BOEING 737-300/400/500

The cockpit door. Can only be unlocked electrically . Can only be locked with A.C. power. Can only be locked with a key .

The cockpit door blow-out panels. Open into the cockpit. Open into the cabin . Open either way for pressure equalization .

To exit the cockpit with the cockpit door jammed closed. Grasp the cockpit door emergency exit handle on the lower part of the door and rotate clockwise . Grasp the cockpit door emergency exit handle on the lower part of the door and push. Grasp the cockpit door emergency exit handle on the upper part of the door and pull forward.

When turning the B737-300 on the ground. The landing gear geometry and sweep back of the wings result in the tail arc being less than the wing tip arc. The landing gear geometry and sweep back of the wings result in the tail arc being greater than the wing tip arc. The landing gear geometry and sweep back of the wings result in the tail arc being greater than the wing tip arc.

When turning the B737-300 on the ground. The landing gear geometry and sweep back of the wings result in an inward motion of the wing tips and tail . The landing gear geometry and sweep back of the wings result in an inward motion of the tail only . The landing gear geometry and sweep back of the wings result in an outward motion of the wing tips and tail.

When turning the B737-300 on the ground the minimum pavement width for an 180 degree turn is. 65 feet (19.81 meters ). 55 feet (16.76 meters.). 60 feet (18.29 meters.).

The cargo compartments are. Pressurized approximately equal to cabin pressure. Not pressurized . Pressurized to their normal differential pressure of 2.0 P.S.I.

In case of sudden loss of aircraft pressurization, the pressure relief of the cargo compartments is accomplished by. Pressure relief valves set at 7.45 P.S.I.D . Pressure relief valves set at 8.65 P.S.I.D . Blow-out panels.

To use the APU for air conditioning, on the ground/engines shut down, -3/4/500 series aircraft, you should select?. Isolation Valve Switch AUTO APU Bleed Air Switch ON Left or Right Air Conditioning Pack Switch AUTO or HIGH. Isolation Valve Switch OPEN. APU Bleed Air Switch ON. Left and Right Air Conditioning Pack Switch AUTO or HIGH. APU Bleed Air Switch ON. Left and Right Air Conditioning Pack Switch AUTO or HIGH. Isolation Valve Switch OPEN. APU Bleed Air Switch ON. Left and Right Air Conditioning Pack Switch AUTO or HIGH . Isolation Valve Switch CLOSED. APU Bleed Air Switch ON. Left or Right Air Conditioning Pack Switch AUTO or HIGH .

The turbo-fan valve of an air conditioning unit opens?. When the aircraft is on the ground only . When the aircraft is on the ground or flaps are extended. Only when the RAM DOOR FULL OPEN light is illuminated.

If the Left DUCT OVERHEAT light illuminates on the B737-300/500 aircraft. The cargo compartment mixer valves will automatically program to the full hot position . The passenger cabin mixer valves will automatically program to the full cold position. The control cabin mixer valves will automatically program to the full cold position.

If the Right PACK TRIP OFF light illuminates: Select a warmer temperature on the control cabin Temperature Selector and press the TRIP RESET switch . Select a warmer temperature on the passenger cabin Temperature Selector and press the TRIP RESET switch. Press and hold the TRIP RESET switch for 30 seconds only.

The RAM DOOR FULL OPEN lights are normally illuminated. When on the ground, during slow flight with the flaps not fully retracted or anytime the landing gear is retracted. Only during slow flight with the flaps not fully retracted. When on the ground or during slow flight with the flaps not fully retracted.

The RAM DOOR FULL OPEN lights are normally extinguished. During the cruise. During the takeoff and climb. Just before landing.

The E & E compartment is cooled by?. The equipment cooling system. The ram air system . The AUTO or STANDBY pressurization systems.

The air supply for the Re-circulating Fan is?. Exhaust air from the main cabin and electrical equipment bay and forward outflow valve collected in a shroud located above the aft cargo compartment . Exhaust air from the main cabin and electrical equipment bay collected in a shroud located above the aft cargo compartment . Exhaust air from the main cabin and electrical equipment bay collected in a shroud located above the forward cargo compartment.

The APU may be operated with APU bleed only up to a maximum altitude of. 7,000ft. 10,000ft . 35,000ft .

