QDB Flashcards

Question

40. When will the Cockpit Voice Recorder automatically operate? a) Anytime the battery switch is ON b) Anytime DC power is available c) In flight only d) The area microphone is activated anytime 115V AC is applied to airplane.

Answer 1

The area microphone is activated anytime 115V AC is applied to airplane. Area Microphone - OA-M Part B vol. 2 – 5.10.11 Active anytime 115V AC is applied to airplane.

Answer 2

The VHF-2 and Cabin/Service interphone receiver switches may be selected at the same time.

Answer 3

Transfer bus 2 Transformer Rectifier Units - OA-M Part B vol. 2 – 6.20.8 The TRs convert 115 volt AC to 28 volt DC, and are identified as TR1, TR2, and TR3. TR1 receives AC power from transfer bus 1. TR2 receives AC power from transfer bus 2. TR3 normally receives AC power from transfer bus 2 and has a backup source of AC power from transfer bus 1. Any two TRs are capable of supplying the total connected load. Under normal conditions, DC bus 1, DC bus 2, and the DC standby bus are connected via the cross bus tie relay. In this condition, TR1 and TR2 are each powering DC bus 1, DC bus 2, and the DC standby bus. TR3 powers the battery bus and serves as a backup power source for TR1 and TR2. The cross bus tie relay automatically opens, isolating DC bus 1 from DC bus 2, under the following conditions: • At glide slope capture during a flight director or autopilot ILS approach. This isolates the DC busses during approach to prevent a single failure from affecting both navigation receivers and flight control computers • Bus transfer switch positioned to OFF. In–flight, an amber TR UNIT light illuminates if TR1, or TR2 and TR3 has failed. On the ground, any TR fault causes the light to illuminate.

Answer 4

AC ground service bus 2 Battery Charger Transformer/Rectifier - OA-M Part B vol. 2 – 6.20.9 Single Battery The purpose of the battery charger is to restore and maintain the battery at full electrical power. The battery charger is powered through AC ground service bus 2. The battery charger provides a voltage output tailored to maximize the battery charge. Following completion of the primary charge cycle, the battery charger reverts to a constant voltage TR mode. In the TR mode, it powers loads connected to the hot battery bus and the switched hot battery bus. The battery charger TR also powers the battery bus if TR3 fails. With loss of AC transfer bus 1 or the source of power to DC bus 1, the AC and DC standby busses are powered by the battery/battery charger.

Answer 5

Galley on transfer bus 2 are shed first. Automatic Load Shedding (Engine Generators) - OA-M Part B vol. 2 – 6.20.4 For single generator operation, the system is designed to shed electrical load incrementally based on actual load sensing. The galleys and main bus on transfer bus 2 are shed first; if an overload is still sensed, the galleys and main bus on transfer bus 1 are shed; if overload still exists, the IFE buses are shed. When configuration changes to more source capacity (two generator operation), automatic load restoration of the main busses, galley busses and IFE buses occurs; manual restoration of galley and main bus power can be attempted by moving the CAB/UTIL Power Switch to OFF, then back ON. APU Automatic Load Shedding - OA-M Part B vol. 2 – 6.20.4 In flight, if the APU is the only source of electrical power, all galley busses and main buses are automatically shed. If electrical load still exceeds design limits, both IFE busses are also automatically shed. On the ground, the APU attempts tocarry a full electrical load. If an overload condition is sensed, the APU sheds galley busses and main busses until the load is within limits. Manual restoration of galley and main bus power can be attempted by moving the CAB/UTIL Power Switch to OFF, then back ON.

Answer 6

The BUS TRANS switch must be in the AUTO position. BUS TRANSFER Switch - OA-M Part B vol. 2 – 6.10.7 AUTO (guarded position) – BTBs operate automatically to maintain power to AC transfer busses from any operating generator or external power • DC cross tie relay automatically provides normal or isolated operation as required. OFF – isolates AC transfer bus 1 from AC transfer bus 2 if one IDG is supplying power to both AC transfer busses • DC cross tie relay opens to isolate DC bus 1 from DC bus 2.

Answer 7

Only the associated GEN OFF BUS light and SOURCE OFF light illuminate

Answer 8

TR 1or TR2 and TR3 has failed Transformer Rectifier Units - OA-M Part B vol. 2 – 6. The TRs convert 115 volt AC to 28 volt DC, and are identified as TR1, TR2, and TR3. TR1 receives AC power from transfer bus 1. TR2 receives AC power from transfer bus 2. TR3 normally receives AC power from transfer bus 2 and has a backup source of AC power from transfer bus 1. Any two TRs are capable of supplying the total connected load. Under normal conditions, DC bus 1, DC bus 2, and the DC standby bus are connected via the cross bus tie relay. In this condition, TR1 and TR2 are each powering DC bus 1, DC bus 2, and the DC standby bus. TR3 powers the battery bus and serves as a backup power source for TR1 and TR2. The cross bus tie relay automatically opens, isolating DC bus 1 from DC bus 2, under the following conditions: • At glide slope capture during a flight director or autopilot ILS approach. This isolates the DC busses during approach to prevent a single failure from affecting both navigation receivers and flight control computers • Bus transfer switch positioned to OFF. In–flight, an amber TR UNIT light illuminates if TR1, or TR2 and TR3 has failed. On the ground, any TR fault causes the light to illuminate

Answer 9

Isolate DC bus 1 from DC bus 2 OA-M Part B vol. 2 – 6.20.7

Answer 10

AC transfer bus 1 or DC bus 1 loses power Standby Power System - OA-M Part B vol. 2 – 6.20.11 Normal Operation The standby system provides 115V AC and 24V DC power to essential systems in the event of loss of all engine or APU– driven AC power. The standby power system consists of: • static inverter • AC standby bus • DC standby bus • battery bus • hot battery bus • switched hot battery bus • main battery During normal operation the guarded standby power switch is in AUTO and the battery switch is ON. This configuration provides alternate power sources in case of partial power loss as well as complete transfer to battery power if all normal power is lost. Under normal conditions the AC standby bus is powered from AC transfer bus 1. The DC standby bus is powered by TR1, TR2, and TR3; the battery bus is powered by TR3; the hot battery bus and switched hot battery bus are powered by the battery/battery charger. Alternate Operation Single Battery The alternate power source for standby power is the battery. With the standby power switch in the AUTO position, the loss of all engine or APU electrical power causes the battery to power the standby loads, both in the air and on the ground. The AC standby bus is powered from the battery via the static inverter. The DC standby bus, battery bus, hot battery bus, and switched hot battery bus are powered directly from the battery. The standby power switch provides for automatic or manual control of power to the standby buses. In the AUTO position, automatic switching from normal to alternate power occurs if power from either AC transfer bus 1 or DC bus 1 is lost. Positioning the switch to BAT overrides automatic switching and places the AC standby bus, DC standby bus, and battery bus on battery power. The battery switch may be ON or OFF. If the battery switch is OFF, the switched hot battery bus is not powered. Positioning the standby power switch to OFF de–energizes both the AC standby bus and the DC standby bus and illuminates the STANDBY PWR OFF light.

Answer 11

Ground Power has been plugged in and meets airplane power quality standards Ground Power Available (GRD POWER AVAILABLE) Light - OA-M Part B vol. 2 – 6.10.7 Illuminated (blue) – ground power is connected and meets airplane power quality standards.

Answer 12

Both IDG’s will come on line automatically if the APU fails or is shut down AC Power System - OA-M Part B vol. 2 – 6. Each AC power system consists of a transfer bus, a main bus, two galley busses, and a ground service bus. Transfer bus 1 also supplies power to the AC standby bus. If the AC source powering either transfer bus fails or is disconnected, the transfer bus can be powered by any available source through the tie bus with the bus tie breakers (BTBs). With the airplane on the ground and both generator control switches OFF, or with both engines shut down, selecting the GRD PWR switch ON connects external power to both transfer busses. Likewise, selecting either APU GEN switch ON connects APU power to both transfer busses. Whichever source is selected last powers both busses. It is not possible to power one transfer bus with external power and one transfer bus with APU power. The transfer busses can be powered from the engine generators by momentarily positioning the related generator switch to ON. This closes the related generator circuit breaker (GCB) and connects the generator to the transfer bus. Whenever external power or APU is powering both transfer busses, and engine generator power is applied to its onside transfer bus, external power or APU continues to supply power to the remaining transfer bus. In flight, each engine generator normally powers its own transfer bus. If an engine generator is no longer supplying power, the BTBs automatically close to allow the other engine generator to supply both transfer busses through the tie bus and BTBs. The APU can power either or both busses through the BTBs. The system also incorporates an automatic generator on–line feature in case the airplane takes off with the APU powering both transfer busses. If the APU is either shut down or fails, the engine generators are automatically connected to their related transfer busses. This action occurs only once in flight and only under the circumstances described above.

