QDB Flashcards
Witch statement is correct?
a) P6 and P18 CB panels are located both behind the co-pilot seat.
b) P6 CB panel are located behind the captain’s seat and P18 CB panel are located on the
left hand side of the observer seat.
c) P18 CB – panels are located behind the captain seat, P6-CB panel are located behind the
Co-pilot seat.
d) There are only P6 CB- panel on the flight deck. P18 CB panels are located in the E&E
compartment
P18 CB – panels are located behind the captain seat, P6-CB panel are located behind the Co-pilot seat.
VOL-2 – 1.20.2
How is the crew oxygen system pressure checked?
a) By direct reading of the gauge, viewed on walk-around inspection.
b) By checking that the PASS OXY ON light is illuminated.
c) By reading the Servo pneumatic gauge on the right forward panel.
d) By reading the Electrical gauge on the aft overhead panel.
By reading the Electrical gauge on the aft overhead panel.
Oxygen gage is electrical: BAT BUS, with Bat switch ON.
The bottle is located in the aft EE compartment with an access door in the forward cargo compartment.
Breakable green plastic discharge indication dick on the fuselage skin held in place by snap-ring, shows
cylinder discharge from overpressure. Is flush-mounted to the fuselage skin just aft of the electronic equipment compartment external access door.
If the PASS OXY ON amber light illuminates, what does this indicate?
a) Passenger oxygen system pressure is low.
b) Oxygen shutoff valve is ON.
c) Passenger oxygen system is activated.
d) Power loss to passenger oxygen activation system.
Passenger oxygen activation system is activated.
Passenger system activated manually or automatically to drop masks.
MASTER CAUTION and OVERHEAR annunciator illuminate
When Fasten seat belt switch in’’AUTO’’ position….
FASTEN SEAT BELT signs will illuminate when flaps or gear are extended.
Extinguish when flaps or gear are retracted.
The LOGO light are located ….
a) On top of the horizontal stabilizer
Cabin Oxygen system is activated at a cabin altitude of….
a) 14.000 Ft
What will cause the AUTO FAIL light to illuminate?
a) Loss of Generator number 1.
b) Operational controller fault or cabin altitude above 14.000 feet.
c) Excessive rate of cabin pressure change (+/- 2.000 sea level ft. min)
d) Selecting alternate system.
Excessive rate of cabin pressure change (+/- 2.000 sea level ft. min)
AUTO FAIL light illuminates if:
o Loss of DC power
o Controller fault
o Outflow valve control fault
o Excessive differential pressure (>8.75 psi)
o Excessive rate of cabin pressure change (+/-2.000 sea level feet/minute) o High cabin altitude (above 15.800 feet)
- With illumination of the AUTO FAIL light, the pressure control automatically transfer to the other auto controller (ALTN mode).
Moving the pressurization mode selection to -ALTN position extinguishes the AUTO FAIL light, however the ALTN light remain illuminated to indicate single channel operation.
When does the cabin start to pressurize?
a) On the ground at high power setting.
b) At main wheel lift off.
c) At nose wheel lift off.
d) When the cabin and cargo door close.
On the ground at high power setting.
Takeoff phase
o Both engine N1’s is greater than 50%
o Both engine N2’s is greater than 84%
o Outflow valve modulate close and pressurize the airplane to approximately 0.1 psid below field
elevation. This prevents the uncomfortable pressure bump at the airplane rotation.
Climb phase
o Starts when the air/ground system indicates that the left and right landing gear are in the air.
Both pack switches are in the ‘’AUTO’’ position. The aircraft is in flight with flaps down, and one pack is inoperative. Why will the other pack not go in to the ‘HI FLOW’’ mode?
a) Because the pack cannot accept bleed air from both engines.
b) When one pack is inoperative, the opposite pack is locked in to the ‘’normal’’ flow mode.
c) The opposite pack automatically goes in to the ‘’HI FLOW’’ mode.
d) To ensure that there will be adequate thrust in the event of a single engine situation.
To ensure that there will be adequate thrust in the event of a single engine situation.