Maximum continuous EGT for Garrett APU operation is?. 760 degrees C. 710 degrees C. 725 degrees C.

The APU LOW OIL PRESSURE light is. Disarmed when the APU switch is in the OFF position. Always illuminated when the APU switch is in the OFF position . Inhibited during APU start .

The APU will automatically shut-down. When the battery switch is placed OFF at any time . When the battery switch is placed OFF on the ground only. When the battery switch is placed OFF in flight only .

The overspeed light will illuminate when a start is aborted prior to the APU reaching normal operating speed and. The APU overspeed reset switch in the E & E compartment must be reset . Will extinguish when the APU switch is placed to START again. No further attempts should be made to start the APU .

Should the APU switch fail to shut-off the APU. The APU Overspeed Reset switch in the E & E compartment has been tripped and must be reset . Trip the APU fuel valve CB . Pull up the APU Fire Warning Switch.

If the APU runs down due to fuel starvation, the Master Caution annunciation system will show. APU only . APU and services being operated by the APU . Services being operated by the APU only.

The APU fire warning system gives aural and visual warnings. Only in the flight decK . In the flight deck and the APU compartment . In the flight deck and main wheel well.

The STAB OUT OF TRIM light operates. At any time . Only when the autopilot is engaged. Only when the speed trim system is activated .

If the autopilot disengage light is flashing amber, this indicates that. The autopilot has reverted to CWS pitch or roll while in CMD. The autopilot has reverted to CMD pitch or roll while in CWS . The light test switch is being held in position 2 .

If the auto throttle disengagement light is flashing amber, this indicates. An auto throttle airspeed error if speed is not held within +15 or -15 knots of commanded speed when in flight, flaps up and autothrottle engaged in MCP SPD or FMC SPD mode . The light test switch is being held in position 2 . An autothrottle airspeed error if speed is not held within +10 or -5 knots of commanded speed when in flight, flaps not up and autothrottle engaged in MCP SPD or FMC SPD mode.

VNAV mode is terminated by. Selecting a different pitch mode or Glideslope capture or De-selecting LNAV or Extending the wing flaps beyond 25 degrees . Selecting a different pitch mode or Localiser capture or Extending the wing flaps beyond 5 degrees . Selecting a different pitch mode or Glideslope capture or Extending the wing flaps beyond 15 degrees.

If during a climb with autopilot engaged, the ALT HOLD switch is pressed, the aircraft will. Continue the climb to the preset altitude selected, as the altitude acquire mode is now armed . Stop the climb and level off. Continue the climb to the next 1000 feet level and level off.

With autopilot A engaged, the altitude selected on the MCP is referenced to. The standby altimeter . The Captain's altimeter. The First Officer's altimeter .

VNAV climbs and descents are. Constrained by the selected MCP altitude. Constrained by the selected cabin altitude . Not constrained by the selected MCP altitude .

When intercepting the ILS with APP mode armed. On all airplanes, localizer must be captured prior to glideslope . The second autopilot can be engaged . On some airplanes, glideslope may be captured prior to localizer.

With the Boom/Mask switch in the MASK position, transmission of a message is possible. By using the oxygen mask only . By using the oxygen mask or hand microphone only. By using the oxygen mask and headset only .

The VHF-1 transmitter selector switch on an audio selector panel is illuminated. Reception on VHF-1 is automatically provided. Reception is achieved by pulling and rotating VHF-1 receiver switch . The ALT/NORM switch must be in NORM to obtain reception at a comfortable volume level .

The Service Interphone switch on the Aft overhead panel when selected ON. Allows communications between the flight deck and the flight attendants when using the Flight Interphone system . Deactivates external jacks sockets from the Service Interphone system . Adds external jack sockets to the Service Interphone system.

Communications between the cockpit and the ground crew is possible by using the. Flight interphone system, or the Flight interphone system, or the Service Interphone system provided the Service Interphone switch is ON. The Service interphone system irrespective of the position of the Service Interphone switch . Flight interphone system only .

The GRD CALL switch on the Fwd Overhead panel when pressed. Sounds a horn in the nose wheel and main wheel wells until released . Sounds a horn in the nose wheel and main wheel wells until the ground crew select the GRD CALL CANCEL switch on the External Power receptacle panel . Sounds a horn in the nose wheel well until released.

