the wheel well fire light has extinguished itself, and then, only if itabsolutely necessary. It has been stated, lowering the landing geardeflects flames away from the fire detection loops, but the possibility ofthe fire condition may still exist after the extinguishing of the warninglight.



B737-200

• The wheel well fire detection loops are located in the ceiling area of themain wheel well. The detector is a fenwall metallic (thermistor) design.Should any portion of the detector is heated to temperatures above4000F, the thermistor detector will actuate the fire warning circuit. Elec-trical power for the wheel well fire detection system is provided by 28VDC from the battery bus and 115V AC from the No.1 transfer bus.

• The wheel well fire detection system may be inoperative for dispatchprovided the brakes are inspected and are cool prior to departure. Re-fer to your MEL.

Engine Fire Warning Handles

• The engine fire handles are normally placed in the locked and downposition until emergency pilot action requires the pulling of the firehandle. The fire handles are locked down until the engine overheatand/or fire warning circuits detects a change of temperature around therespective detector. The associated engine’s handle will then automati-cally unlock, allowing free movement of the handle by the pilot. Thesystem has also been designed with a manual override plunger, allow-ing manual unlocking of the fire handles. The manual override plungeris located under the respective handle.

• (Oral Topic) Pulling the engine fire handle initiates the following actionsto occur:1. Arms the fire extinguisher circuit.2. Closes the fuel shutoff valves.3. Closes the bleed air valves.4. Closes the thrust reverser shutoff valve.5. Closes the hydraulic shutoff valve.6. Trips the generator control relay and breaker (after 7 sec. delay).7. Deactivates the hydraulic low pressure light.

• (Oral Topic) After pulling the associated engine fire handle to the UPposition, the handle is then rotated either LEFT or RIGHT to dischargethe respective extinguishing bottle. Discharging the bottle is accom-plished by electrically “firing” a squib, that punctures the seal of thebottle. This allows freon agent to be discharged around the exteriorsections of the engine. After discharging the fire bottle, the green bottledischarge light will illuminate, indicating the bottle has been successfullydischarged.



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• The fire extinguisher bottles contains an extinguishing agent (freon),that is pressurized with nitrogen to 800 psi (at 700F). If a bottle tem-perature reaches 2660F, the bottle will automatically discharge into thewheel well area. This is indicated by a ruptured disc.



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APU Fire Handle

• The APU fire handle is normally placed in the locked and down positionuntil emergency pilot action requires the pulling of handle. The firehandle is locked down until the APU fire warning circuits detects a firearound the sensor. Upon sensing the fire condition (4000F APU engine,7750F APU tailpipe area, or 4300F APU exhaust), the APU fire handlewill automatically unlock, allowing free movement of the handle by thepilot. The system has also been designed with an manual overrideplunger, allowing manual unlocking of the APU fire handle.

• (Oral Topic) Pulling the APU fire handle initiates the following actions tooccur:1. Arms the fire extinguisher circuit.2. Closes the APU fuel shutoff valve.3. Closes the bleed air valve.4. Closes the APU inlet door.5. Trips the APU generator control relay and breaker.

• (Oral Topic) After pulling the APU fire handle to the UP position, thehandle is then rotated either LEFT or RIGHT to discharge the singleextinguishing bottle. Discharging the bottle is accomplished by electri-cally “firing” a squib, that punctures the seal of the bottle. This allowsfreon agent to be discharged into the APU shroud area. After discharg



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ing the fire bottle, the green bottle discharge light will illuminate, indicat-ing the bottle has been successfully discharged.

• The APU extinguisher bottle is located in the fuselage, just foward of theAPU compartment. The spherical bottle is pressurized with nitrogen to600 psi at 700 F. Electrical power for discharging is provided from thehot battery bus.

• Two bottle discharge indicators are provided to give exterior visualwarnings of a normal discharge and high pressure discharges of theextinguisher bottle. A missing YELLOW discharge disc from the rearsection of the aircraft indicates the APU extinguisher bottle has beendischarged. A missing RED discharge disc, indicates bottle tempera-ture has exceeded 2660 F thus discharging the contents. This is some-times referred as a thermal discharge of the extinguisher bottle.

Engine Fire Warning Light

• The red engine fire warning lights is located within the engine firehandles. The illumination of the respective warning light indicates theassociated fire detection circuit has detected a fire condition (6000F).The fire warning system also includes aural warnings produced by thefire bell and additional visual warnings, as provided by the illumination ofthe master fire warning lights located on the glare shield panel. Press-ing either master fire warning light will silence the fire alarm bell andextinguish the master fire warning lights.

APU DET INOP Light

• The illumination of the amber APU DET INOP (detector inoperativelight) indicates a malfunction within the APU fire detection system hasoccurred. The master caution lights and the OVHT/DET annunciatorlights will also illuminate.

• The fire detection system has been designed with fire short circuit dis-criminators. This provides protection against false fire warnings withinthe APU fire detection system. The testing of this function has beenprovided by the selection of the overheat/inop and fire test switch.Movement of the switch to the OVHT-INOP test position, illuminates theAPU DET INOP light.



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• During ground and flight operations, the illumination of the APU DETINOP light requires the selection of the APU switch to the OFF position.It is recommended that you do not operate the APU, since an APU firewould not be detected and the APU would continue to operate.

APU Bottle Discharge Light

• The illumination of the amber APU BOTTLE DISCHARGE light indicatesthe fire extinguishing agent has been discharged into the APU shroudarea. The light is part of a pressure switch system, whenever bottlepressure decreases below 250 psi, the pressure switch closes and illu-minates the amber light.

Engine Bottle Discharge Light

• The illumination of either the L BOTTLE DISCHARGE light or the RBOTTLE DISCHARGE light indicates the fire extinguishing agent hasbeen discharged from the respective fire bottle. The light is part of apressure switch system, whenever bottle pressure decreases below 250psi, the pressure switch closes and illuminates the amber light.



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• The engine’s spherical extinguisher bottles are located in the wheel wellarea. The bottles are pressurized with nitrogen to 800 psi (at 700F).Electrical power for discharging is provided from the hot battery bus.Should bottle temperatures exceed 2660 F, the extinguisher bottles willdischarge into the wheel well area, thus causing the RED blowout discto be blown free. This provides an external warning of an thermaldischarge of the extinguisher bottles. Anytime a bottle has been dis-charged, the bottle must be completely replaced by maintenance.

Fire Warning Bell Cutout Switch

• The fire warning bell cutout switch has been provided to silence the firebell and APU horn. The selection of this switch also cancels the masterfire warning lights. A common oral question is referenced to the twoways of silencing the fire bell and/or APU horn. Pushing either the FIREWARN light or the bell cutout switch will silence the aural warnings.

Extinguisher Test Switch

• The extinguisher test switch has been provided to test the bottle dis-charge circuits of all fire extinguisher bottles. This test feature alsoincludes checking the engine selector valves for proper integrity.

Extinguisher Test Lights

• (Simulator Hint) The illumination of the green extinguisher test lightsindicates the discharge circuits are normal. During training, instructorswill fail the respective light during preflight. This indicates an malfunc-tion within the discharge circuit has occurred. Failure to bring thismalfunction to his/her attention, may cause future problems for youlater, during engine fire drills and maneuvers. Thoroughly check theoperation of ALL test lights when performing the test.



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Chapter II

AircraftLimitations

Review



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Operational Limitations

Maximum Operating Altitude

• 37,000’ (35,000’ as installed).

Maximum Takeoff Altitude

• 8300’.

Revenue Flights - Retention Bar Use

• During taxi, takeoff and landing, the escape slide retention bar must beinstalled.

Maximum Recommended Wind for Airstair Operations

• 40 kts.

Runway Slope Limits

• plus/minus 2%.

Maximum Takeoff - Landing Tailwind

• 10 kts.

Maximum Speed

• Observe Vmo pointer and gear/flaps placards.



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Turbulent Airspeeds

• 280 kts and .70 MACH

Flight Crew

Minimum Flight Crew

• Minimum flight crew consists of one pilot and one copilot.

Landing Gear Limitations

Brake Application

• Do not apply brakes until after touchdown.