Answer 13

Low oil pressure in the IDG Generator Drive (DRIVE) Lights - OA-M Part B vol. 2 – 6.10.4 Illuminated (amber) – Integrated drive generator (IDG) low oil pressure caused by one of the following: • IDG failure • engine shutdown • IDG automatic disconnect due to high oil temperature • IDG disconnected through generator drive DISCONNECT switch.

Answer 14

At 56% N2 Engine Start System - OA-M Part B vol. 2 – 7.20.8 Starter operation requires pressurized air and electrical power. Air from the bleed air system powers the starter motor. The APU, an external ground cart, or the other operating engine provides the bleed air source. In the GRD position, the engine start switch uses battery power to close the engine bleed air valve and open the start valve to allow pressure to rotate the starter. When the start valve opens, an amber START VALVE OPEN alert is provided on the upper display unit. The starter rotates the N2 compressor through the accessory drive gear system. When the engine accelerates to the recommended value (25% N2 or max motoring), moving the engine start lever to the IDLE position opens the fuel valves on the wing spar and engine, and causes the EEC to supply fuel and ignition to the combustor where the fuel ignites. Initial fuel flow indications lag actual fuel flow by approximately two seconds, therefore, during engine start, an EGT rise may occur before fuel flow indication. At starter cutout speed (approximately 56% N2), power is removed from the start switch holding solenoid. The engine start switch returns to OFF, the engine bleed air valve returns to the selected position, and the start valve closes.

Answer 15

Either radio altimeter sense below 10 feet Thrust Reverser - OA-M Part B vol. 2 – 7.20.11 Each engine is equipped with a hydraulically operated thrust reverser, consisting of left and right translating sleeves. Aft movement of the reverser sleeves causes blocker doors to deflect fan discharge air forward, through fixed cascade vanes, producing reverse thrust. The thrust reverser is for ground operations only and is used after touchdown to slow the airplane, reducing stopping distance and brake wear. Hydraulic pressure for the operation of engine No. 1 and engine No. 2 thrust reversers comes from hydraulic systems A and B, respectively. If hydraulic system A and/or B fails, alternate operation for the affected thrust reverser is available through the standby hydraulic system. When the standby system is used, the affected thrust reverser deploys and retracts at a slower rate and some thrust asymmetry can be anticipated. The thrust reverser can be deployed when either radio altimeter senses less than 10 feet altitude, or when the air/ground safety sensor is in the ground mode. Movement of the reverse thrust levers is mechanically restricted until the forward thrust levers are in the idle position. When reverse thrust is selected, an electro–mechanical lock releases, the isolation valve opens and the thrust reverser control valve moves to the deploy position, allowing hydraulic pressure to unlock and deploy the reverser system. An interlock mechanism restricts movement of the reverse thrust lever until the reverser sleeves have approached the deployed position. When either reverser sleeve moves from the stowed position, the amber REV indication, located on the upper display unit, illuminates. As the thrust reverser reaches the deployed position, the REV indication illuminates green and the reverse thrust lever can be raised to detent No. 2. This position provides adequate reverse thrust for normal operations. When necessary, the reverse thrust lever can be pulled beyond detent No. 2, providing maximum reverse thrust. Downward motion of the reverse thrust lever past detent No. 1 (reverse idle thrust) initiates the command to stow the reverser. When the lever reaches the full down position, the control valve moves to the stow position allowing hydraulic pressure to stow and lock the reverser sleeves. After the thrust reverser is stowed, the isolation valve closes and the electro–mechanical lock engages. The REVERSER light, located on the aft overhead panel, illuminates when the thrust reverser is commanded to stow and extinguishes 10 seconds later when the isolation valve closes. Any time the REVERSER light illuminates for more than approximately 12 seconds, a malfunction has occurred and the MASTER CAUTION and ENG system annunciator lights illuminate. Note: A pause in movement of the reverse thrust levers past detent No. 1 toward the stow position may cause MASTER CAUTION and ENG system annunciator lights to illuminate. A pause of approximately 18 seconds engages the electro-mechanical lock and prevents the thrust reverser sleeves from further movement. Cycling the thrust reversers may clear the fault and restore normal operation. When the reverser sleeves are in the stow position, an electro–mechanical lock and a hydraulically operated locking actuator inhibit motion to each reverser sleeve until reverser extension is selected. Additionally, an auto–restow circuit compares the actual reverser sleeve position and the commanded reverser position. In the event of incomplete stowage or uncommanded movement of the reverser sleeves toward the deployed position, the auto–restow circuit opens the isolation valve and commands the control valve to the stow position directing hydraulic pressure to stow the reverser sleeves. Once the auto–restow circuit is activated, the isolation valve remains open and the control valve is held in the stowed position until the thrust reverser is commanded to deploy or until corrective maintenance action is taken. WARNING: Actuation of the thrust reversers on the ground without suitable precautions is dangerous to ground personnel.

Answer 16

15 seconds after the start lever is moved to idle during ground starts Abnormal Start Protection (Ground Starts Only) - OA-M Part B vol. 2 – 7.20.8 During ground starts, the EEC monitors engine parameters to detect impending hot starts, engine stalls, EGT start limit exceedances, and wet starts. These protection features do not function during inflight starts. If an impending hot start is detected by a rapid rise in EGT or EGT approaching the start limit, or a compressor stall occurs, the white box surrounding the EGT digital readout flashes white. The flashing white box resets when the start lever is moved to CUTOFF or the engine reaches idle N2. Current versions of EEC software (7.B.Q and later) automatically turn off the ignition and shuts off fuel to the engine for an impending hot start or stall. If the EGT exceeds the starting limit, the EGT display both box and dial, turn red. The EEC automatically turns off the ignition and shuts off fuel to the engine. The alert terminates and the display returns to white when EGT drops below the start limit. Following shutdown of both engines, the EGT box turns red to remind the crew of the exceedance. A wet start occurs if the EGT does not rise after the start lever is moved to IDLE. If a wet start is detected, the EEC turns off the ignition and shuts off fuel to the engine 15 seconds after the start lever is moved to IDLE.

Answer 17

Above the N2 dial on the CDS Inflight Starting - OA-M Part B vol. 2 – 7.20.9 Two methods of starting an engine inflight are available, windmill and crossbleed. None of the ground start protection features are functional during inflight start. Note: At low N2 values, the oil scavenge pump may not provide enough pressure to return oil to the tank, causing a low oil quantity indication. Normal oil quantity should be indicated after start. If crossbleed starting is required, the X–BLD indication (XB for the compact engine display) is displayed above the N2 dial. This indication is based on airplane altitude, airspeed and N2.

Answer 18

Ground starts only Abnormal Start Protection (Ground Starts Only) - OA-M Part B vol. 2 – 7.20.8 During ground starts, the EEC monitors engine parameters to detect impending hot starts, engine stalls, EGT start limit exceedances, and wet starts. These protection features do not function during inflight starts. If an impending hot start is detected by a rapid rise in EGT or EGT approaching the start limit, or a compressor stall occurs, the white box surrounding the EGT digital readout flashes white. The flashing white box resets when the start lever is moved to CUTOFF or the engine reaches idle N2. Current versions of EEC software (7.B.Q and later) automatically turn off the ignition and shuts off fuel to the engine for an impending hot start or stall. If the EGT exceeds the starting limit, the EGT display both box and dial, turn red. The EEC automatically turns off the ignition and shuts off fuel to the engine. The alert terminates and the display returns to white when EGT drops below the start limit. Following shutdown of both engines, the EGT box turns red to remind the crew of the exceedance. A wet start occurs if the EGT does not rise after the start lever is moved to IDLE. If a wet start is detected, the EEC turns off the ignition and shuts off fuel to the engine 15 seconds after the start lever is moved to IDLE.