The Cockpit CALL (blue) light will illuminate along with an associated chime whenever the cockpit is being called. By the flight attendants or the ground crew. By the flight attendants only . By the ground crew only.

The ATTEND (attendants call) is pressed and released in the cockpit and. A three tone chime sound will be heard in the passenger cabin . A two tone chime sound will be heard in the passenger cabin. A single tone chime sound will be heard in the passenger cabin .

The selcal system monitors selected frequencies in use on the. VHF-1 and VHF-2 communications radios . HF-1 and VHF-1 communications radios . HF-1 communications radio.

The STANDBY PWR OFF light (amber) illuminated means. The Battery Bus is inactive . The AC standby bus is inactive. The DC standby bus is inactive .

Generator Drive Disconnect switch when operated. Disconnects the Generator Drive from the engine in the event of a Generator Drive malfunction. Disconnects the Generator Drive from the generator in the event of a Generator Drive malfunction . Disconnects the Generator Drive from the engine in the event of a Generator Drive malfunction, only if the engine has been shut down first .

Recoupling of the Generator Drive drive shaft to the engine may be accomplished. At any time on the ground or in flight . At any time in flight provided the Generator Breaker and the Generator Control Relay have previously been tripped . Only on the ground.

After operating the Generator Drive Disconnect switch the Generator Drive HIGH OIL TEMP (if illuminated) will. Immediately extinguish as the power to the light comes from the Generator Drive Disconnect Switch . Remain on, until the Generator Drive oil has cooled. Flash repeatedly indicating successful Generator Drive uncoupling.

The Generator Drive LOW OIL PRESSURE lights are. Amber lights on the forward overhead panel which will illuminate when Generator Drive oil pressure is below minimum operating limits. Red lights on the forward overhead panel which will illuminate when Generator Drive oil pressure is below minimum operating limits . Amber lights on the forward overhead panel which will flash repeatedly when Generator Drive oil pressure is below minimum operating limits .

One of the basic principles of the electrical system is. The AC sources of power can be connected in parallel if necessary . There is no paralleling of the AC sources of power. The AC sources of power are always connected in parallel .

The AC STANDBY BUS power can be supplied from. The APU or engine generators, external power or the battery through the Static Inverter. The APU or engine generators only . The APU or engine generators, external power or directly from the Hot Battery Bus .

A Generator Breaker can be closed. When power quality from the generator is correct. At any time the engine is running at or above idle power . Irrespective of power quality from the generator .

Should a Generator Drive develop a fault and be disconnected from the engine, before corrective action is taken by the pilots. All electrical systems ( loads ) supplied by that generator will be lost A few non-essential electrical systems ( loads ) will be lost. No electrical systems ( loads ) supplied by that generator will be lost as the Bus . Transfer systems supplies all the aircraft's electrical systems (loads) .

When both engines are running with external power connected then. External power will automatically disconnect when either engine generator is brought on-line . External power will automatically disconnect when both engine generators are brought on line. External power must be selected OFF before the engine generators are .

The Engine Overheat and Fire Protection system has. Two dual element overheat/fire detection loops installed in each engine nacelle . Two overheat/fire detection loops, each consisting of four detector elements, installed in each engine nacelle. Four single element overheat/fire detection loops installed in each engine nacelle .

An overheat or fire alert will be given with the OVHT DET switch in NORMAL and both loops serviceable when. Either or both elements of a detector signal an overheat or fire condition . One of the detector elements of loop A and one of the detector elements of loop B sense a fire or overheat. Either element of a detector signals an overheat or fire condition .

The FAULT light illuminates. The other two answers are correct. If both loops sense a fault with the OVHT DET switch in the NORMAL position. If the OVHT DET switch is to 'A' or 'B' and the selected loop senses a fault .

With the OVHT DET switch in the B position. Either A or B detector loops can initiate an overheat or fire warning . Only the A detector loop can initiate an overheat or fire warning .the A detector loop can initiate an overheat or fire warning . Only the B detector loop can initiate an overheat or fire warning.

Holding the TEST switch to the FAULT/ INOP position. Tests the fault detection circuits for both engines and the APU. Tests the fault detection circuits for the APU only . Tests the fault detection circuits for both engines only .

