Airplane General Quiz

Airplane General

The dimensions of a B737 MAX -8200 are 117' 10" wingtip to wingtip and 40' 10" tail height.
The minimum width of pavement for a 180° turn is 80 feet.
Do not attempt to turn away from an object within 23 feet of the nose or 15 feet of a wingtip.
The Maximum Landing Weight (MLW) of the B737 MAX -8200 is 152,800 lbs.
A LOCK FAIL (amber) light indicates a failed door lock or Flight Deck Access System (FDAS) switch in the OFF position.
The flight crew is warned of a correct emergency access code entry by a flight deck chime and AUTO UNLK (amber) light.
If the FLT DK DOOR Lock selector is placed momentarily to DENY, entry is denied and keypad entry is blocked for several minutes.
The flight deck door unlocks after a time delay if the correct emergency access code is entered and DENY is not selected.
The first officer's side window (R2) can be opened from the exterior of the plane for emergency access.
There are two overwing emergency exit doors and one mid-exit emergency door on each side of the B737 MAX -8200.
Illuminated overwing/mid-exit annunciators indicate if the related exit door is not closed and locked.
The aft attendant panel EMERGENCY EXIT LIGHTS switch overrides the flight deck switch for illuminating lights only.
The cabin ceiling and side wall lights at full intensity power the photoluminescent floor path marking system along the cabin aisle, provided strips are not covered or blocked.
Cockpit dome light is powered by Standby Electrical Power during emergency evacuations.
Crew oxygen is supplied by a single cylinder.
The crew oxygen bottle pressure is displayed on the aft overhead panel.
The maximum allowable pressure drop during a crew oxygen pressure drop test is 100 psi.
Crew oxygen requirements for dispatch with 2 pilots and 1 flight deck jumpseater are found in the QRH/Operational Information (OI).
100% oxygen at positive pressure is supplied to a flight crew oxygen mask above 27,000 feet or if selected on the mask.
Passenger oxygen is provided by a chemical oxygen generator located in each Passenger Service Unit (PSU).
Passenger service units (PSU) oxygen masks deploy automatically at 14,000 feet or manually when the PASS OXYGEN switch on the aft overhead panel is positioned ON.
The flight crew can confirm that PSU door solenoids have been powered by the PASS OXY ON light on the aft overhead panel.
Passenger oxygen flows for 12 minutes once initiated, and the oxygen cannot be shut off once initiated.
Protective Breathing Equipment (PBE/Smoke Hoods) provides useful oxygen for 20 minutes.
Potable water tank quantity is shown on the aft flight attendant station.
Potable water tanks are pressurized with bleed air from engines or APU, or by an electric air compressor.
Cargo compartments are pressurized during flight.
Passenger information signs automatically illuminate when the landing gear or flaps are extended.