Autobrakes Use

• The autobrakes must be selected to RTO or OFF for takeoff.

Landing Gear Towing - Hydraulic Pressurization

• System A hydraulic pressure must be depressurized for towing.

Type of Airplane Operation

Types of Airplane Operations

• The airplane is eligible for the following types of operation when re-quired equipment is installed and approved in accordance with theapplicable regulations:

VFR.Night flight.Instrument (IFR).Icing conditions.Over water operations.



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Air Conditioning & Pressurization Limitations

Maximum Differential Pressure

• 8.65 psi.

Operating Differential Pressure

• 7.5 +/- .1 psi.• 7.8 +/- .1 psi (check for correct controller)

Maximum Takeoff - Landing Cabin Differential

• .125 psi

Autopilot - Flight Director System Limitations

Autopilot Use With Depressurized Hydraulics

• (SP77) Use of autopilot pitch channel above .81 Mach is restrictedwith hydraulic system A or B depressurized.

Autopilot Use During Takeoff

• (SP77) Use of autopilot not authorized for takeoff or landing.

• (SP177) Use of autopilot not authorized for takeoff.

Autopilot Roll Channel Restrictions

• (SP77) Do not use autopilot roll channel above 30,000’ with yawdamper inoperative.



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Altitude Mode Use Restriction

• Do not use ALT HOLD mode when Captain's alternate static source isselected.

Autopilot Single Channel Operation Restrictions

• (SP177) For single channel operations, the autopilot shall not beengaged below 50’ AGL.

Autopilot Operational Procedures

• Flight crews must check MCP settings after any electrical power inter-ruptions.

• Flight crews must check ALT display to ensure desired altitude isdisplayed, following changes in the ALT selection in the MCP window.

• Flight crews must closely monitor altitude during all altitude changes toensure that the autopilot captures and levels off at the desired altitude.

• Flights should use standard callouts, crew coordination, and cross-checking techniques to detect any non-selected MCP display changes.

Performance Data Computer System Limitations

PDCS Requirements

• Do not use the PDCS information unless the engine configuration dis-played on the PDCS is the same as the engine configuration of theairplane.



B737-200

Fuel Management & Range Requirements

• Fuel management and range calculation valves presented by the PDCShave not been evaluated by the FAA.

Verification of EPR Values

• Verify that the representative takeoff EPR limits displayed on the CDUand EPR indicators agree with the predetermined limits obtained fromthe flight manual.

Hydraulic Power Limitations

Minimum Fuel For Hydraulic System B Ground Operations

• 1,676 lbs in TANK #2

Flight Control Limitations

Maximum Flap Extension Altitude

• 20,000’.

Minimum Recommended Altitude For Speedbrake Usage

• 500’.

Alternate Flaps Duty Cycle For Flight Operations

• One cycle, 25 minutes OFF.



B737-200

Electrical Power Limitations

TR Voltage Range

• 24 - 30V

Battery Voltage Range

• 22 - 30V

Maximum CSD Oil Temperature

• 1570C

Maximum CSD Oil Temperature Rise

• 200C

Ice & Rain Protection Limitations

Engine Thermal Anti-ice Requirements

• Engine TAI must be ON when icing conditions exist or are anticipated,except during climb and cruise below -400C SAT.

Wing Thermal Anti-ice System Requirements

• Do not operate wing anti-ice on the ground when OAT is above 100C(500F).

Maximum Speed - Altitude With Window Heat Inoperative

• 250 kts below 10,000’.



B737-200

Pitot Heat Requirements

• Pitot heat must be ON for takeoff.

Minimum N1 RPM During Icing Condition

• Minimum N1 RPM for operating in icing conditions except for landingexcept for landing:

TAT between 00 and 100C 40% N1TAT below 00C 55% N1TAT below -6.50C (moderate/severe) 70% N1

Fuel Limitations

Maximum Fuel Temperature

• 490C.

Minimum Fuel Temperature - Freeze Point

• Fuel freeze point + 30C.

Maximum Wing Tank Fuel Quantity

• 10,120 lbs (each).

Maximum Center Tank Fuel Quantity

• 16,351 lbs.

Maximum Allowable Fuel Imbalance - Flight Operations

• Maximum allowable fuel imbalance between tanks No.1 & No.2 is 1500lbs for taxi, takeoff, and flight.



B737-200

Maximum Allowable Fuel Imbalance - Landing

• Maximum allowable fuel imbalance between tanks No.1 & No.2 is 1300lbs for landing.

Fuel Loading Requirements

• Main tanks No.1 and No.2 must be full if the center tank contains morethan 1000 lbs. With less than 1000 lbs in the center tank, partial maintank fuel may be loaded, provided the effects of balance have beenconsidered.

Fuel Usage Requirements

• Use center tank fuel to depletion, followed by main tank fuel.

Fuel Type Requirements

• Always communicate with maintenance control before adding any typeof fuel other than JET-A.

Navigational Equipment Limitations

Weather Radar Requirements

• Do not operate weather radar during fueling, near fuel spills, or people.

Weather Radar Warm-up Requirements

• Warm up radar in STBY position only.



B737-200

HF Radio Requirements

• Do not operate HF radios during fueling or near fuel spills.Gross Weight & CG Limitations (as installed)

Maximum Taxi Weight

• 117,500 lbs.

Maximum Takeoff Weight

• 117,000 lbs (may be further restricted by takeoff, enroute, and landingperformance).

Maximum Landing Weight

• 107,000 lbs (may be further restricted by field length or climb limit.

Maximum Zero Fuel Weight

• 95,000 lbs.

C.G. Limits

• Must use an approved weight and balance system.

Maximum Inflight Weight

• Flaps 0 116,500 lbs.

• Flaps 30/40 106,000 lbs.



B737-200

Pneumatics Limitations

Maximum External Air Pressure

• 60 psig.

Maximum External Air Temperature

• 2320C (4500C).

Configuration Deviation List

Missing Airframe & Engine Parts

• When operation is schedule with certain secondary airframe and engineparts missing, the airplane must be operated in accordance with thelimitations specified in the basic airplane flight manual, and asamended by the CDL Appendix.

Flight Maneuvering Load Acceleration Limits

Load Accelerations Limitations

• Flaps Up: +2.5g to -1.0g.

• Flaps Down +2.0g to -0.0g.

TCAS Limitations

TCAS Weather Approval

• TCAS is approved for use in IMC and VMC conditions.



B737-200

TCAS Compliance Requirements

• Compliance with TCAS resolution advisories is required unless thePilot-In-Command determines that doing so would jeopardize the safeoperation of the flight.

TCAS Response Maneuvers

• Maneuvers in response to a TCAS resolution advisory which are in theopposite direction of that recommended by that advisory are prohibited.Certain circumstances may require a change in aircraft configuration orengine power setting in order to comply with a TCAS resolution advi-sory. Consult the aircraft operating manual, cockpit operating manual,or the approved flight manual for the aircraft specific situations.

Power Plant Limitations

Minimum Engine Starting Pressures

• Minimum pneumatic starting pressures of 30 psig (SL - decreasing onehalf psig per 1000’ above SL) are required prior to starter engagement.

Reverse Thrust Usage

• Reverse thrust for ground use only, intentional use of reverse thrustinflight is prohibited.

Ignition Requirements for Takeoff & Landing

• Ignition is required to be ON for takeoff, landing, and during engine anti-ice operations.



B737-200

High Intensity Ignition Duty Cycle (A or B, IGN L or IGN R, or FLT)

• Optimum life 10 minutes ON, 20 minutes OFF

Low Intensity Ignition Duty Cycle (as installed)

• FLIGHT: 2 minutes ON, 3 minutes OFF; 2 minutes ON, 23 minutesOFF.

• LOW IGN: Continuous

Starter Duty Cycle

• Normal Start: 30 seconds ON, 60 seconds OFF.

• Slow Start: 60 seconds, 60 seconds OFF (2 cycles only, then 5minutes cooling).

• Motoring: 2 minutes ON, 5 minutes cooling (fuel off).