Holding the TEST switch to the FAULT/ INOP position. Will illuminate the Master Caution, OVHT/DET annunciator, FAULT and APU DET INOP lights. Will illuminate the Master Caution, OVHT/DET annunciator, FAULT, WHEEL WELL and APU DET INOP lights . Will illuminate the Master Caution, FAULT annunciator and APU DET INOP light .

The Engine Fire Extinguisher system contain. Three freon bottles which may be discharged into both engines . Two freon bottles which may be discharged into either engine. Three freon bottles which may be discharged into both engines and the APU .

During the Engine Fire drill, the pneumatic ISOLATION VALVE switch should be selected to. AUTO . CLOSE. OPEN .

The power source for the Engine Overheat and Fire Detection is?. Hot Battery Bus (28V DC) . Battery Bus (28V DC). NO.1 Transfer Bus (115V AC) .

The Feel Differential Pressure amber light is. Armed when the trailing edge flaps are up. Armed when the trailing edge flaps are not up . Armed at all times .

A single mach trim FCC channel failure is indicated. By the immediate illumination of the Mach trim Fail light . By pressing to test the Mach Trim Fail light . By the illumination of the Mach Trim Fail light on a Master Caution annunciator recall.

The Yaw Damper operation results in. No rudder pedal movement provided the autopilot is engaged . Rudder pedal movemen . No rudder pedal movement.

Loss of hydraulic system B pressure. Does not cause yaw damper disengagement but will illuminate the amber YAW DAMPER light . Does not cause yaw damper disengagement or illumination of the amber YAW DAMPER light. Causes yaw damper disengagement and illumination of the YAW DAMPER light .

The Yaw Damper uses. Hydraulic system B pressure only. Both hydraulic systems A and B pressure . Hydraulic system A pressure only .

The Rudder Trim Control is. Spring-loaded to neutral and may be rotated left or right provided the autopilot is disengaged . Spring-loaded to neutral and may be rotated left or righ. Spring-loaded to neutral and must be pushed down before rotating left or right .

The Ground Spoilers are powered by. The A hydraulic system. The B hydraulic system . The A and B hydraulic system .

The Flap Load Limiter (B737-300) will retract the trailing edge flaps from the 40 position to the 30 position if the airspeed exceeds. 158 knots. 155 knots . 153 knots .

The primary flight controls are. Ailerons, Spoilers, Elevators and Rudder . Ailerons, Elevators and Rudder. Ailerons, Horizontal Stabiliser, Elevators and Rudder .

How many Air Data Computers are fitted to the B737. 2 (1 AC powered and 1 DC powered) . 3 (2 AC powered and 1 DC powered) . 2 (both AC powered).

The Flight Recorder will operate on the ground. When either engine is operating only . When the Test switch is selected to TEST or either engine is operating. At all times .

The alternate static system provides static pressure inputs to. The Standby Airspeed indicator / Standby Altimeter and No.2 ADC . The Standby Airspeed indicator / Standby Altimeter and No.1 ADC . The Standby Airspeed indicator / Standby Altimeter only.

The Weather Radar ON / OFF selector is located on the. EFIS control panel. Weather radar control panel . MCP control panel .

The standby altimeter indicator utilises the. The alternate static source. The normal static source . The No.1 ADC and the alternate static source.

The Standby Horizon is. Normally powered when any one or more generators are on line . Normally powered at all times even with the loss of all AC generators. Only powered from the No2 AC Transfer Bus .

The Standby Horizon ILS selector when selected to OFF. Retracts the ILS pointers and ILS failure flags from view. Retracts the ILS pointers only from view . Retracts the ILS failure flags from view only .

On the ground the TAT indication is approximately the outside air temperature. Provided the pitot heat is OFF. Provided the pitot heat is ON . Irrespective of the pitot heat being ON or OFF .

Before selecting the ILS test facility. Select an ILS frequency on the associated VHF NAV control panel . Select an ILS frequency on the both VHF NAV control panels . Select an ILS or VOR frequency on the associated VHF NAV control panel .

Fuel Quantity indicators will display. Blank if a malfunction occurs . An ERR symbol if a malfunction occurs. An 88888 reading if a malfunction occurs .