Hydraulics

Hydraulic reservoirs are pressurized to ensure positive fluid flow to all hydraulic pumps.
Foaming in flight is indicated by blinking LOW PRESSURE lights for both pumps and a momentarily appearing MASTER CAUTION and HYD light.
Ailerons, rudder, elevators, and elevator feel, are shared components between Hydraulic Systems A and B.
Components of Hydraulic System A/B that are not recoverable when a system fails are landing gear retraction, ground spoilers, leading edge flaps & slats retraction and autobrakes.
Components powered by the Standby Hydraulic System are rudder, thrust reversers, and leading edge flaps and slats (extend only).
The Engine-Driven Pump (EDP) supplies approximately 6 times the fluid volume as the Electric Motor-Driven Pump (EMDP)
RF (refill) is displayed next to the hydraulic quantity when the airplane is grounded with both engines shut down or after landing with flaps up.
Hydraulic quantity is considered low when RF (refill) is displayed next to the A or B quantity during preliminary preflight.
Normal hydraulic system pressure is 3000 psi, and maximum hydraulic system pressure is 3500 psi.
The ELEC 2 OVERHEAT light illuminates if the system A Electric Motor-Driven Pump (EMDP) or the fluid used to cool and lubricate it overheats.
The minimum fuel level in each main wing tank is 1675 lbs to ensure adequate cooling for the hydraulic fluid for ground operation of the Electric Motor Driven Pumps (EDMP).
Pushback is not permitted without both system A hydraulic pumps turned off.
A Power Transfer Unit (PTU) supplies additional hydraulic fluid to operate Autoslats and Leading Edge Flaps and Slats at the normal rate when system B Engine-Driven Pump (EDP) is inoperative.
The PTU does not transfer hydraulic fluid, but uses system A pressure to power a hydraulic Motor-Driven Pump, which pressurizes system B hydraulic fluid.
The PTU operates when airborne and flaps are not up, when System B Engine-Driven Pump (EDP) pressure drops below the set limits.
The Landing Gear Transfer Unit (LGTU) supplies the volume of hydraulic fluid needed to raise the landing gear at normal speed when System A Engine-Driven Pump (EDP) volume is lost.
The LGTU operates when airborne, when Engine 1 RPM drops below a limit value, and when the landing gear lever is in the up position or the main landing gear is not up and locked.
The Standby Hydraulic System is typically used as a backup if System A and/or B pressure is lost.
System A or B pressure loss alone does not activate the Standby Hydraulic System
The Standby Hydraulic System is manually activated by moving either FLT CONTROL A or B switch to the STBY RUD position or moving the ALTERNATE FLAPS Master switch to the ARM position.
The Standby Hydraulic System automatically activates if airborne, wheel speed greater than 60 knots, flaps are extended, or if there is a loss of hydraulic system A or B and the FLT CONTROL switch is ON
If Hydraulic System A and/or B fail, the Standby Hydraulic System provides alternate operation for the affected Thrust Reverser.

Flight Controls

Hydraulic Systems A and B normally power the primary flight controls.
The Ailerons, elevators, and rudder are mechanically connected to cables in a system that provides the roll control from the spoilers (with a jammed aileron system).
Force applied to the First Officer's control wheel provides roll control from the spoilers, with a jammed spoiler system the Captain’s control wheel provides roll control from the ailerons.
Aileron trim is prohibited with the autopilot engaged due to autopilot overpowering the trim.
More than 10° (approximately) of control wheel displacement initiates Flight Spoiler deflection.
FEEL DIFF PRESS light indicates excessive differential pressure in the Elevator Feel Computer (caused by a hydraulic system loss).
If the Captain’s or First Officer’s control column is jammed apply force against the jam, whichever column moves, can provide adequate control
The Elevator Jam Landing Assist system, aids when a landing is made with a jammed elevator, using spoilers to control flight path.
The Elevator Jam Landing Assist system is activated when the Elevator Jam Landing Assist switch is ON, and the flap position is 1 or greater, and autopilot is disengaged.
Stabilizer trim is accomplished with either the autopilot system or manually via cables that allow the pilot to position the stabilizer by rotating the stabilizer trim wheels.
A cutout switch can disconnect inputs to the stabilizer trim motor from either the autopilot trim or main electric stabilizer trim system.
The Control Column Actuated Stabilizer Trim Cutout switches interrupt operations of the Main Electric Stabilizer Trim, Autopilot Trim and Speed Trim function when the control column movement opposes trim direction in either forward or aft trim modes.
FORWARD control column movement interrupts MCAS nose up commands.
AFT control column movement does not affect MCAS nose down trim commands.
The Flight Control Computer (FCC) software is a layer of redundancy that mitigates latent failure of the Control Column Actuated Stabilizer Trim Cutout, cutting out automatic stabilizer commands that extend beyond the cutout position of the Control Column Actuated Stabilizer Trim Cutout, and for the main electric stabilizer trim commands.
The STAB TRIM switch on the aft electronics panel, overrides the control column position when in the OVRD position, allowing operation of the Main Electric Stabilizer Trim regardless of the control column position.
An intermittent horn occurs if an attempt is made to takeoff with the stabilizers trim out of the green band.
The speed trim system (STS) provides both speed and pitch stability augmentation (with the autopilot engaged).
Speed trim function provides speed stability and MCAS provides pitch stability.
The speed trim function improves flight characteristics in low-weight, aft center of gravity, and high thrust conditions when the autopilot is not engaged.
Speed trim returns the airplane to a trimmed speed by commanding the stabilizer in the direction opposite of the speed change.