B737-200

Maximum EGT Limits

(Condition/Rating) -9 -9A -15

• Acceleration n/a n/a 6300C

• T/O (5 MIN) 5800C 5900C 6200C

• MAX CONT 5400C 5450C 5800C

• GND START (>150C) 4200C 4200C n/a

• GND START (<150C) 3500C 3500C n/a

• GND START (JT8D-15) n/a n/a 5500C

• FLT START (JT8D-15) n/a n/a 6200C

Maximum N1 Limits

• MAX N1 100% 100% 102.4%

• MAX N2 100% 100% 100%

Oil Pressure Limits

• OIL PRESS (MAX) 55 PSI same same

• OIL PRESS (MIN) 40 PSI same same

Oil Temperature

• OIL TEMP (MAX) 1570C 1570C 1650C

• OIL TEMP(15 min.) 120-1570C 120-1570C 130-165C

• OIL TEMP (CONT) 1200C 1200C 1300C



B737-200

APU Limitations

Maximum APU EGT

• MAX EGT 7600C

• MAX CONT. EGT 7100C.

APU Bleed Air & Electrical Power Usage

• The maximum altitude for using APU bleed air and electrical power is10,000’.

APU Bleed Air Usage

• The maximum altitude for using APU bleed air is 17,000’.

APU Electrical Power Usage

• The maximum altitude for using APU electrical power is 35,000’.

APU Maximum Operating Altitude

• The maximum altitude for operating the APU is 35,000’.

APU Bleed Valve Position

• The APU bleed valve must be in the CLOSED position when:a. Ground air connected and isolation valve open.b. LH engine bleed valve open.c. Isolation valve and right engine bleed valve open.

APU Bleed Valve Position - Starting

• APU bleed valve may be open during engine start, but avoid enginepower above idle.



B737-200

Classroom Notes:



B737-200

Chapter III

AircraftSchematics

Review



B737-200

Dimensions Schematic



B737-200

Electrical System Schematic



B737-200

Fire Protection Schematic



B737-200

Flight Controls - Leading Edge Devices Schematic



B737-200

Flight Controls - Roll Control Schematic



B737-200

Flight Controls - Elevator Control Schematic



B737-200

Fuel System Schematic



B737-200

System A Hydraulics Schematic



B737-200

System B Hydraulics Schematic



B737-200

Standby Hydraulic System Schematic



B737-200

Ice & Rain System Schematic



B737-200

Main Landing Gear Schematic



B737-200

Nose Gear Schematic



B737-200

Compass System Schematic (As Installed)



B737-200

Attitude System Schematic (As Installed)



B737-200

Pneumatics Schematic



B737-200

Ram Air Schematic



B737-200

Pressurization Schematic



B737-200

Engine Fuel & Oil System Schematic



B737-200

Thrust Reverser Schematic



B737-200

Auxiliary Power Unit Schematic



B737-200

Brake System Schematic



B737-200

Flight Training ProfilesPrecision Approach

Approach Callouts• PNF "Outer Marker___Ft,

(crossing altitude) crosscheck".• PF “Target___”.• PNF "500' Above DH, On

Target ___+/-,Sink___ft."• PF "100' (Look Left-Right,

Look Ahead)• PNF “Minimums”• PF “Landing, Missed

Approach”.

Missed Approach Profile• Call “G/A Thrust”.• Rotate to G/A attitude

(approx. 15o)• "Flaps 15, Positive Rate,

Gear Up"• Climb at Vref + 15 kts to

1000’.• Tune Radios.• At 1000’, retract flaps on

schedule.

190 kts “Flaps 1”

Before OM“Flaps 30/40”

G/S 1-dot“Flaps 25”


Glide Slope Alive“Gear Down,

Flaps 15, Landing Ck”

One Engine Inop Approach Profile

• Plan Flaps 15 landing. Monitor fuel balance.• Glide Slope alive, call "Flaps 15, Engine Inop Landing Check".• Center rudder trim prior to landing.

One Engine Inop Missed Approach

• Slowly advance thrust lever to G/A thrust.• Rotate to G/A attitude.• Call "Flaps 1, Positive Rate, Gear Up"• Climb at Vref 15 + 5 kts to flap retraction altitude (1000’ AGL).• Call “Tune Radios”.• Accelerate, retract flaps on schedule, call “Max. Continuous Thrust”.• Limit Bank to 15o until reaching Vref + 15 kts.



B737-200

Takeoff ProfileNormal & Engine Failure

Takeoff Roll:

• Complete all takeoff checklists.• Advance thrust levers, engines stabilized: Press TOGA switch or

manually advance thrust levers. Maintain directional control withsteering tiller and rudders, standard crosswind techniques apply.

• Rotate at 30 per second.• Target Attitude: 150 (200)

• Engine Failure: Prior to V2, accelerate to and maintain V2. After V2,maintain speed attained at time of failure. Above V2+25, increase pitchto maintain V2+25. Limit bank angle to 15o until reaching V2+15. Tar-get Attitude: 110 (200)

400’ AGL

• Call “Heading Select”.• Obstacle Clearance: Compliance with applicable SPECIAL departure

procedure is mandatory. In absence of a SPECIAL departure, do notturn prior to 1000’ AGL unless WX is greater than 1000-3.

• Engine Failure: Initiate Emergency Procedures above 400’ AGL.

1000’ AGL

• Begin acceleration to BSEC. Retract flaps on speed schedule.• Engine Failure: Accelerate in level flight if engine fails. Complete

checklist, tune radios, notify ATC, check weather, notify cabin crew, andnotify company.

3000’ AGL

• Normal profile is completed. Enter climb phase of flight.



B737-200

Takeoff ProfileNormal & Engine Failure Callouts

Takeoff Roll:

• PF "Set___(EPR/N1)”.• PNF "80 kts, Thrust & Flaps Normal, V1- Rotate".• PF "Positive Rate - Gear Up".• Engine Failure:

PF “Max. Power - Positive Rate - Gear Up”.

400’ AGL

• PF “Heading Select”.• Engine Failure:

PF ‘What’s the Problem”.PF “Memory Items”.

Use the word “CONFIRMED”, to confirm and authorize PNF’smemory items and immediate actions.

(Example: Engine Fire, Severe Damage, or Separation Cklist)• PF “Memory Items”.• PNF “No.1 Thrust Lever Close, Confirm the No.1 Thrust Lever”• PF “Confirmed - Close”

1000’ AGL

• PF “Climb Thrust, Flaps 1 (or Flaps Up), Set BSEC”• Engine Failure: (Retract flaps on speed schedule).

PF “Flaps Up, Set Max. Continuous Thrust”.PF (Calls one of the following checklist)

“Engine Fire, Severe Damage or Separation”“Engine Failure and Shutdown”

“Inflight Engine Start”3000’ AGL

• PF “Vnav - After Takeoff” or “After Takeoff Checklist”



B737-200

Takeoff ProfileRejected Takeoff

Note: V1 is an action point. Engine failure has already been recognized andACTION initiated by the Captain to either CONTINUE or REJECT thetakeoff.

The Captain will simultaneously:

• Apply maximum manual brakes.• Close throttles and disengage A/T.• Manually Extend speedbrakes.• Apply maximum reserve thrust consistent with weather conditions.

The First Officer will:

• Verify all Captain required actions and call out any omissions.• Call out 60 kts.• Notify tower.

Note: If another takeoff is planned, review Brake Cooling Schedule.

As appropriate, Captain calls to cabin: “Remain Seated”

“Engine Failure” or other abnormality.“REJECT”



B737-200

Flight Training ProfilesClean Stall

Initial Setup

• Flight Director OFF.• Set Vref for Flaps 40.

• Set EPR/N1 Bugs to G/A.

Flight Requirements

• Power OFF above 210 kts.

• Trim to 210 kts.

• At 180 kts set 50% N1.

• Hold heading and altitude.

Recovery

• Advance thrust levers.

• Call "G/A Thrust".

• Minimum loss of altitude.

• Return to entry altitude.

• Maintain 210 kts.



B737-200

Flight Training ProfilesTakeoff Stall

Initial Setup



Recovery




• Roll wings level.

• Accelerate to Vref, call “PositiveRate - Gear Up”.


• Flap retraction on schedule.


Flight Requirements


• Extend flaps on schedule.