To defuel No. 1 tank. Select the No. 1 Main tank fuel pumps ON, the crossfeed valve CLOSED and the Manual Defuelling valve OPEN . Select the No. 1 and No. 2 Main tank fuel pumps ON, the Crossfeed valve OPEN and the Manual Defuelling valve OPEN . Select the No. 1 Main tank fuel pumps ON, the Crossfeed valve OPEN and the Manual Defuelling valve OPEN.

The Fuel Temperature Indicator reads the temperature of the fuel in. The No. 1 tank. The No. 2 tank . The Centre tank .

The manual De-fuelling valve is located. Inboard of the No. 2 engine . Outboard of the No. 2 engine. Outboard of the No. 1 engine .

The Fuel Temperature Indicator reads the temperature of the fuel in. The No. 1 tank. The No. 2 tank . The Centre tank .

The manual De-fuelling valve is located. Inboard of the No. 2 engine . Outboard of the No. 2 engine. Outboard of the No. 1 engine .

Dripsticks (or Floatsticks) are installed in. Classics: 4 in each main tank. NG's: 5 in each main tank and 2 in the centre tank . Classics: 6 in each main tank. NG's: 7 in each main tank and 6 in the centre tank . Classics: 5 in each main tank. NG's: 6 in each main tank and 4 in the centre tank.

The Fuelling Valve Position Lights on the External Fueling Panel illuminate (blue) when. The respective fuelling valve is in transit . The respective fuelling valve is OPEN . The respective fuelling valve is OPEN and fuel is being transferred into the tank.

When the APU is inoperative and no external power is available, refuelling can be accomplished as follows. Battery switch OFF Standby Power switch BAT The entire fuel system will operate normally . Battery switch ON Standby Power switch BAT The entire fuel system will operate normally. Battery switch ON Standby Power switch BAT The entire fuel system will operate normally, except for the fuel shut-off system .

Complete loss of system 'B' pressure will deactivate. The outboard flight spoilers. The inboard flight spoilers . The ground spoilers .

If an electric hydraulic pump OVERHEAT light comes on. The pump will be automatically switched off and the OVERHEAT LIGHT extinguished . Turn the associated system electric and engine hydraulic pumps OFF . Turn the associated system electric hydraulic pump OFF.

A leak in the hydraulic system B engine driven pump or its associated lines would be indicated by. System B contents falling to one quarter full (conventional engine instruments) or 25% (EIS) . System B contents falling to half full (conventional engine instruments) or 40% (EIS). System B contents falling to three quarters full (conventional engine instruments) or 64% (EIS) .

The standby system LOW PRESSURE light is armed. Only when the standby pump operation has been selected or automatic standby function is activated. At all times . Only when the standby pump operation has been selected or either Spoiler switch has been selected to OFF .

The standby hydraulic pump only supplies pressure to the. Standby rudder and leading edge slats . Standby rudder actuator, leading edge devices and thrust reversers. Standby rudder, leading edge devices and brakes .

If either Flight Control switch is moved to the STBY RUD position. The standby pump will be activated and the standby hydraulic LOW QUANTITY light will be armed . The standby pump will be de-activated allowing system A pressure to power the rudder and the standby hydraulic LOW PRESSURE light will be armed . The standby pump will be activated and the standby hydraulic LOW PRESSURE light will be armed.

Low fluid quantity in the A system reservoir can be indicated by one of the following. Mechanical indication on the reservoir. LOW QUANTITY light on the centre instrument panel . LOW QUANTITY light on the centre instrument panel along with a Master Caution ENG annunciator light .

If a total failure of both pumps supplying system B pressure occurs, which of the primary flight controls will be totally inoperative. None. Aileron and elevator . Rudder only .

The A and B hydraulic reservoirs are pressurised by. Air from the 14th stage only . Air from the pneumatic manifold. Hydraulic fluid from the standby reservoir .

If the No 1 window inner pane cracks and arcing begins. The maximum cabin differential pressure must be reduced to 5 PSID . Initiate drill for emergency descent . No pressurisation adjustments are necessary.

The right No.2 window heating input is controlled by. The right SIDE window heat control system. The right No.1 window heat control system . The left No.1 window heat control system .