Answer 19

N1 Electronic Engine Control (EEC) - OA-M Part B vol. 2 – 7.20.3 Each engine has a full authority digital EEC. Each EEC has two independent control channels, with automatic channel transfer if the operating channel fails. With each engine start or start attempt, the EEC alternates between control channels. The EEC uses thrust lever inputs to automatically control forward and reverse thrust. N1 is used by the EEC to set thrust in two control modes: normal and alternate. Manual selection of the control mode can be made with the EEC switches on engine panel. EEC Normal Mode In the normal mode, the EEC uses sensed flight conditions and bleed air demand to calculate N1 thrust ratings. The EEC compares commanded N1 to actual N1 and adjusts fuel flow to achieve the commanded N1. The full rated takeoff thrust for the installed engine is available at a thrust lever position less than the forward stop. Fixed or assumed temperature derated takeoff thrust ratings are set at thrust lever positions less than full rated takeoff. The maximum rated thrust is available at the forward stop. The EEC limits the maximum thrust according to the airplane model as follows: • 737-800 – CFM56-7B27 rating EEC Alternate Mode The EEC can operate in either of two alternate modes, soft or hard. If required signals are not available to operate in the normal mode, the EEC automatically changes to the soft alternate mode. When this occurs, the ALTN switch illuminates and the ON indication remains visible. In the soft alternate mode, the EEC uses the last valid flight conditions to define engine parameters which allows the mode change to occur with no immediate change in engine thrust. Thrust rating shortfalls or exceedances may occur as flight conditions change. The soft alternate mode remains until the hard alternate mode is entered by either retarding the thrust lever to idle or manually selecting ALTN with the EEC switch on the aft overhead panel. Note: Loss of either DEU results in a loss of signal to both EECs. The EEC ALTN lights illuminate and each EEC reverts to the alternate mode to prevent the engines from operating on a single source of data. When the hard alternate mode is entered, the EEC reverts to the alternate mode thrust schedule. Hard alternate mode thrust is always equal to or greater than normal mode thrust for the same lever position. If the hard alternate mode is entered by reducing the thrust lever to idle while in the soft alternate mode, the ALTN switch remains illuminated and the ON indication remains visible. When ALTN is selected manually, the ON indication is blanked.

Answer 20

Hydraulic system A for the no. 1 thrust reverser, hydraulic system B for the no. 2 thrust reverser Thrust Reverser - OA-M Part B vol. 2 – 7.20.11 Each engine is equipped with a hydraulically operated thrust reverser, consisting of left and right translating sleeves. Aft movement of the reverser sleeves causes blocker doors to deflect fan discharge air forward, through fixed cascade vanes, producing reverse thrust. The thrust reverser is for ground operations only and is used after touchdown to slow the airplane, reducing stopping distance and brake wear. Hydraulic pressure for the operation of engine No. 1 and engine No. 2 thrust reversers comes from hydraulic systems A and B, respectively. If hydraulic system A and/or B fails, alternate operation for the affected thrust reverser is available through the standby hydraulic system. When the standby system is used, the affected thrust reverser deploys and retracts at a slower rate and some thrust asymmetry can be anticipated. The thrust reverser can be deployed when either radio altimeter senses less than 10 feet altitude, or when the air/ground safety sensor is in the ground mode. Movement of the reverse thrust levers is mechanically restricted until the forward thrust levers are in the idle position. When reverse thrust is selected, an electro–mechanical lock releases, the isolation valve opens and the thrust reverser control valve moves to the deploy position, allowing hydraulic pressure to unlock and deploy the reverser system. An interlock mechanism restricts movement of the reverse thrust lever until the reverser sleeves have approached the deployed position. When either reverser sleeve moves from the stowed position, the amber REV indication, located on the upper display unit, illuminates. As the thrust reverser reaches the deployed position, the REV indication illuminates green and the reverse thrust lever can be raised to detent No. 2. This position provides adequate reverse thrust for normal operations. When necessary, the reverse thrust lever can be pulled beyond detent No. 2, providing maximum reverse thrust. Downward motion of the reverse thrust lever past detent No. 1 (reverse idle thrust) initiates the command to stow the reverser. When the lever reaches the full down position, the control valve moves to the stow position allowing hydraulic pressure to stow and lock the reverser sleeves. After the thrust reverser is stowed, the isolation valve closes and the electro–mechanical lock engages. The REVERSER light, located on the aft overhead panel, illuminates when the thrust reverser is commanded to stow and extinguishes 10 seconds later when the isolation valve closes. Any time the REVERSER light illuminates for more than approximately 12 seconds, a malfunction has occurred and the MASTER CAUTION and ENG system annunciator lights illuminate. Note: A pause in movement of the reverse thrust levers past detent No. 1 toward the stow position may cause MASTER CAUTION and ENG system annunciator lights to illuminate. A pause of approximately 18 seconds engages the electro-mechanical lock and prevents the thrust reverser sleeves from further movement. Cycling the thrust reversers may clear the fault and restore normal operation. When the reverser sleeves are in the stow position, an electro–mechanical lock and a hydraulically operated locking actuator inhibit motion to each reverser sleeve until reverser extension is selected. Additionally, an auto–restow circuit compares the actual reverser sleeve position and the commanded reverser position. In the event of incomplete stowage or uncommanded movement of the reverser sleeves toward the deployed position, the auto–restow circuit opens the isolation valve and commands the control valve to the stow position directing hydraulic pressure to stow the reverser sleeves. Once the auto–restow circuit is activated, the isolation valve remains open and the control valve is held in the stowed position until the thrust reverser is commanded to deploy or until corrective maintenance action is taken. WARNING: Actuation of the thrust reversers on the ground without suitable precautions is dangerous to ground personnel.

Answer 21

AC standby bus Engine Ignition System - OA-M Part B vol. 2 – 7.20.9 Each engine has two igniter plugs. The EEC arms the igniter plug(s) selected by the ignition select switch. The left igniter plug receives power from the associated AC transfer bus. The right igniter plug receives power from the AC standby bus. Auto-Relight An auto-relight capability is provided for flameout protection. Whenever the EEC detects an engine flameout, both igniters are activated. A flameout is detected when an uncommanded rapid decrease in N2 occurs, or N2 is below idle RPM.

Answer 22

Respective Start Lever is placed to CUTOFF or Respective Engine Fire Handle is pulling. Engine Fuel System - OA-M Part B vol. 2 – 7.20.6 Fuel is delivered under pressure from fuel pumps located in the fuel tanks. The fuel flows through a fuel spar shutoff valve located at the engine mounting wing stations. The fuel passes through the first stage engine fuel pump where pressure is increased. It then passes through two fuel/oil heat exchangers where IDG oil and main engine oil heat the fuel. A fuel filter then removes contaminants. Fuel automatically bypasses the filter if the filter becomes saturated. Before the fuel bypass occurs, the fuel FILTER BYPASS alert illuminates on the fuel control panel. The second stage engine fuel pump adds more pressure before the fuel reaches the hydro mechanical unit (HMU). To meet thrust requirements, the EEC meters fuel through the HMU. The spar fuel shutoff valve and engine fuel shutoff valve allow fuel flow to the engine when both valves are open. The valves are open when the engine fire warning switch is in and the start lever is in IDLE. Both valves close when either the start lever is in CUTOFF or the engine fire warning switch is out. SPAR VALVE CLOSED and ENG VALVE CLOSED lights located on the overhead panel indicate valve position. Fuel flow is measured after passing through the engine fuel shutoff valve and is displayed on the display unit. Fuel flow information is also provided to the FMS.

Answer 23

The EEC provides flameout protection and red line protection for N1 and N2. The EEC does not provide EGT redline exceedance protection. Structural Limit Protection - OA-M Part B vol. 2 – 7. The EEC provides N1 and N2 redline overspeed protection in both normal and alternate modes. The EGT limit must be observed by the crew because the EEC does not provide EGT redline exceedance protection.

Answer 24

The APU will operate in the no load mode for 1 minute, then shutdown APU Shutdown - OA-M Part B vol. 2 – 7.30.3 Operate the APU for one full minute with no bleed air load prior to shutdown. This cooling period is recommended to extend the service life of the APU. When the APU switch is moved to OFF, this time delay is met automatically. Moving the APU switch to OFF trips the APU generator, closes the APU bleed air valve and extinguishes the APU GEN OFF BUS light. Shutdown occurs automatically after 60 seconds. When the APU speed decreases sufficiently during shutdown, the fuel valve and inlet door close. If the fuel valve does not close, the FAULT light will illuminate after approximately 30 seconds. An immediate shutdown can be accomplished by pulling the APU fire switch.