Flight Instruments and Displays

Static Air Temperature (SAT) is displayed on the CDU PROGRESS page.
If a single Display Processing Computer (DPC) fails, the system automatically selects another DPC to drive all four displays.
The DISPLAYS SOURCE switch on the DISPLAYS panel allows manual selection of which DPC drives the displays, overriding the automation.
A failed DPC 1 or 2 is indicated as amber on the Primary Flight Display (PFD).
The DISPLAYS CONTROL PANEL (amber) indicates a failed EFIS Control Panel.
When an EFIS control panel fails, displays can be controlled by the remaining panel through either BOTH ON 1 or BOTH ON 2.
Weather Radar is automatically set to OFF if an EFIS panel fails in-flight.
Outboard DU failures result in the automatic reconfiguration of displays, with the adjacent MFD replaced by a reduced size PFD
Inboard DU failures are reconfigured, and the Max Display System (MDS) ensures each pilot has a PFD and Engine Display is always present.
In cases of inboard DU failures and the display of the Engine Display, display select switches can be used to show the Engine Display on the desired outboard DU.
The Integrated Standby Flight Display (ISFD) is used during standby power operation.
The ISFD is powered by the Battery Bus.
Maximum Operating Speed (red/black) is displayed at the bottom of the amber bar.
The amber bar indicates maximum maneuver speed when flaps are up.
The amber bar indicates placard speed for the next normal flap setting sequence when flaps are not fully retracted.
V2+15 indication (white/sharktooth) is displayed for takeoff.
Minimum Maneuver Speed (amber) provides 1.3g capability to the stick shaker below 20,000 feet and above 20,000 feet for low-speed buffet.
The Minimum Speed (red/black) indication above the amber bar indicates the speed at which the stick shaker occurs.
Flight Path Vector (FPV) symbol represents the airplane's flight path angle vertically and drift angle laterally.
Pitch Limit Indication (PLI) is displayed in amber when flaps are not in the up position or at slow speeds with them up.
The PLI indicates the pitch limit for the stick shaker actuation.
Bank Pointer turns amber when bank angle is 35 degrees or more and Slip/Skid indication turns amber when bank angle is 35 degrees or more.
Roll/Yaw Assymetry alert illuminates when the autopilot requires more than 75% of its roll authority due to unusual asymmetric forces.
The Roll Command Arrow displays when bank angle exceeds 45 or 65 degrees, depending on pitch attitude.
Radio Altitude is displayed in the bottom center of the attitude indication area, and highlighted in white for 10 seconds if it falls below 2,500 ft AGL.
The rising runway indicator displays when a usable localizer signal is present and Radio Altitude is less than 2,500 ft.
Landing Altitude Reference Bar indicates height above the touchdown, with a white bar between 500-1000 ft above, and an amber bar between 0-500 ft, above the landing altitude.
Part Time Mini-map displays during Inboard Display failure.
The Mini-Map displays TCAS targets, terrain, weather, true airspeed, and ground speed/wind direction/wind speeds.
Failures/invalid information on the PFD/ND are indicated by flags or dashes.

Fuel

Normal fuel feed schedule is from the center tank until it is nearly empty, then pressure is fed from the respective main tanks.
The Nitrogen Generation System (NGS) converts bleed air to Nitrogen-Enriched Air (NEA) for center tank use to reduce flammability
The NGS is indicated by a dark right-hand main wheel well indicator panel
A tripped NGS is indicated by an amber INOPERATIVE light.
Center tank pumps automatically shut down after a short time if a low output pressure is detected.
Engines can often still suction feed fuel from center tank if both main AC fuel pumps fail.
Engines cannot suction feed fuel from the center tank.
A crossfeed valve allows fuel to be transferred from any tank to each engine.
The crossfeed valve is closed when no light (or dim light) is illuminated.
The crossfeed valve is in transit when a bright blue light is illuminated.
A bright blue crossfeed light indicates the valve is open and that the position of the valve and the CROSSFEED selector are inconsistent..
A dim blue crossfeed light indicates the valve is closed.
Spar and Engine Fuel Shutoff Valves are closed by pulling the respective engine fire switch.
Related engine or spar fuel shutoff valve is in transit or if the valve position and engine start lever or engine fire switch are inconsistent, an ENG VALVE CLOSED (dim blue) and SPAR VALVE CLOSED (dim blue) light is illuminated.
ENGINE VALVE CLOSED (blue) and SPAR VALVE CLOSED (blue) lights extinguish when engines are running.
The APU is supplied by the left side of the fuel manifold during operation of all AC fuel pumps
The APU can suction feed from Main Tank No. 1
Usable fuel quantity for each Main Tank is 8,500 lbs, and for the Center Tank is 28,600 lbs.
The CDU displays the Total Fuel Quantity on the PERF INIT page.
Fuel system checks for low pressure, fuel quantity issues, reserve fuel, and fuel disagreement between the totalizer and the FMC calculations by utilizing various annunciations on the Engine Display.