With any control cabin window heating inoperative, speed should be restricted to. 250 Kts at altitudes below 15,000ft . 280 Kts at altitudes below 15,000ft . 250 Kts at altitudes below 10,000ft.

If a No.1 window outer pane cracks. Turn window heat off, limit IAS to 250kts below 10,000 ft. Cabin differential pressure must be reduced to a maximum of 5 PSI . Cabin differential pressure must be reduced to a maximum of 2 PSI .

If a window overheat light illuminates this indicates that. The associated window has reached an overheat condition and it must be selected OFF manually to remove power before any damage takes place . The power has been automatically removed from the associated window system. The power has been automatically removed from the associated window system and will be re-introduced automatically when the window has cooled .

With all window heaters selected ON and OVHT selected on the window heat test switch. The ON lights will not be affected but observe a rise in on the AC ammeters . All OVERHEAT lights illuminate and the ON lights will extinguish immediately . All OVERHEAT lights illuminate and the ON lights may extinguish immediately, or remain illuminated for as long as 70 seconds.

The window heat switches must be. ON to make a PWR or OVHT test. OFF to make a PWR and OVHT test only . OFF when making an OVHT test only .

With all windscreen heating switches ON and the aircraft on the ground. All lights remain extinguished until the aircraft is airborne because of touchdown relays . If a green ON light extinguishes and an OVERHEAT light comes ON, the Master Caution light will also come ON. If a green ON light extinguishes and an OVERHEAT light comes ON, the Master Caution light will not come ON .

After the overheat lights have illuminated during OVHT test. The window heat switches must be momentarily positioned to RESET to reset the system . No action is required as the resetting of the system is automatic . The window heat switches must be momentarily positioned to OFF to reset the system.

The alternate anti-skid system has. Two anti-skid valves. One alternate brake metering valve connected to all alternate antiskid valves . Four anti-skid valves .

The alternate brake source selector valve.   Prevents hydraulic system B from powering the alternate brake system when hydraulic system A is operating .normally . Prevents hydraulic system A from powering the alternate brake system when hydraulic system B is operating normally.     Isolates accumulator pressure .

The landing gear is normally operated. By system B hydraulic power but has a manual extension facility as a back-up . By system A hydraulic power but has a manual extension facility as a back-up. By system B hydraulic power but has a manual extension facility and system A hydraulic power as back-ups .

  With any landing gear not locked down and the flaps greater than 15 degrees.  A steady horn will sound which cannot be silenced (reset) with the Horn Cut-out switch, but can be silenced if either thrust lever is in a high forward thrust position .   A steady horn will sound which can be silenced (reset) with the Horn Cut-out switch . A steady horn will sound which cannot be silenced (reset) with the Horn Cut-out switch.

With either or both thrust levers retarded to idle, the landing gear not down and locked and the flaps are up. Red warning lights will illuminate and a steady horn will sound . Red warning lights will illuminate. Red warning lights will illuminate and intermittent horn will sound .

Rudder pedal steering.     Deactivated whenever the gear is up and locked . Active at all times irrespective of the gear position, as it can only demand +/- 7 degrees of nose wheel movement . Deactivated as the nose gear strut extends after takeoff.

Rudder pedal steering.   Is available up to +/- 7 degrees on the captain's rudder pedals only . Is activated anytime the nose gear strut is extended . Can be overridden by the nose wheel steering wheel.

Hydraulic power for the normal brakes is supplied by. System B. System A . System A to outboard and system B to the inboard brakes .

The alternate brake system is powered by. System B hydraulics . System A hydraulics. Accumulator pressure .

  The 737 brake system has. Two hydraulic brake accumulators . No hydraulic brake accumulators . One hydraulic brake accumulator.

Before selecting the ILS test facility. Select an ILS or VOR frequency on the associated VHF NAV control panel . Select an ILS frequency on the both VHF NAV control panels . Select an ILS frequency on the associated VHF NAV control panel.

If the No. 1 ADF mode selector is selected to TEST, the No. 1 ADF needle on both RDMIs should. " Classics: Indicate 45 degrees left of the lubber line. NG's: Pointer slews to 135 degrees relative bearing. ".   All series: Rotate clockwise continuously . All series: Indicate 90 degrees right of the lubber line .