Answer 25

The battery or AC power from transfer bus 1 Electrical Requirements for APU Operation - OA-M Part B vol. 2 – 7.30.3 APU operation requires the following: • APU fire switch on the overheat/fire panel must be IN • APU fire control handle on the APU ground control panel must be IN • battery switch must be ON. Electrical power to start the APU comes from No. 1 transfer bus or the airplane battery. With AC power available, the starter generator uses AC power to start the APU. With no AC power, the starter generator uses battery power to start the APU. Moving the battery switch to OFF on the ground or in the air automatically shuts down the APU because of power loss to the electronic control unit.

Answer 26

The aircraft is in flight, APU bleed air switch is on and either or both pack switches are in High position.

Answer 27

60 seconds APU Shutdown - OA-M Part B vol. 2 – 7.30.3 Operate the APU for one full minute with no bleed air load prior to shutdown. This cooling period is recommended to extend the service life of the APU. When the APU switch is moved to OFF, this time delay is met automatically. Moving the APU switch to OFF trips the APU generator, closes the APU bleed air valve and extinguishes the APU GEN OFF BUS light. Shutdown occurs automatically after 60 seconds. When the APU speed decreases sufficiently during shutdown, the fuel valve and inlet door close. If the fuel valve does not close, the FAULT light will illuminate after approximately 30 seconds. An immediate shutdown can be accomplished by pulling the APU fire switch.

Answer 28

It suction feeds APU Fuel Supply - OA-M Part B vol. 2 – 7.30.1 Fuel to start and operate the APU comes from the left side of the fuel manifold when the AC fuel pumps are operating. If the AC fuel pumps are not operating, fuel is suction fed from the No. 1 tank. During APU operation, fuel is automatically heated to prevent icing.

Answer 29

On the ground or inflight

Answer 30

Close the fuel APU shutoff valve, APU bleed Air Valve, and APU inlet door APU Fire Extinguishing - OA-M Part B vol. 2 – 8.20.1 The APU fire extinguisher system consists of one APU fire extinguisher bottle, an APU fire switch, an APU BOTTLE DISCHARGE light, and an EXT TEST switch. The APU ground control panel located in the right main wheel well also contains an APU fire warning light, an APU BOTTLE DISCHARGE switch, an APU fire control handle and APU HORN CUTOUT switch. The APU fire switch is normally locked down to prevent inadvertent shutdown of the APU. Illumination of the APU fire switch unlocks the switch. The switch may also be unlocked manually. Pulling the APU Fire switch up: • provides backup for the automatic shutdown feature • deactivates the fuel solenoid and closes the APU fuel shutoff valve • closes the APU bleed air valve • closes the APU air inlet door • trips the APU generator control relay and breaker • allows the APU fire switch to be rotated for discharge • arms the APU fire extinguisher bottle squib.

Answer 31

Engine overheat detection powered by the battery bus, engine fire detection powered by the battery bus and engine fire extinguishing powered by the hot battery bus. Engine Fire Protection - OA-M Part B vol. 2 – 8.20.1 Engine fire protection consists of these systems: • engine overheat and fire detection powered by the battery bus • engine fire extinguishing powered by the hot battery bus.

Answer 32

An illuminated FAULT light Engine Overheat and Fire Detection - OA-M Part B vol. 2 – 8.20.1 Each engine contains two overheat/fire detector loops. Each loop provides both fire and overheat detection. As the temperature of a detector increases to a predetermined limit, the detector senses an overheat condition. At higher temperatures, the detector senses a fire condition. Normally, both detector loops must sense a fire or overheat condition to cause an engine overheat or fire alert. The ENG OVERHEAT light or engine fire switch remains illuminated until the temperature drops below the onset temperature. An OVHT DET switch for each engine, labeled A, B, and NORMAL, permits selection of either loop A or B, or both A and B, as the active detecting loops. The system contains a fault monitoring circuit. If one loop fails with the OVHT DET switch in NORMAL, that loop is automatically deselected and the remaining loop functions as a single loop detector. There is no flight deck indication of single loop failure. If both loops fail on an engine, the FAULT light illuminates and the system is inoperative. If the OVHT DET switch is positioned to A or B, the system operates as a single loop system. The non–selected loop is not monitored. If the selected loop fails, the FAULT light illuminates and the system is inoperative. The indications of an engine overheat are: • both MASTER CAUTION lights illuminate • the OVHT/DET system annunciator light illuminates • the related ENG OVERHEAT light illuminates. The indications of an engine fire are: • the fire warning bell sounds • both master FIRE WARN lights illuminate • the related engine fire switch illuminates • all related engine overheat alert indications illuminate

Answer 33

One or more detectors in the loops have failed. Cargo Fire TEST - OA-M Part B vol. 2 – 8.20.9 The indications for the Cargo Fire test are: • the fire warning bell sounds • both master FIRE WARN lights illuminate • the extinguisher test lights illuminate • the FWD and AFT cargo fire warning lights illuminate when all detectors in selected loops (s) respond to the fire test • the cargo fire bottle DISCH light illuminates Note: The fire warning BELL CUTOUT switch on the Overheat/Fire Protection panel can silence the fire warning bell and extinguish the master FIRE WARN lights. Note: During a Cargo Fire Test, the DETECTOR Fault light will illuminate if one or more detectors in the loop(s) has failed. Note: Individual detector faults can only be detected by a manually initiated test. The MASTER CAUTION light does not illuminate. Note: At the end of cargo fire testing, up to a four second delay may occur to allow all applicable indications to extinguish at the same time.

Answer 34

Close the fuel, hydraulic and bleed air valves Engine Fire Extinguishing - OA-M Part B vol. 2 – 8.20.8 The engine fire extinguisher system consists of two engine fire extinguisher bottles, two engine fire switches, two BOTTLE DISCHARGE lights, and an EXT TEST switch. Either or both bottles can be discharged into either engine. The engine fire switches are normally locked down to prevent inadvertent shutdown of an engine. Illumination of an engine fire switch or ENG OVERHEAT light unlocks the engine fire switch. The switches may also be unlocked manually. Pulling the engine fire switch up: • closes both the engine fuel shutoff valve and the spar fuel shutoff valve • closes the engine bleed air valve resulting in loss of wing anti–ice to the affected wing and closure of bleed air operated pack valve • trips the generator control relay and breaker • closes the hydraulic fluid shutoff valve. The engine driven hydraulic pump LOW PRESSURE light is deactivated • disables thrust reverser for the related engine. • allows the engine fire switch to be rotated for discharge • arms one discharge squib on each engine fire extinguisher bottle. Rotating the engine fire switch electrically “fires” a squib, discharging the extinguishing agent into the related engine. Rotating the switch the other way discharges the remaining bottle. The L or R BOTTLE DISCHARGE light illuminates a few seconds after the engine fire switch is rotated, indicating the bottle has discharged.

Answer 35

Engine Engine Overheat and Fire Detection Each engine contains two overheat/fire detector loops. Each loop provides both fire and overheat detection. As the temperature of a detector increases to a predetermined limit, the detector senses an overheat condition. At higher temperatures, the detector senses a fire condition. Normally, both detector loops must sense a fire or overheat condition to cause an engine overheat or fire alert. The ENG OVERHEAT light or engine fire switch remains illuminated until the temperature drops below the onset temperature. APU Fire Detection A single fire detection loop is installed on the APU. As the temperature of the detector increases to a predetermined limit, the detector senses a fire condition. The APU fire switch remains illuminated until the temperature of the detector has decreased below the onset temperature. Main Wheel Well Fire Detection A single fire detector loop is installed in the main wheel well. As the temperature of the detector increases to a predetermined limit, the detector senses a fire condition. The WHEEL WELL fire warning light remains illuminated until the temperature of the detector has decreased below the onset temperature.

Answer 36

Engines, APU, cargo compartment, and toilets There are fire detection and extinguishing systems for: • engines • APU • lavatories • cargo compartments The engines also have overheat detection systems. The main gear wheel well has a fire detection system, but no fire extinguishing system.

Answer 37

Engine no.1, and engine no.2 Fault Light Illuminated (amber) – with the overheat detector switch in NORMAL – indicates both detector loops for an engine have failed. Illuminated (amber) – with the overheat detector switch in A or B – indicates the selected loop for an engine has failed. Note: MASTER CAUTION and OVHT/DET system annunciator lights do not illuminate.