• At 170 kts, “Gear Down, Flaps15, set 60% N1”.

• Enter a 250 bank.

• Trim to 150 kts.

• Hold altitude.

• Stall shaker at Vref.



B737-200

Flight Training ProfilesLanding Stall

Initial Setup



Flight Requirements


• Extend flaps on schedule.

• At 170 kts, “Gear Down, Flaps15, set 70% N1”.

• Trim to Vref plus 4 kts.

• Flaps 30 on speed schedule.

• Hold heading and altitude.

• Stall shaker at Vref minus 15.

Recovery




• Accelerate to Vref, call “PositiveRate - Gear Up”.


• Flap retraction on schedule.




B737-200

Flight Training ProfilesPrecision Approach

Approach Callouts• PNF "Outer Marker___Ft,



Look Ahead)• PNF “Minimums”• PF “Landing, Missed

Approach”.Missed Approach Profile• Call “G/A Thrust”.• Rotate to G/A attitude




schedule.

Glide Slope Alive“Gear Down,



Before OM“Flaps 30/40”

170 kts “Flaps 5” G/S 1-dot

“Flaps 25”


• Plan Flaps 15 landing. Monitor fuel balance.• Glide Slope alive, call "Flaps 15, Engine Inop Landing Check".• Center rudder trim prior to landing.





B737-200

Flight Director Approach

(Oral Topic) Flight director approach profile:• Arm VOR/LOC when cleared to intercept localizer.• Arm AUTO APP or APP mode after localizer intercept (see caution).• MCP: Engage second A/P in CMD for dual channel approach.• If FLARE does not arm by 800' RA, second autopilot will disconnect.• 400' RA - Nose up pitch bias introduced (dual channel).• 50' RA - disconnect A/P if single channel.Intercept roll mode can be:• VOR/LOC, HDG SEL, LNAV, CWS ROLL (as installed). After LOC

capture, HDG bug may be set as desired.In order to arm and capture the localizer the respective:• F/D and /or autopilot must be ON (A/P in CMD) (as installed).• VHF NAV radio must be manually tuned to the ILS frequency.• (300) Non-EFIS "NAV" switch must be in VOR/ILS (raw data), EFIS

equipment may remain in the MAP mode using the ADI raw data pre-sentation.

• Select AUTO APP or APP mode to arm the glide slope capture feature.Mode control panel:• Dual channel autopilot capability is achieved by engaging the second A/

P in CMD after the APP mode is selected and the above criteria is metfor the second NAV radio prior to 1500' RA. Both autopilots mustremain engaged for autoland.

Glideslope capture (2/5 dot. green annunciation):• ALT ALERT deactivated (except SP-77), set as desired.• MCP Equipment must choose a pitch mode (LVL CHG, VERT SPD,

VNAV, CWS PITCH) to fly within 2/5 dot for G/S capture. SP-77 has amanual G/S capture feature..

• Capture from above the G/S may be abrupt, initially.• PDC A/T equipment engages in GA. (300 N1 Limit engages in GA).Dual channel flare/autolanding:• 50’ - 42’: Flare Active: F/D's bias out of view.• M27’: A/T or Pilot retards throttles to idle. Touchdown.• A/T disengages 2 seconds after touchdown (if in use).• Pilot disengages autopilot.• See your OPS SPECS for autoland restrictions.

Caution:Autopilot can capture and descend on the glide slope even though not onthe LOC. Ensure that descent on the G/S will meet all restrictions or waituntil "LOC ALIVE" before arming AUTO APP or APP mode.



B737-200

Flight Training ProfilesNon-Precision Approach

Approach Callouts• PNF "VOR, NDB ___Ft,



Look Ahead)• PNF “Minimums”• PF “Holding, Landing, Missed

Approach”.

Missed Approach Profile• Call “G/A Thrust”.• Rotate to G/A attitude




schedule.


• Plan Flaps 15 landing. Monitor fuel balance.• Inbound course alive, call "Flaps 15, Engine Inop Landing Check".• Center rudder trim prior to landing.



Inbound Course Alive“Gear Down,



Before FAF“Flaps 30/40”




B737-200

Flight Training ProfilesSteep Turns



B737-200



B737-200

Index



B737-200

Index

A

A/T Takeoff Mode, .................................................................................185AC Meters Selector, ................................................................................39AC Voltmeter, ..........................................................................................39Acceleration, .........................................................................................274Accumulator Precharge Pressures, ......................................................235Aileron Trim Wheel, ...............................................................................252Air Conditioning & Pneumatic Sensors Locations, ................................116Air Conditioning & Pressurization Limitations, ......................................264Air Conditioning & Pressurization: Class Notes, ...................................317Air Conditioning Pack Switch, ...............................................................125Air Conditioning System Annunciator, ...................................................182Air Mix Valve Indicator, ..........................................................................117Air Mix Valves, .......................................................................................118Air Temperature Source Selector, .........................................................117Air Temperature/True Airspeed Indicator, ..............................................234Air-Ground Sensor Failure, ...................................................................228Aircraft Defueling, ...................................................................................29Aircraft Refueling, ...................................................................................29Aircraft Schematics Review, .................................................................277Airplane General: Class Notes, .............................................................313Airspeed Cursor Control Knob, .............................................................194ALT/NORM Switch, ...............................................................................162Alternate Flap Master Switch, .................................................................18Alternate Flaps Duty Cycle For Flight Operations, ...............................266Alternate Flaps Operation, ......................................................................19Alternate Flaps Position Switch, .............................................................19Alternate Reverser Hydraulic Pressure, ................................................155Alternate Vertical Gyro, .........................................................................196Altimeter, ...............................................................................................197Altitude Alert Speaker, ..........................................................................151Altitude Display Indicator (SP177), .......................................................188Altitude Horn Cutout Switch, .................................................................101Altitude Loss after Takeoff or Go-Around, .............................................243Altitude Mode Use Restriction, .............................................................265



B737-200

Altitude Selector (SP177), .....................................................................188Anti-Collision Light Switch, ....................................................................139Anti-Ice System Annunciator, ................................................................181Anti-skid Control Switch, .......................................................................222Anti-skid Control Unit, ...........................................................................221Anti-skid Inop Lights, .............................................................................222Anti-skid System, ..................................................................................221Approach Callouts, .........................................................................307,309Approach Profile, ...................................................................................308APU Aborted Starts, ...............................................................................61APU Alternate Starting, ..........................................................................61APU Automatic Shutdown Protection, ....................................................63APU Bleed Air & Electrical Power Usage, ............................................275APU Bleed Air Switch, ..........................................................................131APU Bleed Air Usage, ...........................................................................275APU Bleed Valve Position, ....................................................................275APU Bleed Valve Position - Starting, ....................................................275APU Bottle Discharge Light, .................................................................259APU DET INOP Light, ...........................................................................258APU Electrical Power Usage, ................................................................275APU Exhaust Temperature Indicator, ......................................................59APU Fault Light, ......................................................................................57APU Fire, ................................................................................................64APU Fire Detection, ................................................................................65APU Fire Handle, ..................................................................................257APU Fuel Control Unit, ............................................................................62APU Fuel Control Unit (FCU), .................................................................62APU Generator Ammeter, .......................................................................59APU Generator Off Bus Light, ................................................................55APU Generator Switch, ...........................................................................55APU High Temperature Light, .................................................................57APU Horn, ......................................................................................178,254APU Inflight Start Attempts, ....................................................................61APU Inlet Area Icing, ..............................................................................64APU Limitations, .........................................................................56,62,275APU Low Oil Quantity Light, ...................................................................56APU Maximum Operating Altitude, .......................................................275APU Oil Pressure Light, ..........................................................................57APU Overspeed Light, ............................................................................58APU Related Notes, ................................................................................62APU Start Sequence, ..............................................................................60