AirflowControl - OA-MPartBvol.2–2.31.1
With both air conditioning pack switches in AUTO and both packs operating, the packs provide “normal air flow”. However, with one pack not operating, the other pack automatically switches to “high air flow” in order to maintain the necessary ventilation rate. This automatic switching is inhibited when the airplane is on the ground, or inflight with the flaps extended, to insure adequate engine power for single engine operation. Automatic switching to “high air flow” occurs if both engine bleed air switches are OFF and the APU bleed air switch is ON, regardless of flap position, air/ground status or number of packs operating.
What are the restrictions in the event a radio altimeter is inoperative?
a) Do not use the autopilot system.
b) Do not use the associated FCC for landing. You may use it for approach.
c) Do not use the associated autopilot for approach.
d) There are no restrictions.
Do not use the associated autopilot for approach
OPERATIONS (O) MEL - 34-20 RADIO ALTIMETER SYSTEMS
Note: For aeroplanes with -1, -2, or -3 SMYD, an invalid signal from radio altimeter number 1 will result in failure of both stick shakers to self-test.
1. Ensure that weather minimums or operating procedures are not dependent upon its use.
2. With radio altimeter(s) inoperative, do not use the associated autopilot, flight director or autothrottle for approach and landing.
3. For aeroplanes with FCC Operational Program Software (OPS) 2212-HNP-03B-05 or later installed, if the remaining radio altimeter fails:
- How are the N1 limits and target N1 values normally provided to the A/T?
a) By the A/T computer.
b) By the AFDS.
c) By the Flight Management Computer (FMC).
d) By the Manual Selection.
By the Flight Management Computer (FMC)
Thrust mode display
Display green - active N1 limit reference mode
With N1 Set outer knob (on engine display control panel) in AUTO, active N1 limit is displayed by
reference N1 bugs
With N1 Set outer Knob (on engine display control panel) in either 1, 2 or BOTH (other than AUTO), the
thrust mode display annunciation is MAN
Active N1 limit is normally calculated by FMC Autothrottle Limit (A/T LIM) indication
Illuminated (white) – the FMC is not providing the A/T system with N1 limit values. TheA/T is using a degraded N1 thrust limit from the related EEC
Which of the following occurs when a TOGA switch is pressed for a flight director go- around from a single A/P ILS approach?
a) Autopilot disengage / TOGA mode of the flight director engage and the A/T advances thrust levers to reduced go-around N1.
b) TOGA mode of the Autopilot / Flight director engage and A/T advance thrust levers to reduced go-around N1
c) Autopilot disengage, flight director bars retract and A/T advance thrust levers to reduced go-around N1.
d) Autopilot disengages / TOGA mode of the flight director engages and A/T disengages.
Autopilot disengage / TOGA mode of the flight director engage and the A/T advances thrust levers to reduced go-around N1.
F/D Go–Around - OA-M Part B vol. 2 – 4.20.19
If both A/Ps are not engaged, a manual F/D only go–around is available under the following conditions: • inflight below 2000 feet RA
• inflight above 2000 feet RA with flaps not up or G/S captured
• not in takeoff mode.
With the first push of either TO/GA switch:
• A/T (if armed) engages in GA and advances thrust toward the reduced go–around N1 to produce 1000 to 2000 fpm rate of climb. The A/T Engaged Mode annunciation on the FMA indicates GA
• autopilot (if engaged) disengages
• pitch mode engages in TO/GA and the Pitch Engaged Mode annunciation on the FMA indicates TO/GA
• F/D pitch commands 15 degrees nose up until reaching programmed rate of climb. F/D pitch then commands target airspeed for each flap setting based on maximum takeoff weight calculations
• F/D roll commands approach ground track at time of engagement. The Roll Engaged Mode annunciation on the FMA is blank
A/P Go–Around - OA-M Part B vol. 2 – 4.20.18
The A/P GA mode requires dual A/P operation and is available after FLARE armed is annunciated and prior to the A/P sensing touchdown.