Electrical

The Hot Battery Bus is always connected to the battery.
The Switched Hot Battery Bus is powered when the battery switch is ON.
Battery voltage range is 22–30 volts.
BAT DISCHARGE (amber) indicates excessive battery discharge.
Integrated Drive Generators (IDGs) provide 115 Volt 400 cycle Alternating Current (AC) power.
TR UNIT (amber) light indicates a failed TR1 or TR2/TR3.
ELEC (amber) light indicates a fault in the DC power system or Standby Power system.
The ELEC (amber) light is inhibited while in flight.
Principles of 737 electrical operation are non-paralleling of AC sources and the source of power being connected to the related Transfer Bus disconnecting an existing source.
Illumination of the GRD POWER AVAILABLE light indicates ground power is connected and meets the aircraft's power quality standards.
APU GEN OFF BUS light indicates that the APU is running and not powering the Transfer Bus.
SOURCE OFF light illuminates until the other APU GEN switch is moved to ON.
AC Transfer Buses are powered by External Power or APU.
If both Transfer Buses are powered, the output from the associated generator is supplied to the related Bus while the other power source remains in operation.
If External Power or APU is powering both Transfer Buses, and engine generator power is applied to its onside Transfer Bus, the External power or APU continues to supply power to the other Transfer Bus.
The purpose of the Generator Drive (DRIVE) is to maintain power supply to the Integrated Drive Generator (IDG) if low oil pressure is detected, engine shutdown occurs, or high temperature is sensed.
Single generator operations cause the transfer of load between galleys and main buses on Transfer Bus 2 if an overload is sensed.
If there is still a demand overload, the galleys and main buses on Transfer Bus 1 will be shed; followed by IFE buses if the overload still exists.
If the APU is the only electrical source, galley and main buses are shed automatically, and if the load still exceeds design limits, both IFE Buses are shed automatically.
A ground service switch is used to manually control ground service buses to facilitate servicing of the aircraft with external power without activating Transfer Buses.
If generator 1 fails inflight, the AC Transfer Bus 1 also loses power.
The Battery Charger restores and maintains battery power.

Fire Protection

Engine, APU, and Lavatory systems have fire detection and protection capabilities.
The main wheel well has its own fire detection system.
The cargo compartments are protected by a smoke detection and suppression system.
Each engine has two overheat/fire loops .
A single overheat/fire detector loop malfunction does not result in a loss of fire detection.
A MASTER CAUTION light illuminates if the engine has overheat problems, and a fire warning bell sounds when an engine fire occurs
A single engine overheat can activate both MASTER CAUTION lights and the OVHT/DET system annunciator light, also the related Overheat light.
Engine fires will trigger both MASTER/CAUTION lights, both master FIRE WARN lights, the OVHT/DET system annunciator light, and the related engine overheat light, as well as illuminate lights for the related engine fire switch, and related engine start lever.
There are two extinguisher bottles for engine fires; these may discharge when a fire is detected.
Both bottles can suppress fire for one engine.
The overhead system panels with gray backgrounds are affected by the pulling of a system fire switch
Pulling the engine fire switch will close the engine fuel shutoff valve, spar fuel shutoff value, engine bleed air valve, and related hydraulic shutoff valve.
The L or R BOTTLE DISCHARGE light indicates if a bottle has discharged or if pressure is low.
The APU has a single fire detection loop.
A failure in the APU loop will result in a loss of APU fire detection capability.
A fire in the APU will trigger a fire warning bell, both master fire warning lights, the APU fire annunciator light, and an APU fire warning horn in ground operation along with an emergency light that flashes.
There is one fire extinguisher bottle for APU fire events.
The APU ground control panel is located right in the main wheel well.
Pulling the APU fire switch will close the APU air inlet door and shut down the fuel solenoid along with the APU fuel shutoff valve.
The APU generator control relay and breaker will also trip if the APU fire switch is pulled.
The nose wheel well does not have fire detection.