B737-200

APU Start Switch, ...................................................................................60APU System Annunciator, .....................................................................181APU Winter Operations, ..........................................................................64Arcing-Delaminated-Shattered-Cracked Windows, .................................79ASP - ALT/NORM Switch (as installed), ...............................................162ASP - Amplifier Switch, .........................................................................162ASP - Filter Switch, ...............................................................................160ASP - MASK/BOOM Switch, ................................................................162ASP - Oxygen/Boom Switch, ................................................................161ASP - PTT Switch, ................................................................................160ASP - PTT Switch, ................................................................................162ASP - Receiver Switch, .........................................................................160ASP - Receiver Switch, .........................................................................161ASP - Transmit Light, ............................................................................160ASP - Transmitter Selector, ...................................................................160ASP - Transmitter Selector, ...................................................................162Asymmetric Flap Protection, .................................................................220Asymmetrical Flap Protection System Review, ......................................19Attendant Call Switch, .............................................................................72Attitude Director Indicator, ....................................................................196Attitude System Schematic, ..................................................................292Attitude Warning Flag, ...........................................................................196Audio Selector Panel (overview), ..........................................................159Auto Fail Light, ......................................................................................103Auto Mode Flight Profile, .......................................................................109Auto Mode Test, ....................................................................................114Auto Spoiler System, ............................................................................206Autobrake Inop Light, ............................................................................224Autobrake Select Switch, ......................................................................224Autobrake Selection Criteria, ................................................................224Autobrake System, ................................................................................223Autobrakes Use, ....................................................................................263Autolanding Procedures, .......................................................................308Automatic Flight: Class Notes, ..............................................................321Automatic Load Shedding, ......................................................................52Automatic Load Shedding Protection, .....................................................40Autopilot - Flight Director System Limitations, ......................................264Autopilot Actuator, .................................................................................245Autopilot Engage Paddles (SP177), ......................................................189Autopilot Operational Procedures, ........................................................265Autopilot Roll Channel Restrictions, ......................................................264



B737-200

Autopilot Single Channel Operation Restrictions, .................................265Autopilot Use During Takeoff, ................................................................264Autothrottle Arm Switch (SP177), .........................................................184Autothrottle Disengage Switches, .........................................................248Autothrottle System,..............................................................................185Auxiliary Power Unit, ......................................................................298,299Auxiliary Power Unit Schematic, ...........................................................298Auxiliary Power Unit: Class Notes, ........................................................325Auxiliary Vertical Gyro, ............................................................................23

BBattery, ....................................................................................................41Battery Charger, ......................................................................................42Battery Switch, ........................................................................................40Battery Voltage Range, .........................................................................267Bleed Air Valves, ...................................................................................255Bleed Trip Off Light, ..............................................................................128Boric Acid, ...............................................................................................41Brake Accumulator, ...............................................................................236Brake Application, .................................................................................263Brake Snubbers, ....................................................................................237Brake System Schematic, .....................................................................299Brakes, ..................................................................................................235Braking Action Reports, ........................................................................225Bus Transfer Switch, ...............................................................................51

CC.G. Limits, ...........................................................................................270C.V.R. (Cockpit Voice Recorder), ............................................................99C.V.R. Area Microphone, .........................................................................99C.V.R. Erase Switch, ...............................................................................99C.V.R. Headset Jack, ............................................................................100C.V.R. Monitor Indicator, .......................................................................100C.V.R. Monitor Light, .............................................................................100C.V.R. Test Switch, ................................................................................100Cabin Altimeter-Differential Pressure Indicator, ....................................101Cabin Altitude Indicator, ........................................................................106Cabin Altitude Selector, .........................................................................106Cabin Altitude Warning System, ...........................................................102Cabin Emergency Exit Lights, .................................................................70



B737-200

Cabin Rate of Climb Indicator, ..............................................................102Cabin Rate Selector, .............................................................................106Center Console, ....................................................................................244Center Tank Fuel Pump Low Pressure Light, .........................................32Center Wing Fuel Boost Pumps, .............................................................33Change-Over Switch (SP177), ..............................................................186Checklist Philosophy, ..............................................................................11Circuit Breaker Light Control, ..................................................................66Classroom Notes, .................................................................................311Clock, ....................................................................................................200Cockpit Call Light, ...................................................................................72Cockpit Lighting, ....................................................................................141Cockpit Voice Recorder (C.V.R.), ............................................................99Command Bars, ....................................................................................182Communications: Class Notes, .............................................................329Compass Synchronization, .....................................................................14Compass System Review, ......................................................................15Compass System Schematic, ...............................................................291Compass Transfer Switch, ......................................................................15Configuration Deviation List, .................................................................271Course Deviation Bar, ...........................................................................203Cowl Valve Open Light, ...........................................................................84Crew Oxygen Masks, ............................................................................167Crew Oxygen Shutoff Valve, .................................................................167Crew Portable Oxygen System, ............................................................170Crewmember Oxygen Panel, ................................................................168Crossfeed Selector, .................................................................................27Crossfeed Valve Open Light, ...................................................................27CSD - Constant Speed Drive, .................................................................44CSD Limitations, .....................................................................................46

D

Data Thumbwheel Switches, .................................................................174DC Ammeter, ..........................................................................................38DC Meters Selector, ................................................................................39DC Power APU Boost Pump, ..................................................................64DC Voltmeter, ..........................................................................................38Decision Height Light, ...........................................................................199Defueling, ................................................................................................29DG/Slaved Switch, ..................................................................................14



B737-200

Digital Total Air Temperature Indicator, .................................................232Digital Error Codes, ...............................................................................210Dim Entry Lights, ..................................................................................141Dimensions Schematic, ........................................................................278Documentary Data Thumbwheel Switches, ..........................................174Dome Light Switch, ...............................................................................158Doors System Annunciator, ..................................................................181Dripsticks, .............................................................................................208Dual Bleed Light, ...................................................................................121Dual Channel Autopilot, ........................................................................308Duct Overheat Light, .............................................................................118

EEGT Limitations, ...................................................................................214Electric Motor Pump Switch, ...................................................................90Electric Vertical Speed Indicators Displays, ..........................................203Electrical Power Limitations, .................................................................267Electrical System Annunciator, .............................................................180Electrical System Description, ................................................................37Electrical System Schematic, ...............................................................279Electrical: Class Notes, .........................................................................333Emergency Battery Packs, ......................................................................70Emergency Braking, ..............................................................................236Emergency Cockpit Lighting, ................................................................141Emergency Equipment: Class Notes, ...................................................337Emergency Exit Lights, ....................................................................69,141Emergency Exit Lights Not Armed Light, ................................................70Emergency Flood Lights, ......................................................................158Emergency Instrument Flood Light, ......................................................141Emergency Oxygen Lever, ....................................................................168Engine Anti-Ice Switch, ...........................................................................86Engine Anti-Ice Valve Light, ....................................................................85Engine Bleed Air Switch, .......................................................................129Engine Bottle Discharge Light, ..............................................................259Engine Driven Pump Switch, ..................................................................90Engine Fire Warning Handles, ..............................................................255Engine Fire Warning Light, ....................................................................258Engine Fuel & Oil System Schematic, ..................................................296Engine Fuel Shutoff Valves, ....................................................................24Engine Fuel Valve Closed Light, .............................................................24



B737-200

Engine Inlet Pressure (Pt2), ..................................................................211Engine Overheat Light, .........................................................................253Engine Pressure Ratio (EPR), ..............................................................211Engine Start Panel, .................................................................142,143,144Engine System Annunciator, .................................................................181Engine Thermal Anti-ice Requirements, ...............................................267Engine Vibration Levels, ..........................................................................86Entry Lights, ..........................................................................................141EPR Flight Mode Selector, ....................................................................235Equipment Cooling Off Light, ..................................................................68Equipment Cooling Switch, .....................................................................67Event Switch, ........................................................................................173Excessive Descent Rate, ......................................................................243Excessive Deviation below Glide Slope, ...............................................243Excessive Terrain Closure Rate, ...........................................................243Exhaust Gas Temperature, ...................................................................213External Power Contactors, ....................................................................37External Power Interphone Station, ......................................................157External Power Receptacle, ....................................................................49Extinguisher Bottles, .............................................................................260Extinguisher Test Lights, .......................................................................260Extinguisher Test Switch, ......................................................................260