Answer 38

Arms the alternate flaps position switch Alternate Extension In the event that hydraulic system B fails, an alternate method of extending the LE devices and extending and retracting the TE flaps is provided. The TE flaps can be operated electrically through the use of two alternate flap switches. The guarded ALTERNATE FLAPS master switch closes a flap bypass valve to prevent hydraulic lock of the flap drive unit and arms the alternate flaps position switch. The ALTERNATE FLAPS position switch controls an electric motor that extends or retracts the TE flaps. The switch must be held in the DOWN position until the flaps reach the desired position. No asymmetry or skew protection is provided through the alternate (electrical) flap drive system. When using alternate flap extension the LE flaps and slats are driven to the full extended position using power from the standby hydraulic system. In this case the ALTERNATE FLAPS master switch energizes the standby pump and the ALTERNATE FLAPS position switch, held in the down position momentarily, fully extends the LE devices. Note: The LE devices cannot be retracted by the standby hydraulic system.

Answer 39

There are four (4) flight spoilers located on the upper surface of each wing each hydraulic system A and B is dedicated to a different set of spoiler pairs to provide isolation and maintaining symmetric operation in the event of hydraulic system failure. Speed Brakes The speed brakes consist of flight spoilers and ground spoilers. Hydraulic system A powers all four ground spoilers, two on the upper surface of each wing. The SPEED BRAKE lever controls the spoilers. When the SPEED BRAKE lever is actuated all the spoilers extend when the airplane is on the ground and only the flight spoilers extend when the airplane is in the air. The SPEEDBRAKES EXTENDED light provides an indication of spoiler operation in-flight and on the ground. In-flight, the light illuminates to warn the crew that the speed brakes are extended while in the landing configuration or below 800 feet AGL. On the ground, the light illuminates when hydraulic pressure is sensed in the ground spoiler shutoff valve with the speed brake lever in the DOWN position. In-Flight Operation Operating the SPEED BRAKE lever in flight causes all flight spoiler panels to rise symmetrically to act as speed brakes. Caution should be exercised when deploying flight spoilers during a turn, as they greatly increase roll rate. When the speed brakes are in an intermediate position roll rates increase significantly. Moving the SPEED BRAKE lever beyond the FLIGHT DETENT causes buffeting and is prohibited in flight. A lever stop feature is incorporated into the SPEED BRAKE lever mechanism. The lever stop prevents the SPEED BRAKE lever from being moved beyond the FLIGHT DETENT when the airplane is in flight with the flaps up. In the event of the loss of electrical power the lever stop is removed and full speed brake lever movement is available.

Answer 40

An excessive differential hydraulic pressure is sensed in the elevator feel computer between system A and system B Elevator Feel System The elevator feel computer provides simulated aerodynamic forces using airspeed (from the elevator pitot system) and stabilizer position. Feel is transmitted to the control columns by the elevator feel and centering unit. To operate the feel system the elevator feel computer uses either hydraulic system A or B pressure, whichever is higher. When either hydraulic system or elevator feel pitot system fails, excessive differential hydraulic pressure is sensed in the elevator feel computer and the FEEL DIFF PRESS light illuminates.

Answer 41

Flap 10 through 40 positions. Flap Load Relief The flaps/slat electronics unit (FSEU) provides a TE flap load relief function which protects the flaps from excessive air loads. This function is operative at the flaps 10, 15, 25, 30 and flaps 40 positions. The FLAP lever does not move, but the flap position indicator displays flap retraction and re–extension. When the flaps are set at 40, the TE flaps: • retract to 30 if airspeed exceeds 163 knots • re–extend when airspeed is reduced below 158 knots. When the flaps are set at 30, the TE flaps: • retract to 25 if the airspeed exceeds 176 knots • re–extend when airspeed is reduced below 171 knots. When the flaps are set at 25, the TE flaps: • retract to 15 if the airspeed exceeds 191 knots • re–extend when airspeed is reduced below 186 knots. When the flaps are set at 15, the TE flaps: • retract to 10 if the airspeed exceeds 201 knots • re–extend when airspeed is reduced below 196 knots. When the flaps are set at 10, the TE flaps: • retract to 5 if the airspeed exceeds 211 knots • re–extend when airspeed is reduced below 206 knots

Answer 42

When any gear strut compresses Ground Operation During landing, the auto speed brake system operates when these conditions occur: • SPEED BRAKE lever is in the ARMED position • SPEED BRAKE ARMED light is illuminated • radio altitude is less than 10 feet • landing gear strut compresses on touchdown Note: Compression of any landing gear strut enables the flight spoilers to deploy. Compression of the right main landing gear strut enables the ground spoilers to deploy. • both thrust levers are retarded to IDLE • main landing gear wheels spin up (more than 60 kts). The SPEED BRAKE lever automatically moves to the UP position and the spoilers deploy. If a wheel spin-up signal is not detected, when the air/ground system senses ground mode (any gear strut compresses) the SPEED BRAKE lever moves to the UP position and flight spoiler panels deploy automatically. When the right main landing gear strut compresses, a mechanical linkage opens the ground spoiler bypass valve and the ground spoilers deploy. If the SPEED BRAKE lever is in the DOWN position during landing or rejected takeoff, the auto speed brake system operates when these conditions occur: • main landing gear wheels spin up (more than 60 kts) • both thrust levers are retarded to IDLE • reverse thrust levers are positioned for reverse thrust. The SPEED BRAKE lever automatically moves to the UP position and spoilers deploy. After an RTO or landing, if either thrust lever is advanced, the SPEED BRAKE lever automatically moves to the DOWN detent and all spoiler panels retract. The spoiler panels may also be retracted by manually moving the SPEED BRAKE lever to the DOWN detent.

Answer 43

A needle split is displayed on the flap position indicator Asymmetry and Skew Detection, Protection and Indication The FSEU monitors the TE flaps for asymmetry and skew conditions. It also monitors the LE devices for improper position and skew conditions on slats 2 through 7. If a flap on one wing does not align with the symmetrical flap on the other wing, there is a flap asymmetry condition. A skew condition occurs when a TE flap or LE slat panel does not operate at the same rate causing the panel to twist during extension or retraction. Trailing Edge Flap Asymmetry and Skew When the FSEU detects a trailing edge asymmetry or skew condition the FSEU: • closes the TE flap bypass valve • displays a needle split on the flap position indicator. Leading Edge Device Improper Position or Skew When the FSEU detects a LE device in an improper position or a LE slat skew condition, the LE FLAPS TRANSIT light remains illuminated and one of the following indications is displayed on the LE DEVICES annunciator panel: • amber TRANSIT light illuminated • incorrect green EXT or FULL EXT light illuminated • no light illuminated. There is no skew detection of the outboard slats, 1 and 8, or for the LE flaps. Slat skew detection is inhibited during autoslat operations

Answer 44

Hydraulic system B - Hydraulic system A through a PTU when a loss of pressure is sensed from Hydraulic system B EDP after take- off. Autoslats Autoslat operation is normally powered by hydraulic system B. An alternate source of power is provided by system A through a power transfer unit (PTU) if a loss of pressure is sensed from the higher volume system B engine driven pump. The PTU uses system A pressure to power a hydraulic motorized pump, pressurizing system B fluid to provide power for the autoslat operation. (Refer to Chapter 13, Hydraulics, Power Transfer Unit) At flap positions 1, 2, 5, 10, 15, and 25 an autoslat function is available that moves the LE slats to full extended if the airplane approaches a stall condition. The autoslat system is designed to enhance airplane stall characteristics at high angles of attack during takeoff or approach to landing. When TE flaps 1 through 25 are selected, the LE slats are in the extend position. As the airplane approaches the stall angle, the slats automatically begin driving to the full extended position prior to stick shaker activation. The slats return to the extend position when the pitch angle is sufficiently reduced below the stall critical attitude. Power Transfer Unit - Chapter 13, Hydraulics, Power Transfer Unit The purpose of the PTU is to supply the additional volume of hydraulic fluid needed to operate the autoslats and leading edge flaps and slats at the normal rate when system B engine–driven hydraulic pump volume is lost. The PTU uses system A pressure to power a hydraulic motor–driven pump, which pressurizes system B hydraulic fluid. The PTU operates automatically when all of the following conditions exist: • system B engine–driven pump hydraulic pressure drops below limits • airborne • flaps not up.