Air Systems

The air conditioning, pressurization, wing and engine thermal anti-icing, hydraulic reservoirs, engine starting and water tank pressurization all use bleed air.
Excessive engine bleed air temperature, overpressure or underpressure cause the BLEED light to illuminate; this also occurs when the isolation valve fails to operate correctly after takeoff or go-around (45 seconds after flaps up)
To remedy this, the engine thrust should be limited to idle when the DUAL BLEED light is illuminated.
The DUAL BLEED light illuminates when the APU bleed valve is open and either Engine 1 or Engine 2 bleed air switch is ON, with the ISOLATION VALVE open.
Single pack operation can maintain pressurization at the maximum certified altitude.
Both packs cannot operate with APU bleed air inflight or when on the ground, only one pack at a time from a single engine.
If the pack trip off or either of the primary and standby pack controls fail then the PACK light illuminates.
A failure of the flow control valve will cause both pack lights to illuminate after takeoff, if both pack switches were off 45 seconds after the flaps were made.
Pack output air temperature is determined by the zone that needs cooling the most.
ZONE TEMP lights illuminate when the duct temperature overheats or if the flight deck temperature control fails.
Recirculating fans reduce the pack load and engine bleed air demand.
Smoke detection in the equipment cooling system will cause the E/E cooling supply fan (s) to turn off for around 5 minutes, followed by the activation of overboard exhaust valve and the shutoff of the recirculation fans.
The OFF light of the SUPPLY/EXHAUST will not illuminate in this event.

Anti-Ice and Rain

Windshield wipers cause scratching of the dry windshield.
Windows 1 and 2 on each side of the flight deck are heated.
Pulling the captain’s or first officer’s WINDSHEILD AIR control activates an alternate system for defrosting window number 1 and 2.
Power to the windows is automatically removed by a panel light if an overheat condition is detected.
A window heat ON light that turns off indicates either a switch is OFF, an overheat condition exists a system failure, or the system is at the correct temperature.
The PWR TEST selection on the probe heat panel activates an automatic supply of power to both probe heat systems when the engines are running.
On electrical standby, only the captain's pitot probe is heated.
ENG ANTI-ICE switch activation turns on engine anti-ice with the engine start switch in the CONT position which turns the Anti-ice valves on.
Minimum Maneuver Speeds and Speedshaker logic are adjusted when the ENG ANTI-ICE switch is ON, and all return to normal without Wing Anti-Ice usage in flight.
Wing Anti-Ice is required when icing conditions exist or are anticipated on the ground, between engine start and takeoff.

Landing Gear

Landing gear extension/retraction is supplied by Hydraulic System A.
Landing gear lever lock prevents the gear from being moved to the up position on the ground.
In flight, the air/ground system automatically opens the lever lock.
Main gears are held in place using mechanical uplocks
Nose gear is held in place using an overcenter lock and enclosed by doors mechanically linked to the gear
Hydraulic pressure is removed from the landing gear system 10 seconds after all gears are up and locked.
A main gear tire with loose tread is prevented from causing damage by having it stop retracting and free falling back to the down position, however that fitting has to be replaced
Landing gear is extended by Hydraulic System A pressure releasing the uplocks using hydraulic pressure and air loads
The landing gear indicator lights on the aft overhead panel are different and run through a separate circuit from those on the main panel.
Both sets of landing gear indicator lights use two lights, per gear, to confirm the down and locked position.
Manual landing gear extension is possible even if the landing gear lever is in any position, with hydraulic system A pressure, however gear retraction is disabled.
Normal nose wheel steering is available when the nose gear is down and compressed by the weight of the plane.
Hydraulic systems A and B are used to provide pressure for normal and alternate nose wheel steering, respectively.