FF Outflow Closed Light, .........................................................................123Fast Synchronization, ..............................................................................16Fast Synchronization Process, ...............................................................14FCOP - Emergency Lever, ....................................................................168FCOP - Oxygen Diluter Lever, ..............................................................168FCOP - Oxygen Flow Indicator, ............................................................169FCOP - Oxygen Supply Lever, ..............................................................169FDR - Event Switch, .............................................................................173FDR - Flight Recorder Test Switch, ......................................................171FDR - OFF Light, ..................................................................................174FDR - Recording Time Remaining Indicator, ........................................173FDR - Repeat Switch, ...........................................................................174FDR - Trip and Date Light, ....................................................................174FDR - Trip and Date Selectors, .............................................................173Feel Differential Pressure Light, ..............................................................21Fire Handle Override Plunger, .......................................................255,257



B737-200

Fire Protection Schematic, ....................................................................280Fire Protection: Class Notes, ................................................................341Fire Short Circuit Discriminators, ..........................................................258Flap Bypass Valve, ................................................................................250Flap Drive System,................................................................................250Flap Gates, ...........................................................................................250Flap Lever, ............................................................................................250Flap Load Limiter, .................................................................................250Flight Limitations, ...................................................................................19Flight Altitude Indicator, ........................................................................105Flight Altitude Selector, .........................................................................105Flight Control Limitations, .....................................................................266Flight Control Low Pressure Light, ..........................................................18Flight Control Switches, ..........................................................................16Flight Control System Annunciator, ......................................................180Flight Controls - Elevator Control Schematic, .......................................283Flight Controls - Leading Edge Devices Schematic, .............................281Flight Controls - Roll Control Schematic, ..............................................282Flight Controls: Class Notes, ................................................................345Flight Crew, ...........................................................................................263Flight Crew Oxygen Indicator, ...............................................................163Flight Crew Oxygen Masks, ..................................................................167Flight Crew Oxygen Shutoff Valve, .......................................................167Flight Crewmember Oxygen Panel, ......................................................168Flight Data Recorder, ............................................................................172Flight Director - Engine Failure, ............................................................183Flight Director Approach, ......................................................................308Flight Director Switch (SP177), .............................................................182Flight Director Takeoff Mode, ................................................................183Vertical Speed Indicator, .......................................................................203Flight Instruments: Class Notes, ...........................................................349Flight Interphone System, .....................................................................157Flight Maneuvering Load Acceleration Limits, ......................................271Flight Mode Annunciator Panels (SP177), .....................................190,191Flight Recorder Test Switch, .................................................................171Flight Spoilers, ......................................................................................205Flight Spoiler Switch, ..............................................................................20Flight Training Profiles - Clean Stall, .....................................................304Flight Training Profiles - Landing Stall, ..................................................306Flight Training Profiles - Non-Precision Approach, ................................309Flight Training Profiles - Precision Approach, .......................................307



B737-200

Flight Training Profiles - Steep Turns, ...................................................310Flight Training Profiles - Takeoff Stall, ...................................................305Flight Training Profiles Review, .............................................................385Flight Training Profiles Review, .............................................................300Flight-Ground Switch, ...........................................................................107Flow Control Valve, ...............................................................................113Flow Pattern Philosophy, .........................................................................11Fluid Balance Lines, .......................................................................238,239Flux Valve System, ..................................................................................14Flux Valves, ......................................................................................15,202Forward Instrument Panel, ....................................................................193Forward Outflow Valve, .........................................................................113Frequency Meter, ....................................................................................38Fuel Boost Pump Power Source, ............................................................31Fuel Compensator Leakage, .................................................................210Fuel Filter Icing Light, .............................................................................25Fuel Flow Indicator, ...............................................................................215Fuel Heat Switch, ....................................................................................26Fuel Heat Valve, ......................................................................................26Fuel Heat Valve Failure, ..........................................................................26Fuel Heat Valve Open Light, ...................................................................26Fuel Imbalance, .......................................................................................25Fuel Limitations, ....................................................................................268Fuel Loading Requirements, .................................................................269Fuel Management & Range Requirements, ..........................................266Fuel Pump Low Pressure Light, ..............................................................34Fuel Pump Switch, ..................................................................................34Fuel Quantity Indication System, ..........................................................208Fuel Quantity Test Switch - Analog Indicators, .....................................209Fuel Quantity Test Switch - Digital Indicators, ......................................210Fuel Shutoff Valves, ..............................................................................255Fuel Surge Tanks, ...................................................................................35Fuel System Annunciator, .....................................................................180Fuel System Limitations, .........................................................................25Fuel System Schematic, .......................................................................284Fuel Tank Capacities, ..............................................................................36Fuel Tank Transfer, ..................................................................................28Fuel Tanks, ..............................................................................................35Fuel Temperature Indicator, ....................................................................24Fuel Transfer, ..........................................................................................28Fuel Type Requirements, ......................................................................269



B737-200

Fuel Usage, .............................................................................................25Fuel Usage Requirements, ...................................................................269Fuel Valve Closed Light, ..........................................................................24Fuel Vent System, ...................................................................................35Fuel: Class Notes, .................................................................................353

G

Galley Power Switch, ..............................................................................40Gasper Fan Switch, ..............................................................................119Generator AC Ammeter, .........................................................................48Bus Off Light, ..........................................................................................52Generator Control Unit (GCU), ...............................................................55Generator Drive Disconnect Switch, .......................................................45Generator Drive High Oil Temperature Light, ..........................................45Generator Drive Low Oil Pressure Light, ................................................44Generator Drive Oil Temperature Indicator, ............................................46Generator Drive Temperature Switch, .....................................................46Generator Loss (dual), ............................................................................54Generator Off Bus Light, .........................................................................53Generator Switch, ...................................................................................54Glide Slope Warning Flag, ....................................................................203Go-around Mode, ...........................................................................186,249GPWS Flap/Gear Inhibit Switch, ..........................................................242GPWS Inoperative Light, ......................................................................241GPWS System Test Switch, ..................................................................241Greenwich Mean Time, .........................................................................200Gross Weight & CG Limitations, ...........................................................270Ground Call Switch, ................................................................................72Ground Interconnect Switch, ..................................................................88Ground Power, ........................................................................................49Ground Power Available Light, ................................................................48Ground Power Switch, ............................................................................48Ground Proximity Warning System, ......................................................242Ground Service Switch, ..........................................................................50Ground Service Bus, ...............................................................................51Ground Spoilers, ...................................................................................205

H HF Radio Requirements, .....................................................................270Heading Control Knob, ............................................................................14



B737-200

Heading Warning Flag, ..........................................................................202Heading/Bank Angle Selector (SP177), ................................................187Heater OFF Light, ...................................................................................80High Energy Continuous Ignition, .........................................................143High Energy Ignition, ...............................................................142,143,144High Intensity Ignition Duty Cycle, ........................................................273Horizontal Situation Indicator, ...............................................................202Horizontal Stabilizer Trim Control System,............................................244How To Use The Cockpit Review, ...........................................................11Hydraulic Brake Pressure Indicator, ......................................................235Hydraulic Malfunction: Loss of Standby System, ...................................97Hydraulic Malfunction: Loss of System A Pressure, ..............................95Hydraulic Malfunction: Loss of System B Pressure, ..............................96Hydraulic Malfunction: Manual Reversion, .............................................97Hydraulic Overheat Light - Electric Pump, ..............................................89Hydraulic Power Limitations, .................................................................266Hydraulic Pump Low Pressure Light, ......................................................89Hydraulic Shutoff Valve, ........................................................................255Hydraulic System, ...................................................................................91Hydraulic System A Fluid Quantities (200), ..........................................239Hydraulic System A Quantity Indicator, ................................................239Hydraulic System Annunciator, .............................................................181Hydraulic System B Low Quantity Light, ...............................................238Hydraulic System Pressure Indicator, ...................................................237Hydraulic: Class Notes, .........................................................................357

I

IAS/MACH Display (SP177), .................................................................186Ice & Rain Protection Limitations, .........................................................267Ice & Rain System Schematic, .............................................................288Ice & Rain: Class Notes, .......................................................................361Icing Conditions, ......................................................................................87Igniter Service Life, .................................................................142,143,144Ignition Requirements for Takeoff & Landing, .......................................272Inboard Brakes (200), ...........................................................................236Inboard Landing Lights, .........................................................................135Inoperative Flight Directors, ..................................................................183Inoperative Pack Valves, .......................................................................125Instrument Comparator Lights, .............................................................196Instrument Comparator System, ...........................................................195