Answer 45

When control wheel is displaced more than approximately 10 degrees of roll Flight Spoilers Four flight spoilers are located on the upper surface of each wing. Each hydraulic system, A and B, is dedicated to a different set of spoiler pairs to provide isolation and maintain symmetric operation in the event of hydraulic system failure. Hydraulic pressure shutoff valves are controlled by the two flight SPOILER switches. Flight spoiler panels are used as speed brakes to increase drag and reduce lift, both in flight and on the ground. The flight spoilers also supplement roll control in response to control wheel commands. A spoiler mixer, connected to the aileron cable-drive, controls the hydraulic power control units on each spoiler panel to provide spoiler movement proportional to aileron movement. The flight spoilers rise on the wing with up aileron and remain faired on the wing with down aileron. When the control wheel is displaced more than approximately 10°, spoiler deflection is initiated.

Answer 46

Flap extended. Stabilizer Trim Stabilizer trim switches on each control wheel actuate the electric trim motor through the main electric stabilizer trim circuit when the airplane is flown manually. With the autopilot engaged, stabilizer trim is accomplished through the autopilot stabilizer trim circuit. The main electric and autopilot stabilizer trim have two speed modes: high speed with flaps extended and low speed with flaps retracted. If the autopilot is engaged, actuating either pair of stabilizer trim switches automatically disengages the autopilot. The stabilizer trim wheels rotate whenever electric stabilizer trim is actuated. The STAB TRIM MAIN ELECT cutout switch and the STAB TRIM AUTOPILOT cutout switch, located on the control stand, are provided to allow the autopilot or main electric trim inputs to be disconnected from the stabilizer trim motor. Control column actuated stabilizer trim cutout switches stop operation of the main electric and autopilot trim when the control column movement opposes trim direction. When the STAB TRIM override switch is positioned to OVERRIDE, electric trim can be used regardless of control column position. Manual stabilizer control is accomplished through cables which allow the pilot to position the stabilizer by rotating the stabilizer trim wheels. The stabilizer is held in position by two independent brake systems. Manual rotation of the trim wheels can be used to override autopilot or main electric trim. The effort required to manually rotate the stabilizer trim wheels may be higher under certain flight conditions. Grasping the stabilizer trim wheel will stop stabilizer motion.

Answer 47

Illuminates when the radio altimeter is less than 800 feet, or the TE flaps are extended beyond flap 10 Speed Brakes The speed brakes consist of flight spoilers and ground spoilers. Hydraulic system A powers all four ground spoilers, two on the upper surface of each wing. The SPEED BRAKE lever controls the spoilers. When the SPEED BRAKE lever is actuated all the spoilers extend when the airplane is on the ground and only the flight spoilers extend when the airplane is in the air. The SPEEDBRAKES EXTENDED light provides an indication of spoiler operation in-flight and on the ground. In-flight, the light illuminates to warn the crew that the speed brakes are extended while in the landing configuration orbelow 800 feet AGL. On the ground, the light illuminates when hydraulic pressure is sensed in the ground spoiler shutoff valve with the speed brake lever in the DOWN position.

Answer 48

All LE Devices EXTEND light illuminated

Answer 49

By the fuel LOW indication when both indications exist. Fuel Imbalance (IMBAL) Alert Displayed (amber) – • main tanks differ by more than 453 kgs • displayed below main tank with lower fuel quantity • inhibited when airplane is on ground • inhibited by fuel LOW indication when both indications exist • displayed until imbalance is reduced to 91 kgs The fuel quantity digits on tank with lower fuel quantity turn amber

Answer 50

Either tank; 2.000 lbs (907 kgs). Fuel LOW Alert Displayed (amber) – • fuel tank quantity less than 907 kgs in related main tank • display remains until fuel tank quantity is increased to 1134 kgs The fuel quantity digits on tank(s) with low fuel quantity turn amber.

Answer 51

No fuel alert indication will be displayed Fuel Configuration (CONFIG) Alert Displayed (amber) – • either engine running • center fuel tank quantity greater than 726 kgs; and • both center fuel tank pump switches positioned OFF The quantity digits on the center tank fuel quantity indicator turn amber. Display remains until – • both engines not running • center fuel tank quantity less than 363 kgs • one center fuel tank pump switch ON The quantity digits on the center tank fuel quantity indicator return to normal.

Answer 52

Main tank differ by more than 1.000 lbs / 453 kgs. Fuel Imbalance (IMBAL) Alert Displayed (amber) – • main tanks differ by more than 453 kgs • displayed below main tank with lower fuel quantity • inhibited when airplane is on ground • inhibited by fuel LOW indication when both indications exist • displayed until imbalance is reduced to 91 kgs The fuel quantity digits on tank with lower fuel quantity turn amber

Answer 53

Center tank fuel is greater than 1.600 lbs (726 kgs), both center tank pumps pressure is low and either engine is running. Fuel Configuration (CONFIG) Alert Displayed (amber) – • either engine running • center fuel tank quantity greater than 726 kgs; and • both center fuel tank pump switches positioned OFF The quantity digits on the center tank fuel quantity indicator turn amber. Display remains until – • both engines not running • center fuel tank quantity less than 363 kgs • one center fuel tank pump switch ON The quantity digits on the center tank fuel quantity indicator return to normal.

Answer 54

No. 1 main fuel tank is about 1⁄2 depleted and the forward boost pump in tank 1 is operating. Center Tank Fuel Scavenge Jet Pump With the main tank fuel pump No. 1 FWD Switch ON, the center tank fuel scavenge jet pump operates automatically to transfer any remaining center tank fuel to main tank No. 1. Fuel transfer begins when main tank No. 1 quantity is about one- half. Once the fuel scavenge process begins, it continues for the remainder of the flight.

Answer 55

Illuminated bright blue Crossfeed VALVE OPEN Light Extinguished – crossfeed valve is closed. Illuminated (blue) – • bright – crossfeed valve is in transit, or valve position and CROSSFEED selector disagree. • dim – crossfeed valve is open.

Answer 56

System B ``` System B • ailerons • elevator and elevator feel • leading edge flaps and slats • No. 2 thrust reverser • alternate nose wheel steering • autoslats • trailing edge flaps • rudder • flight spoilers (two on each wing) • normal brakes • autopilot B • landing gear transfer unit. • yaw damper ```

Answer 57

The fluid quantity is at the top of the standpipe System A Hydraulic Leak If a leak develops in the engine–driven pump or its related lines, a standpipe in the reservoir prevents a total system fluid loss. With fluid level at the top of the standpipe, the reservoir quantity displayed indicates approximately 20% full. System A hydraulic pressure is maintained by the electric motor–driven pump. If a leak develops in the electric motor– driven pump or its related lines, or components common to both the engine and electric motor–driven pumps, the quantity in the reservoir steadily decreases to zero and all system pressure is lost.

Answer 58

The standby electric motor driven pump is activated Standby Hydraulic System The standby hydraulic system is provided as a backup if system A and/or B pressure is lost. The standby system can be activated manually or automatically and uses a single electric motor–driven pump to power: • thrust reversers • rudder • leading edge flaps and slats (extend only) • standby yaw damper. Manual Operation Positioning either FLT CONTROL switch to STBY RUD: • activates the standby electric motor–driven pump • shuts off the related hydraulic system pressure to ailerons, elevators and rudder by closing the flight control shutoff valve • opens the standby rudder shutoff valve • deactivates the related flight control LOW PRESSURE light when the standby rudder shutoff valve opens • allows the standby system to power the rudder and thrust reversers. • illuminates the STBY RUD ON, Master Caution, and Flight Controls (FLT CONT) lights. Positioning the ALTERNATE FLAPS master switch to ARM, (refer to Chapter 9, Flight Controls for a more complete explanation): • activates the standby electric motor–driven pump • closes the trailing edge flap bypass valve • arms the ALTERNATE FLAPS position switch • allows the standby system to power the leading edge flaps and slats and thrust reversers.

Answer 59

72% Standby Hydraulic System Leak If a leak occurs in the standby system, the standby reservoir quantity decreases to zero. The LOW QUANTITY light illuminates when the standby reservoir is approximately half empty. System B continues to operate normally, however, the system B reservoir fluid level indication decreases and stabilizes at approximately 72% full.

Answer 60

Loss of system A or B, flaps extended and airborne, or wheel speed greater than 60kts Automatic Operation Automatic operation is initiated when the following conditions exist: • loss of system A or B, and • flaps extended, and • airborne, or wheel speed greater than 60 kts, and • FLT CONTROL switch A or B Hydraulic System ON OR: • the main PCU Force Fight Monitor (FFM) trips Automatic operation: • activates the standby electric motor–driven pump • opens the standby rudder shutoff valve • allows the standby system to power the rudder and thrust reversers. • illuminates the STBY RUD ON, Master Caution, and Flight Controls (FLT CONT) lights.