Communications

The radio associated with the panel being tuned is indicated by the Offside Tuning Light (amber).
Pressing the VHF TEST switch removes automatic squelch functionality to improve reception of weak signals.
A PANEL FAIL (amber) light indicates a failed Radio Tuning Panel.
A FAIL FAIL (amber) light indicates an individual radio failure.
The MASK-BOOM switch on the audio control panel (ACP) activates the oxygen mask microphone for transmissions.
The ALT-NORM switch on the ACP allows operation of the panel in either normal or degraded mode
A degraded radio operates with only Captain VHF 1.
Flight Deck microphones are keyed by the Glareshield, control wheel, or Audio Control Panel (ACP), R/T and I/C switches.
The INT position on the control wheel push-to-talk (PTT) switch selects the oxygen mask or boom microphone and disconnects the ACP transmitter, until it is changed to the OFF or MIC position.
The flight interphone system allows for communication between the flight deck, flight attendants, and ground personnel.

Warning Systems

RED lights indicate warnings that require immediate attention and AMBER lights warnings that require timely attention.
BLUE lights are used for informational purposes only and GREEN lights indicate a fully deployed or extended operational condition.
A MASTER FIRE WARN light illuminates when a fire is detected in the engine, APU or main wheel well, or the cargo hold.
Pressing the illuminated MASTER FIRE WARN switch, will both extinguish the FIRE WARN lights and silence the fire warning bell, also resetting the system for further cautions.

Global Positioning System (GPS) amber lights indicate failure in both or either GPS sensor units (system recall reset light extinguishes).
Instrument Landing System (ILS) amber lights indicate failure in both or either ILS sensor units (system recall reset light extinguishes).
The duration of alignment processes is 5-17 minutes and the airplane is required to remain stationary
Automatic Navigation Realignment (ANR) occurs when a fast realignment is initiated or the airplane remains stationary in NAV mode for 7.5 to 15 minutes after the flight.
The ADIRUs are reset to remove drift errors accumulated during the flight.
Left and right IRS are typically powered by AC Standby Bus and Transfer Bus 2 respectively.
If AC power is lost, IRS switches to backup DC power from the switched hot battery bus, and an ON DC light illuminates.
Backup DC power to the right IRS is lost after 5 minutes if AC power is not restored.
Altitude intervention can be accomplished using the MCP, while on climbs and descents by using the altitude intervention switches to delete altitude restrictions.
Descent below airspeed restriction altitude will not occur in VNAV mode, unless airspeed is at or below the restricted value plus 10 knots
The transition from VNAV PHT to LVL CHG during descent in VNAV mode is triggered when actual speed decreases to minimum speed as it relates to the given altitude.
VNAV remains in VNAV PATH if preparing to execute a missed approach when the MCP altitude is at least 300 feet above current airplane altitude or if the airplane is further than 200 feet below the vertical path, VNAV commands zero vertical speed.

Automatic Flight

The Autopilot Flight Director System (AFDS) consists of 2 flight control computers and automatic modes.
Illuminated switches on the Mode Control Panel (MCP) indicate the selected mode.
All autopilot modes are disengaged by selecting a different command or by disengaging the autopilot (A/P) and turning the Flight Directors (F/Ds) off.
Only one Autopilot (A/P) can be engaged at a time, except during approach mode.
The A/P may disengage if there is an inappropriate engagement, a column or wheel force override, an engage switch, or the TAKEOFF/GO-AROUND switch.
Autopilot is recommended to be the one associated with the operating IRU if there is an inoperable IRU.
Radio Altimeter inoperability results in approximate 2 second autopilot disengagement after LOC and GS capture.
Independent Flight Director (F/D) operation is indicated on the Mode Control Panel (MCP) by illumination of the respective Master (MA) lights.
F/D pitch or roll guidance bars will be removed (with approximately 4 seconds of disagreement) and return to normal operation when guidance is in agreement, a pitch or roll mode change occurs or automatically operates.
The CWS P and CWS R status annunciations show pitch and roll mode engagement when present in the AFDS status display.
When the SINGLE CH is present in ILS approach, the status annunciation will appear when localizer capture is achieved and will disappear when there is a pitch monitor confidence test.