With the first push of either TO/GA switch:
• A/T (if armed) engages in GA and the A/T Engaged Mode annunciation on the FMA indicates GA
• thrust advances toward the reduced go–around N1 to produce 1000 to 2000 fpm rate of climb
• pitch mode engages in TO/GA and the Pitch Engaged Mode annunciation on the FMA indicates TO/GA
• F/D pitch commands 15 degrees nose up until reaching programmed rate of climb. F/D pitch then commands target airspeed for each flap setting based on maximum takeoff weight calculations
• F/D roll commands hold current ground track. The Roll Engaged Mode annunciation on the FMA is blank
• the IAS/Mach display blanks
• the command airspeed cursor automatically moves to a target airspeed for the existing flap position based on maximum takeoff weight calculations.
• If the TO/GA switch is pressed after touchdown and prior to A/T disengagement, A/P channel disengages and the A/T may command GA thrust.
With the second push of either TO/GA switch after A/T reaches reduced go–around thrust:
• the A/T advances to the full go–around N1 limit. TO/GA mode termination from A/P go–around:
• below 400 feet RA, the AFDS remains in the go–around mode unless both A/Ps and F/Ds are disengaged
• above 400 feet RA, select a different pitch or roll mode.
• if the roll mode is changed first:
• the selected mode engages in single A/P roll operation and is controlled by the A/P which was first in CMD
• pitch remains in dual A/P control in TO/GA mode.
• if the pitch mode is changed first:
• the selected mode engages in single A/P pitch operation and is controlled by the A/P which was first in CMD
• the second A/P disengages
• the roll mode engages in CWS R.
• the A/T GA mode is terminated when:
• another pitch mode is selected
• ALT ACQ annunciates engaged.
Note: The pitch mode cannot be changed from TO/GA until sufficient nose–down trim has been input to allow single channel A/P operation. This nose–down trim is automatically added by the A/P to reset the trim input made by the A/P at 400 feet RA and at 50 feet RA during the approach.
With pitch mode engaged in TO/GA, ALT ACQ engages when approaching the selected altitude and ALT HOLD engages at the selected altitude if the stabilizer position is satisfactory for single A/P operation.
• if stabilizer trim position is not satisfactory for single A/P operation:
• ALT ACQ is inhibited
• A/P disengage lights illuminate steady red • pitch remains in TO/GA.
Note: To extinguish A/P disengage lights, disengage A/Ps or select higher altitude on MCP
If the autopilot ALT HOLD mode is overridden with control column pressure, which of the following occurs?
a) A/P disengage
b) LNAV disengage
c) A/P revert to CWS pitch
d) A/P revert to LEVEL CHANGE mode
A/P revert to CWS pitch
Pitch CWS with a CMD Engage Switch Selected - OA-M Part B vol. 2 – 4.20.7
The pitch axis engages in CWS while the roll axis is in CMD when:
• a command pitch mode has not been selected or was deselected
• A/P pitch has been manually overridden with control column force. The force required for override is greater than normal CWS control column force. This manual pitch override is inhibited in the APP mode with both A/Ps engaged. CWS P is annunciated on the FMAs while this mode is engaged.
Command pitch modes can then be selected.
When approaching a selected altitude in CWS P with a CMD engage switch selected, CWS P changes to ALT ACQ. When at the selected altitude, ALT HOLD engages.
If pitch is manually overridden while in ALT HOLD at the selected altitude, ALT HOLD changes to CWS P. If control force is released within 250 feet of the selected altitude, CWS P changes to ALT ACQ, the airplane returns to the selected altitude, and ALT HOLD engages. If the elevator force is held until more than 250 feet from the selected altitude, pitch remains in CWS P.
Which flight modes are annunciated when the autopilot is initially engaged in the COM mode and both F/D’s are OFF?
a) HDG SEL, MCP SPD.
b) HDG SEL, CWS PITCH
c) CWS ROLL, CWS PITCH
d) CWS ROLL, V/S.
CWS ROLL, CWS PITCH
- Is it possible to turn-off the FD Take-off mode, when A/C is below 400 feet RA?
a) Not possible.
b) Possible by disengaging the Auto pilot.
c) Possible by turning off both FD’s on the MCP
d) Only possible by pulling a flight computer circuit braker.