Answer 61

The ailerons and elevators

Answer 62

Supplies additional volume of hydraulic fluid to operate the autoslats and leading edge flaps and slats at the normal rate when system B engine-driven pump volume is lost Power Transfer Unit The purpose of the PTU is to supply the additional volume of hydraulic fluid needed to operate the autoslats and leading edge flaps and slats at the normal rate when system B engine–driven hydraulic pump volume is lost. The PTU uses system A pressure to power a hydraulic motor–driven pump, which pressurizes system B hydraulic fluid. The PTU operates automatically when all of the following conditions exist: • system B engine–driven pump hydraulic pressure drops below limits • airborne • flaps not up.

Answer 63

Autopilot A Standby Hydraulic System The standby hydraulic system is provided as a backup if system A and/or B pressure is lost. The standby system can be activated manually or automatically and uses a single electric motor–driven pump to power: • thrust reversers • rudder • leading edge flaps and slats (extend only) • standby yaw damper.

Answer 64

Leading edge flaps and slats System B • ailerons • rudder • elevator and elevator feel • flight spoilers (two on each wing) • leading edge flaps and slats • normal brakes • No. 2 thrust reverser • autopilot B • alternate nose wheel steering • landing gear transfer unit. • autoslats • yaw damper • trailing edge flaps

Answer 65

No. 1 engine RPM drops below a limit value Landing Gear Transfer Unit The purpose of the landing gear transfer unit is to supply the volume of hydraulic fluid needed to raise the landing gear at the normal rate when system A engine–driven pump volume is lost. The system B engine–driven pump supplies the volume of hydraulic fluid needed to operate the landing gear transfer unit when all of the following conditions exist: • airborne • No. 1 engine RPM drops below a limit value • landing gear lever is positioned UP • either main landing gear is not up and locked.

Answer 66

It may impact a fitting in the wheel well opening causing the gear to free-fall back to down position Landing Gear Retraction When the LANDING GEAR lever is moved to UP, the landing gear begins to retract. During retraction, the brakes automatically stop rotation of the main gear wheels. After retraction, the main gear are held in place by mechanical uplocks. Rubber seals and oversized hubcaps complete the fairing of the outboard wheels. The nose wheels retract forward into the wheel well and nose wheel rotation is stopped by snubbers. The nose gear is held in place by an overcenter lock and enclosed by doors which are mechanically linked to the gear. Hydraulic pressure is removed from the landing gear system with the LANDING GEAR lever in the OFF position. If a main landing gear tire is damaged during takeoff, it is possible that braking of the main gear wheels during retraction may be affected. A spinning tire with a loose tread must be stopped prior to entering the wheel well or it can cause damage to wheel well components. When a spinning tire with loose tread impacts a fitting in the wheel well ring opening, that gear stops retracting and free falls back to the down position. The affected gear cannot be retracted until the fitting is replaced.

Answer 67

When thrust levers are retracted to idle at or above 90 knots Autobrake System The autobrake system uses hydraulic system B pressure to provide maximum deceleration for rejected takeoff and automatic braking at preselected deceleration rates immediately after touchdown. The system operates only when the normal brake system is functioning. Antiskid system protection is provided during autobrake operation. Rejected Takeoff (RTO) The RTO mode can be selected only when on the ground. Upon selection, the AUTO BRAKE DISARM light illuminates for one to two seconds and then extinguishes, indicating that an automatic self–test has been successfully accomplished. To arm the RTO mode prior to takeoff the following conditions must exist: • airplane on the ground • antiskid and autobrake systems operational • AUTO BRAKE select switch positioned to RTO • wheel speed less than 60 knots • forward thrust levers positioned to IDLE. With RTO selected, if the takeoff is rejected prior to wheel speed reaching 90 knots autobraking is not initiated, the AUTO BRAKE DISARM light does not illuminate and the RTO autobrake function remains armed. If the takeoff is rejected after reaching a wheel speed of 90 knots, maximum braking is applied automatically when the forward thrust levers are retarded to IDLE.

Answer 68

Applying manual brakes Landing When a landing autobrake selection is made, the system performs a turn–on– self–test. If the turn–on–self–test is not successful, the AUTO BRAKE DISARM light illuminates and the autobrake system does not arm. Four levels of deceleration can be selected for landing. However, on dry runways, the maximum autobrake deceleration rate in the landing mode is less than that produced by full pedal braking. After landing, autobrake application begins when: • both forward thrust levers are retarded to IDLE • the main wheels spin–up. Note: Landing autobrake settings may be selected after touchdown prior to decelerating through 30 kts of ground speed. Braking initiates immediately if the above conditions are met. To maintain the selected landing deceleration rate, autobrake pressure is reduced as other controls, such as thrust reversers and spoilers, contribute to total deceleration. The deceleration level can be changed (without disarming the system) by rotating the selector. The autobrake system brings the airplane to a complete stop unless the braking is terminated by the pilot. Autobrake – Disarm The pilots may disarm the autobrake system by moving the selector switch to the OFF position. This action does not cause the AUTO BRAKE DISARM light to illuminate. After braking has started, any of the following pilot actions disarm the system immediately and illuminate the AUTO BRAKE DISARM light: • moving the SPEED BRAKE lever to the down detent • advancing the forward thrust lever(s), except during the first 3 seconds after touchdown for landing • applying manual brakes.

Answer 69

Landing gear retraction is disabled Landing Gear Manual Extension If hydraulic system A pressure is lost, the manual extension system provides another means of landing gear extension. Manual gear releases on the flight deck are used to release up locks that allow the gear to free–fall to the down and locked position. The forces that pull the gear down are gravity and air loads. With the manual extension access door open: • manual landing gear extension is possible with the LANDING GEAR lever in any position • normal landing gear extension is possible if hydraulic system A pressure is available • landing gear retraction is disabled. Following a manual extension, the landing gear may be retracted normally by accomplishing the following steps: • close the manual extension access door • move the LANDING GEAR lever to DOWN with hydraulic system A pressure available, and then • position the LANDING GEAR lever to UP.

Answer 70

Auto braking is not initiated With RTO selected, if the takeoff is rejected prior to wheel speed reaching 90 knots autobraking is not initiated, the AUTO BRAKE DISARM light does not illuminate and the RTO autobrake function remains armed. If the takeoff is rejected after reaching a wheel speed of 90 knots, maximum braking is applied automatically when the forward thrust levers are retarded to IDLE.

Answer 71

7° Rudder/Brake Pedals Push full pedal – turns nose wheel up to 7 degrees in either direction. Push top of pedal only – activates wheel brakes. Refer to Chapter 9 Flight Controls for rudder description. Nose Wheel Steering Wheel Rotate – • turns nose wheel up to 78 degrees in either direction. Note: Refer to Chapter 1 for effective steering angle and turning radius. • overrides rudder pedal steering. Nose Wheel Steering Nose wheel steering is available when the nose gear is in the down position and compressed by weight of the airplane. Positioning the landing gear control lever to down makes system A hydraulic pressure available to the steering metering valve. Alternate nose wheel steering can be activated to provide system B pressure to the nose wheels when the NOSE WHEEL STEERING switch is placed to ALT, normal quantity is in the system B reservoir, and the airplane is on the ground. In the event of a hydraulic leak downstream of the Landing Gear Transfer Unit, resulting in a loss of hydraulic system B fluid in the reservoir, a sensor closes the Landing Gear Transfer Valve and alternate steering will be lost. Primary steering is controlled through the nose wheel steering wheel. Limited steering control is available through the rudder pedals. A pointer on the nose steering wheel assembly shows nose wheel steering position relative to the neutral setting. Rudder pedal steering is deactivated as the nose gear strut extends. A lockout pin may be installed in the towing lever to depressurize nose wheel steering. This allows airplane pushback or towing without depressurizing the hydraulic systems.

Answer 72

Aircraft position Look–Ahead Terrain Alerting The GPWS terrain data base contains detailed terrain data near major airports, and data in lesser detail for areas between airports. Terrain within 2,000 feet of airplane barometric altitude shows on the navigation display. The terrain data is not designed to be an independent navigation aid. The terrain display is generated from a data base contained in the GPWS computer and correlated to GPS position. Terrain and weather radar cannot show together on a display. If one pilot selects terrain and the other pilot selects weather radar, each display updates on alternating sweeps. All other displays (TCAS, LNAV routing, etc.) can show with terrain data. Look-ahead terrain alerts are based on the airplane’s position, barometric altitude, vertical flight path, and ground speed

Answer 73

In flight below 2.300 ft R.A. The weather radar automatically begins scanning for windshear when: • thrust levers set for takeoff, even if engine is off or IRS not aligned, or • in flight below 2,300 feet RA (predictive windshear alerts are issued below 1,200 feet RA). Alerts are available approximately 12 seconds after the weather radar begins scanning for windshear. Predictive windshear alerts can be enabled prior to takeoff by pushing the EFIS control panel WXR switch. If windshear is not detected, weather radar returns show only after pushing the EFIS control panel WXR switch.