B737-200

Instrument Comparator Test Switch, .....................................................195Instrument Vibrators, .............................................................................198Isolation Valve, ......................................................................................153Isolation Valve Light, .............................................................................153Isolation Valve Switch, ..........................................................................125

JJackscrew Mechanical Device, .............................................................244

KKidde Sensor Fire Detection Loop, .........................................................65

LL.E.D. Amber Transit Lights, ..................................................................147L.E.D. Annunciator Panel Test Switch, ..................................................148L.E.D. Green EXT/FULL EXT Lights, ....................................................147Landing Altitude Indicator, .....................................................................106Landing Gear Indication System, ..........................................................226Landing Gear Indicator Light, ................................................................225Landing Gear Lever, ..............................................................................227Landing Gear Limitations, .....................................................................263Landing Gear Override Trigger, .............................................................229Landing Gear Solenoid Failure, .............................................................228Landing Gear Towing - Hydraulic Pressurization, .................................263Landing Gear Warning Horn, ................................................................226Landing Gear: Class Notes, ..................................................................365Landing Lights, ......................................................................................134Lavatory Dome Light, ............................................................................141LE FLAPS EXT Light (200 Basic), ........................................................149LE FLAPS EXT Light (200A/300), ........................................................149LE FLAPS TRANSIT Light, ...................................................................148Leading Edge Annunciator Panel (100), ...............................................146Leading Edge Annunciator Panel (200 Basic), .....................................146Leading Edge Annunciator Panel (200A/300/400), ...............................147Leading Edge Devices, .........................................................................150Leading Edge Devices Indicators, .........................................................145Leading Edge Flap Control Valve, .........................................................250Load Accelerations Limitations, ............................................................271Logo Light Switch, .................................................................................140



B737-200

Loss of All AC Power, ............................................................................158Loss of Both Generators, ........................................................................41Low Energy Continuous Ignition, ..........................................................142Low Energy Ignition, ......................................................................143,144Low Oil Pressure Light, .........................................................................218

M

Mach Airspeed Warning Test Switch, ....................................................175Mach Trim, ...............................................................................................22Mach Trim Fail Light, ...............................................................................22Mach Trim System Review, .....................................................................22Mach Trim Test Button, ...........................................................................22Mach Tuck., .............................................................................................22Mach/Airspeed Indicator, ......................................................................194Main Electric Trim Actuator, ..................................................................245Main Gear Downlock Viewer, ................................................................230Main Landing Gear Schematic, .............................................................289Main Outflow Valve, ...............................................................................112Manual Actuation and Reset Handle, ..................................................166Manual Gear Extension Handles, .........................................................229Manual Mode Flight Profile, ..................................................................111Manual Mode Test, ................................................................................115Manual Trim Wheel, ...............................................................................245Manual Trimming, ..................................................................................246Mark II GPWS System, .........................................................................242Marker Beacons, ...................................................................................194MASK/BOOM Switch, ...........................................................................162Master Caution Lights, ..........................................................................179Master Fire Warning Lights, ..................................................................178Master Flight Director Indicator Lights (SP177), ...................................184Max. Motoring, ......................................................................................251Maximum Allowable Fuel Imbalance - Flight Operations, .....................268Maximum Allowable Fuel Imbalance - Landing, ....................................269Maximum APU EGT, .............................................................................275Maximum Cabin Differential Pressure, ..................................................115Maximum Center Tank Fuel Quantity, ...................................................268Maximum CSD Oil Temperature, ..........................................................267Maximum CSD Oil Temperature Rise, ..................................................267Maximum Differential Pressure, .....................................................115,264Maximum EGT Limits, ...........................................................................274



B737-200

Maximum External Air Pressure, ..........................................................271Maximum External Air Temperature, .....................................................271Maximum Flap Extension Altitude, ........................................................266Maximum Fuel Temperature, ................................................................268Maximum Inflight Weight, ......................................................................270Maximum Landing Weight, ....................................................................270Maximum N1 Limits, .............................................................................274Maximum Operating Altitude, ................................................................262Maximum Recommended Wind for Airstair Operations, .......................262Maximum Speed, ..................................................................................262Maximum Speed - Altitude With Window Heat Inoperative, ..................267Maximum Takeoff - Landing Cabin Differential, ....................................264Maximum Takeoff - Landing Tailwind, ...................................................262Maximum Takeoff Altitude, ....................................................................262Maximum Takeoff Weight, .....................................................................270Maximum Taxi Weight, ..........................................................................270Maximum Wing Tank Fuel Quantity, ......................................................268Maximum Zero Fuel Weight, .................................................................270Minimum Engine Starting Pressures, ...................................................272Minimum Flight Crew,............................................................................263Minimum Fuel For Hydraulic System B Ground Operations, ................266Minimum Fuel Temperature - Freeze Point, ..........................................268Minimum N1 RPM During Icing Condition, ...........................................268Minimum Oxygen Dispatch Pressures, .................................................163Minimum Recommended Altitude For Speedbrake Usage, ..................266Mirror Lights, .........................................................................................141Missed Approach Profile, ......................................................................307Missing Airframe & Engine Parts, .........................................................271Mode Selector Switches (SP177), ........................................................187Modulating and Shutoff Valve, ...............................................................130MU Reports, ..........................................................................................225

NN1 RPM Indicator, .................................................................................212N2 RPM Indicator, .................................................................................214Navigation: Class Notes, .......................................................................369Navigational Equipment Limitations, .....................................................269Negative Relief Door, ............................................................................112Nicad Batteries, .......................................................................................70Nickel-cadmium Battery, .........................................................................41



B737-200

No Bleed Takeoffs - C Flow, ..................................................................131No Smoking Passenger Warning Switch, ...............................................71No.7 Burner Can, ..................................................................................142Nose Gear Downlock Viewer, ................................................................231Nose Gear Schematic, ..........................................................................290Nose Wheel Steering, ...........................................................................240

OOFF Flag, ..............................................................................................203OFF Light, .............................................................................................174Off Schedule Descent Light, .................................................................104Oil Distribution System, ........................................................................215Oil Filter Bypass Light, ..........................................................................218Oil Lubrication System, .........................................................................216Oil Pressure Indicator, ..........................................................................215Oil Pressure Limits, ...............................................................................274Oil Quantity Indicator, ...........................................................................216Oil Quantity Sensing Unit, .....................................................................217Oil Tank Capacity, .................................................................................216Oil Temperature, ....................................................................................274Oil Temperature Indicator, .....................................................................216Oil Test Switch, .....................................................................................217One Engine Inop Approach Profile, ...............................................307,309One Engine Inop Missed Approach, ..............................................307,309Operating Differential Pressure, ............................................................264Operational Limitations, ........................................................................262Oral Examinations, ..................................................................................11Outboard Brakes, ..................................................................................236Outboard Landing Lights, ......................................................................134Outflow Valve, .......................................................................................112Outflow Valve Position Indicator, ...........................................................107Outflow Valve Switch, ............................................................................107Overhead System Annunciator, ............................................................181Overheat Detector Switch, ....................................................................253Overheat/Inop and Fire Test Switch, .....................................................254OVHT/DET System Annunciator, ..........................................................181Oxygen Cylinders, .................................................................................169Oxygen Diluter Lever, ............................................................................168Oxygen Flow Indicator, .........................................................................169Oxygen Masks, .....................................................................................167



B737-200

Oxygen Panel, .......................................................................................168Oxygen Shutoff Valve, ...........................................................................167Oxygen Supply Lever, ...........................................................................169Oxygen System Activation, ...................................................................164Oxygen/Boom Switch, ...........................................................................161