Selecting an IRS mode selector from OFF to ALIGN will initiate. The 10 minute alignment cycle. The 7 minute alignment cycle . The 30 second fast alignment cycle .

A flashing IRS ALIGN light indicates. Alignment cannot be completed due to IRS detection of airplane movement . Normal alignment cycle . Alignment cannot be completed due to IRS detection of airplane movement (on some 3/4/500's only), significant difference between previous and entered positions or no present position entry.

The SUPP NAV DATA pages are accessible. On the ground only. At any time . In flight only .

The IRS provides. Attitude and heading information while in ATT mode. Attitude, heading and wind information while in ATT mode . Attitude, heading and groundspeed information while in ATT mode .

The left IRS requires a full alignment so it is selected OFF. All electrical power is. Removed from the system after a 30 seconds shutdown cycle. Removed immediately . Removed from the system after a 10 seconds shutdown cycle .

The AC busses have failed in flight. The IRSs wil. Continue to operate on DC power for 5 minutes only . Continue to operate on DC power but the Right IRS is limited to 5 minutes only. Continue to operate on DC power but the LEFT IRS is limited to 5 minutes only .

The Weather radar WX/TURB mode displays detected turbulence within. 160 Nm . 40 Nm. 80 Nm .

Fast alignment of an IRS is possible. Any time the air/ground sensor is in ground mode . On the ground and stationary only. On the ground or in flight provided one IRS is fully aligned for cross-reference purposes .

The amount of fan air that is ducted through the pre-cooler is controlled by the. Thermostatic pre-cooler valve. Modulating and shut-off valve . Ram air controller .

The isolation valve is. DC operated . AC controlled and pneumatically operated . AC operated.

The A.P.U. bleed valve is. AC controlled and pneumatically operated . Pneumatically controlled and operated . DC controlled and pneumatically operated.

The source air ducted through the pre-cooler is. Fan air. 5th stage bleed air . Pressure controlled ram air .

Both hydraulic reservoirs are pressurised by. Separate engine bleeds directly to the reservoirs . 5th and 9th stage air from engine No. 2 only . Air from the pneumatic manifold.

The engine bleed valves are. Pneumatically activated and operated . AC activated and pneumatically operated . DC activated and pneumatically operated.

The sources of engine bleed are. 5th and 9th stages of the compressor section. 5th and 13th stages of the fan section . 5th and 9th stages of the turbine section .

The pneumatic duct pressure gauge. Indicates pressure available for engine anti-icing . Is DC powered . Indicates pressure in left and right pneumatic ducts.

What initial action is required in the event of a WING BODY OVERHEAT light illuminating. Close the isolation valve and switch off the associated engine bleed. Close the isolation valve and switch off the associated air-conditioning pack . Switch off the associated engine bleed .

The CFM International CFM56-3-B1 is. A high bypass ratio turbo fan engine rated at 20,000 pounds of take-off thrust. A high bypass ratio turbo fan engine rated at 22,000 pounds of take-off thrust . A low bypass ratio turbo fan engine rated at 20,000 pounds of take-off thrust .

The CFM56-3 N1 low pressure turbine consists of. 9 stages . 4 stages. 3 stages .

The CFM56-3 N1 rotor section consists of. A single stage fan and a nine stage booster section . A single stage fan and a two stage booster section . A single stage fan and a three stage booster section.

The START VALVE OPEN light (amber) indicates. The engine starter valve is open irrespective of air being supplied to the air driven starter . The engine starter valve is open and air is being supplied to the air driven starter. The Engine Start switch is in GRD .

Illumination of the OIL FILTER BYPASS ligh. Illuminates the Master Caution ENG annunciator light . Indicates the Oil Filter is being bypassed . Indicates an impending bypass of the oil Scavenge Filter.

The engine Ignition System contains. One DC and one AC high energy system . Two high energy AC systems. Two high energy DC systems .

The PMC INOP light. Indicates the PMC is inoperative when engine speed is above 46% N2 or the PMC is selected OFF. Indicates the PMC is inoperative when engine speed is above 46% N2 or the PMC is selected OFF. Indicates the PMC is inoperative when engine speed is above 46% N2 only .