Answer 74

Spoilers not down with the speedbrakes lever in the DOWN position Intermittent Cabin Altitude/Configuration Warning Takeoff configuration warning is armed when the airplane is on the ground and either or both forward thrust levers are advanced for takeoff. Takeoff configuration warning activates if: • trailing edge flaps are not in the flaps 1 through 25 takeoff range, or • trailing edge flaps are in a skew or asymmetry condition, or have uncommanded motion, or • leading edge devices are not configured for takeoff or have uncommanded motion, or • speed brake lever is not in the DOWN position, or • spoiler control valve is open providing pressurized hydraulic fluid to the ground spoiler interlock valve, or • parking brake is set, or • stabilizer trim not set in the takeoff range. An intermittent warning horn sounds and the TAKEOFF CONFIG warning light illuminates when takeoff configuration warning activates. Cabin altitude warning activates when cabin altitude exceeds 10,000 feet. An intermittent warning horn sounds and the CABIN ALTITUDE warning light illuminates. The warning horn may be silenced by momentarily pressing the ALT HORN CUTOUT switch on the Cabin Altitude Panel. The warning light remains illuminated until the cabin altitude descends below 10,000 feet. WARNING: The Cabin Altitude and Takeoff Configuration Warnings use the same intermittent tone when activated.

Answer 75

Below 1.500 feet AGL

Answer 76

Other aircraft are approximately 40 seconds from closet point of approach Traffic Alert and Collision Avoidance System (TCAS) TCAS alerts the crew to possible conflicting traffic. TCAS interrogates operating transponders in other airplanes, tracks the other airplanes by analyzing the transponder replies, and predicts the flight paths and positions. TCAS provides advisory and traffic displays of the other airplanes to the flight crew. Neither advisory, guidance, nor traffic display is provided for other airplanes which do not have operating transponders. TCAS operation is independent of ground–based air traffic control. To provide advisories, TCAS identifies a three dimensional airspace around the airplane where a high likelihood of traffic conflict exists. The dimensions of this airspace are based upon the closure rate with conflicting traffic. TCAS equipment interrogates the transponders of other airplanes to determine their range, bearing, and altitude. A traffic advisory (TA) is generated when the other airplane is approximately 40 seconds from the point of closest approach. If the other airplane continues to close, a resolution advisory (RA) is generated when the other airplane is approximately 25 seconds from the point of closest approach. The RA provides aural warning and guidance as well as maneuver guidance to maintain or increase separation from the traffic. Non–transponder equipped airplanes are invisible to TCAS. RAs can be generated if the other airplane has a mode C transponder. Coordinated RAs require both airplanes to have TCAS

Answer 77

When thrust lever is retarded to idle and cannot be silenced Landing Gear Configuration Warnings Visual indications and aural warnings of landing gear position are provided by the landing gear indicator lights and landing gear warning horn. Visual Indications The landing gear indication lights are activated by signals from each gear, the LANDING GEAR lever, and the forward thrust lever position as follows: Green light illuminated – landing gear is down and locked. Red light illuminated – • landing gear is in disagreement with LANDING GEAR lever position (in transit or unsafe). • landing gear is not down and locked (with either or both forward thrust levers retarded to idle, and below 800 feet AGL). All lights extinguished – landing gear is up and locked with the LANDING GEAR lever UP or OFF.

Answer 78

158 feet AGL.

Answer 79

10.000 ft.

Answer 80

3° C above the freezing point of the fuel being used or -43° C, whichever is higher.

Answer 81

78 degree 15 minutes.

Answer 82

41.000 ft.

Answer 83

17.000 ft.

Answer 84

453KG (1000 LBS)

Answer 85

Degrees Celsius above the fuel freezing point or -43 c whichever is higher

Answer 86

After landing, the ground personal should be informed not to open the cargo door until all supernumeraries and crew have exited the airplane and fire fighting equipment is nearby.

Answer 87

If force is applied to the First Officer’s control wheel and the aircraft rolls, the pilot has control through the spoilers, or if force is applied to the Captain’s control wheel and the aircraft rolls, the pilot has control though the ailerons.

Answer 88

DC power is lost The overwing emergency exits lock when: • three of the four Entry/Service doors are closed and • either engine is running and • the airplane air/ground logic indicates that the airplane is in the air or both thrust levers are advanced. The overwing emergency exits unlock when any one of the above conditions is not met or DC power is lost.

Answer 89

OXYGEN MASK AND REGULATORS 0N, 100%

Answer 90

Lowest Safety Altitude or 10.000FT, whichever is higher

Answer 91

ENGINE FIRE WARNING Switch………..PULL.

Answer 92

Auto-throttle Disengage, thrust lever retard until indications remain within appropriate limits or the thrust lever is closed. Read Engine Limit / Surge / Stall Checklist.

Answer 93

Disengage the autopilot and Auto Throttle. Apply max thrust and roll wing level. Rotate to an initial pitch of 20 degrees. Retract speed brakes. Do not change gear or flap configuration until Terrain Separation is assured.

Answer 94

Control column hold firmly

Answer 95

If a flight lock has failed locked or a fault is detected the PSEU light, the OVERHEAD annunciator, and the MASTER CAUTION lights illuminate.These indications are inhibited from takeoff until 30 seconds after the airplane is in the ground mode. When the doors are latched and locked and the flight lock is operating properly none of these lights will illuminate

Answer 96

Automatically illuminates the emergency lights, if airplane electrical power to DC bus No. 1 fails or AC power is turned off

Answer 97

Adjust Airplane Attitude / Thrust, check that Probe heat is ON and cross check Mach/ Airspeed Indicators

Answer 98

Select manual pressurization, open OUTFLOW VALVE, and land at nearest suitable airport.

Answer 99

Jammed or Restricted Elevator

Answer 100

Disconnect Autopilot and Auto Throttle. Apply max thrust. Roll wings level and rotate toward 15° pitch. Retract speed brakes. Follow Flight Director TO/GA. Do not change flap or gear configuration until windshear is no longer a factor.

Answer 101

Disengage Auto Pilot and Auto Throttle. Adjust Pitch and Thrust to satisfy RA command

Answer 102

Stabilizer and rudder trim only

Answer 103

A/T disengage; Thrust Lever (affected engine) – confirm – CLOSE; if ENG OVERHEAT light stays illuminated, go to ENGINE FIRE or SEVERE DAMAGE/SEPARATION checklist.

Answer 104

DEU 2 automatically supplies data to all six DU’s. On both PDF’s the message DSPLY SOURCE is displayed

Answer 105

The UP bug is displayed independent of flap position up to approximately 20.000’

Answer 106

After VREF is selected on the APPROACH REF page

Answer 107

If the APP mode is selected on the ND with a VOR frequency tuned

Answer 108

An amber circle (filled)

Answer 109

The predicted airspeed within 10 seconds, based on present airspeed and acceleration

Answer 110

A speed which provides a 1.3 G margin over the stick shaker speed at unaccelerated flight

Answer 111

a. The fo’s efis control panel is supplying identical inputs to the captains ans fo’s displays

Answer 112

a. Non dispatchable CDS fault has occurred

Answer 113

The aircraft descent through the selected minimum altitude

Answer 114

The Captain’s outboard DU has failed. Which statement is correct? a. PFD move to inboard DU, the ND can be selected on the lower DU using the lower DU control panel select switch.

Answer 115

The barometric display is boxed and turns AMBER

Answer 116

a. DME distance is displayed on the PFD, ND and APP mode.

Answer 117

When TERR display selectet switch on the EFIS control panel is activated and automatically when neither pilot has terrain selected in EXP MAP, Center MAP, EXP VOR or EXP APP modes.

Answer 118

a. A red square (filled)

Answer 119

a. A speed that give a 1.3 G margin to high speed buffet

Answer 120

The placard speed for flap 15

Answer 121

Position, Altitude, Speed, Attitude.

Answer 122

Below 2.500 ft RA and will rise toward airplane symbol at 200ft. RA.