PPA Hand Microphone, ...........................................................................252Pack Trip Off Light, ................................................................................126Pack Valves, ..........................................................................................125Parking Brake Lever, .............................................................................247Parking Brake Shutoff Valve, .................................................................248Passenger Address System, .................................................................252Passenger Cabin Emergency Exit Lights, ...............................................70Passenger Cabin Temperature Selector, ...............................................119Passenger Oxygen Indicator, ................................................................164Passenger Oxygen Indicator Light, .......................................................165Passenger Oxygen Switch, ...................................................................165Passenger Service Unit (PSU), .............................................................164PDCS Requirements, ............................................................................265Performance Data Computer System Limitations, ................................265Permanent Magnets, ...............................................................................40Personal Breathing Equipment (PBE), ..................................................170Pitot Heat Requirements, ......................................................................268Pitot Static Heat Switch, .........................................................................79Pneumatic Duct Pressure Indicator, .....................................................124Pneumatics Limitations, ........................................................................271Pneumatics Schematic, ........................................................................293Pneumatics: Class Notes, .....................................................................373Portable Crew Oxygen, .........................................................................170Portable Oxygen System, .....................................................................170Portable Passenger Oxygen, ................................................................169Position Lights, ...............................................................................137,138Power Plant Limitations, ........................................................................272Powerplant: Class Notes, ......................................................................377Pressure Regulator, ..............................................................................130Pressure Relief Valve, ...........................................................................130Pressure Relief Valves, .........................................................................112Pressurization Auto Mode, ....................................................................109Pressurization Limit Placard, ................................................................102



B737-200

Pressurization Limitations, ....................................................................115Pressurization Manual Light, .................................................................105Pressurization Mode Selector, ..............................................................108Pressurization Schematic, ....................................................................295Pressurization Standby Light, ...............................................................104Pressurization System Description, ......................................................111Pressurization System Tests, ................................................................114Probe Heater Lights, ...............................................................................80PTT Switch, ...................................................................................160,162Pulsing Charging, ....................................................................................42

R

Radio Altimeter, .....................................................................................199Radio Altimeter Test Switch, .................................................................199Radio Magnetic Indicator Compass - RMI, ...........................................201Rain Boe - Type III., ................................................................................73Ram Air Schematic, .......................................................................294,295Ram Door Full Open Light, ...................................................................122Receiver Switch, ...................................................................................161Recording Time Remaining Indicator, ...................................................173Red Dome Light Switches, ....................................................................159Refueling, ................................................................................................29Repeat Switch, ......................................................................................174Reset/Manual Activation Handle, ..........................................................164Residual Voltage, ....................................................................................40Residual Volts Button, .............................................................................46Residual Volts Switch, .............................................................................40Respiratory Protection System, ............................................................170Revenue Flights - Retention Bar Use, ...................................................262Reverse Thrust Usage, .........................................................................272Reverser Unlock Light, ..........................................................................217Rudder Pedal Steering, .........................................................................240Rudder Trim Wheel, ...............................................................................251Runway Slope Limits, ............................................................................262Runway Turnoff Lights, ..........................................................................135

SService Interphone Handsets, ..............................................................252Service Interphone Switch, ............................................................157,252Service Interphone System, .................................................................156



B737-200

Simulator Training Hints, .........................................................................17Single Approach Procedures, ........................................................307,309Single System Failures, ........................................................................179Smoke Hoods, .......................................................................................170Speed Brake Armed Light, ....................................................................204Speed Brake Do Not Arm Light, ...........................................................204Speed Brake Lever, ...............................................................................247Speed Brake Test Switches, .................................................................205Speedbrake and Spoiler System,..........................................................205Speedbrake System, .............................................................................204Stabilizer Brake Release Knob, .............................................................251Stabilizer Trim Band Range, ..................................................................245Stabilizer Trim Cutout Switches, ...........................................................246Stabilizer Trim System, .........................................................................245Stabilizer Trim Wheel, ...........................................................................246Stall Warning Off Light, .........................................................................176Stall Warning Switch, ............................................................................176Stall Warning System, ...........................................................................177Stall Warning Test Indicator, ..................................................................177Standby Airspeed Indicator, ..................................................................198Standby Altimeter/Airspeed Indicator, ...................................................198Standby Compass Light, ................................................................141,158Standby Hydraulic Components, .............................................................94Standby Hydraulic Low Quantity Amber Light, .......................................17Standby Hydraulic System, .....................................................................17Standby Hydraulic System Schematic, .................................................287Standby Mode Flight Profile, .................................................................110Standby Mode Test, ..............................................................................114Standby Power Switch, ...........................................................................46Standby Pressurization Mode Failure, ..................................................104Start Levers, ..........................................................................................250Start Valve Open Light, .........................................................................217Starter Duty Cycle, ................................................................................273Synchronization Indicator, .......................................................................15System A Hydraulic Components, ..........................................................92System Annunciator Panel, ...................................................................180System B Hydraulic Components, ..........................................................93System B Hydraulics Schematic, ..........................................................286



B737-200

TTakeoff Configuration Warning Horn, ....................................................227Takeoff Profile - Normal Takeoff, ...........................................................301Takeoff Profile - Normal Takeoff Callouts, .............................................302Takeoff Rejected, ..........................301,302,303,304,305,306,307,309,310Takeoff Warning Horn, ..........................................................................245Takeoff/Go-Around Switches, ...............................................................248TAT-MAX EPR Indicator, .......................................................................234Taxi Light, ..............................................................................................136TCAS Compliance Requirements, ........................................................272TCAS Limitations, .................................................................................271TCAS Response Maneuvers, ................................................................272TCAS Weather Approval, ......................................................................271Thermal Discharge - Fire Extinguisher Bottles, ....................................260Thrust Levers, .......................................................................................249Thrust Reverser Isolation Valve Panel, .................................................152Thrust Reverser Schematic, .................................................................297Thrust Reverser Shutoff Valve, .............................................................255Thumbwheel Switches, .........................................................................174TO/GA Button, ......................................................................................248Total Air Temperature Indicator, ............................................................232Total Air Temperature Indicator, ............................................................233TR Voltage Range, ................................................................................267Trailing Edge Control Valve, ..................................................................250Trailing Edge Flap Bypass Valve, ............................................................18Transfer Bus Off Light, ............................................................................52Transfer Relays, ......................................................................................37Transformer Rectifier, ..............................................................................42Transformer Rectifier Failure, ..................................................................43Trip and Date Light, ...............................................................................174Trip and Date Selectors, ........................................................................173Trip Reset Switch, ..........................................................................126,131Turbine Discharge Pressure (Pt7), ........................................................211Turbofan System, ..................................................................................123Turbulent Airspeeds, .............................................................................263Type of Airplane Operation, ..................................................................263



B737-200

UUnsafe Terrain Clearance - Not in Landing, ..........................................243

VVerification of EPR Values, ...................................................................266Vertical Gyro Transfer Switch, ..........................................................23,196Vertical Speed Display (SP177), ...........................................................188Vertical Speed Indicator, .......................................................................203Vertical Speed Thumbwheel, ................................................................188

WWarning Systems: Class Notes, ...........................................................381Weather Radar Requirements, .............................................................269Weather Radar Warm-up Requirements, ..............................................269Wheel Well Fire Warning Light, .............................................................254Wheel Well Light Switch, .......................................................................231Wheel Well Lights, .................................................................................140White Dome Light, ................................................................................141White Dome Light Switch, .....................................................................158White/Red Dome Light Switches, .........................................................159Window Heat ON Light, ..........................................................................76Window Heat Switch, ..............................................................................77Window Heat Test Switch, ......................................................................77Window Overheat, ...................................................................................78Window Overheat Light, ..........................................................................77Windows, .................................................................................................79Windshield Wiper Selector, .....................................................................74Wing Anti-ice Schematic Decal, ............................................................127Wing Anti-Ice Switch, .........................................................................81,82Wing Anti-ice Valve, ................................................................................84Wing Anti-Ice Valve Open Light, .............................................................81Wing Illumination Switch, ......................................................................139Wing Thermal Anti-ice System Requirements, .....................................267Wing-Body Overheat Light, ...................................................................127Wing-Body Overheat Test Switch, ........................................................120



B737-200

Y

Yaw Control, ..........................................................................................219Yaw Damper Indicator, ..........................................................................219Yaw Damper System, ..............................................................................21Yaw Damper Warning Light and Switch, .................................................21

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