The Fuel Flow Transmitter is located. Inside the MEC . Between the First and Second Stage of the engine driven fuel pump . Downstream of the MEC Fuel Shutoff Valve.

The engine fuel system (not including fuel tank pumps) has. Two single stage engine driven fuel pumps . Two electrical fuel pumps . One engine driven fuel pump with two stages.

The engine fuel system contains. One fuel heater only . One fuel heater and one fuel/oil heat exchanger. One fuel/oil heat exchanger only .

With flap 15 selected and landing gear UP, the landing gear warning horn cannot be silenced with the HORN CUTOUT switch if. Either thrust lever is below 10 degrees or an engine not running and the other thrust lever is less than 30 degrees . Both thrust levers are below approx 30 degrees . Either thrust lever is below 20 degrees or an engine not running and the other thrust lever is less than 34 degrees.

Classics: The Take-off configuration warning is armed when on the ground and either or both Forward Thrust levers are advanced for takeoff. The Take-off warning horn sounds when. Stab trim is in the green band range, or the trailing edge flaps are in the Flaps 1 through 15 take-off range, or the leading edge SLATS are not in the correct position for take-off or the speed brake is NOT in the DOWN position or the Parking Brake is set . Stab trim is NOT in the green band range, or the trailing edge flaps are NOT in the Flaps 1 through 15 take-off range, or the leading edge FLAPS are NOT in the correct position for take-off or the speed brake is NOT in the DOWN position or the Parking Brake is set. Stab trim is NOT in the green band range, or the trailing edge flaps are NOT in the Flaps 1 through 15 take-off range, or the Parking Brake is NOT set .

The landing gear warning horn will sound if either or both thrust levers are approximately in the idle position and the flaps are in the. Up through 40 position provided the aircraft in not on the ground . Up position . 1 through 10 position.

The landing gear warning horn cannot be silenced by the Horn Cut-Out switch, regardless of Forward Thrust lever position, when the flaps are. In the 15 position. In the 15 position or greater than the 15 position . Greater than the 15 position.

The Mach/Airspeed Aural Warning will sound when the Mach number or IAS exceeds. 0.84M or 320 Knots whichever is reached first. 0.84M above 24,000 feet or 340 knots below 24,000 feet whichever is reached first . 0.82M or 340 Knots whichever is reached first.

The Proximity Switch Electronic Unit (PSEU) light illuminates. If an overwing exit flight lock fails to disengage when commanded at any time . is inhibited from thrust lever advance for take-off until 30 seconds after landing. If a fault is detected in the PSEU at any time .

EGPWS terrain display changes from dotted to solid yellow. 20 - 30 seconds from impact . 30 - 60 seconds from impact . 40 - 60 seconds from impact.

The weather radar automatically begins scanning for windshear when: In flight below 1200ft RA . In flight below 2500ft RA . In flight below 2300ft RA.

The maximum aircraft altitude for both APU bleed and electrical load is. 10,000 feet . 17,000 feet . 35,000 feet .

The battery voltage range is. 22 - 30 volts. 22 - 32 volts . 24 - 32 volts .

The minimum fuel required for ground operation of the hydraulic electric pumps is. 456 KGS in respective wing tank . 760 KGS in respective wing tank. 1150 KGS in respective wing tank .

The maximum external air pressure (ground cart) is. 30 PSI (at 232 degrees C) . 40 PSI (at 232 degrees C) . 60 PSI (at 232 degrees C ).

The minimum engine oil pressure is. 13 psi. 14 psi . 15 psi .

When engine bleeds are ON, both air conditioning packs must be set to. AUTO or OFF for approach and landing only . AUTO or OFF for take-off, approach or landing. AUTO or OFF for takeoff only .

The maximum fuel tank quantities (Classics) are. 4,600 kgs in each wing tank, 7,000 kgs in the centre tank . 4,828 kgs in each wing tank, 7449 kgs in the centre tank. 4,728 kgs in each wing tank, 7249 kgs in the centre tank .

The maximum lateral fuel imbalance between wing tanks 1 and 2 must be scheduled to be zero and, for all phases of flight, must not exceed. 453 kgs. 800 Kgs. 900kgs .

What is the minimum engine oil quantity (per engine) for dispatch for an aircraft with EIS. 66% Full . 75% Full. 79% Full .