Possible by turning off both FD’s on the MCP
If not in a VNAV mode when does the A/T speed mode engage automatically?
a) When localizer is captured
b) When ALT ACQ engages
c) When F/D switched to ON
d) None of the above
When ALT ACQ engages
- Which of these AFDS mode allow both autopilots to be engaged at the same time?
a) VNAV
b) VOR/LOC
c) APP
d) LNAV
APP
- During a single engine F/D go-around with a push either TOGA switch:
a) F/D roll commands hold current heading
b) F/D roll commands hold current ground track
c) F/D roll commands hols current heading until passing 400ft.
d) F/D roll commands hold current ground track until passing 400ft.
F/D roll commands hold current heading
Single Engine F/D Go–Around - OA-M Part B vol. 2 – 4.20.20
With a push of either TO/GA switch:
• F/D roll commands hold current ground track. The Roll Engaged Mode annunciation on the FMA is blank
• pitch mode engages in TO/GA and the Pitch Engaged Mode annunciation on the FMA indicates TO/GA
• the F/D target speed is displayed on IAS/Mach display
• the F/D target speed is displayed on the airspeed cursor
• F/D pitch commands 13 degrees nose up. As climb rate increases, F/D pitch commands maintain a target speed.
• if engine failure occurs prior to go–around engagement, then F/D target speed is the selected MCP speed.
• if engine failure occurs after go–around engagement, then F/D target speed depends on whether ten seconds have elapsed since go–around engagement:
• if prior to ten seconds, the MCP selected approach speed becomes target speed
• if after ten seconds and the airspeed at engine failure is within five knots of the go–around engagement speed, the airspeed that existed at go–around engagement becomes target speed
• if after ten seconds and the airspeed at engine failure is more than five knots above go–around engagement speed, then the current airspeed becomes target speed.
Note: The target speed is never less than V2 speed based on flap position
unless in windshear conditions.
- During climbs and descent in LEVEL CHANGE, the airspeed is commanded by which subsystem of the AFDS?
a) FMC
b) Airspeed indicator
c) MCP
d) Auto throttle
MCP
- When VNAV mode is engaged, which system provide command to AFDS pitch and A/T mode?
a) MCP
b) FMC
c) A/T Computer
d) Autopilots flight directors computer
FMC
- You are established on the ILS 3000ft and inadvertently press TOGA once. What will happen?
C.NOTHING WILL HAPPEN AS THE AIRCRAFT HAS NOT DESCENDED BELOW 2000FT (?????? NOT SURE FOR THE ANSWER)
- What is required to erase the cockpit voice recorder tape?
a) Aircraft must be on the ground, parking brakes must be set and press the ERASE bottom for 2 seconds
b) Aircraft must be on the ground, parking brakes must be set and press the ERASE bottom for 1 second
c) Aircraft must be on ground, Ground Power connected and press the ERASE bottom for 2 seconds
d) Aircraft must be on ground, Ground Power connected and press the ERASE bottom for 1 second
Aircraft must be on the ground, parking brakes must be set and press the ERASE bottom for 2 seconds
ERASE Switch (red) - OA-M Part B vol. 2 – 5.10.11
Push (2 seconds) –
• all four channels are erased
• operative only when airplane is on ground and parking brake is set.
- When the Captain is transmitting on VHF – 1:
a) Only VHF-2 Reception is blocked
b) Only VHF-2 Transmission is blocked
c) Both VHF-2 Transmission and Reception are blocked
d) The First Officer can simultaneously transmit on VHF-2 and can receive VHF-2 on his
headset while the Captain is transmitting.
The First Officer can simultaneously transmit on VHF-2 and can receive VHF-2 on his headset while the Captain is transmitting.
- Which of the following statement is true regarding the Cockpit Voice Recorder?
a) Record audio from the Flight Deck area conversation.
b) Is erase automatically if recording are older than 60 minutes.
c) Records audio from the Captain’s, First Officer’s and Observer’s Audio Selector Panel and
Flight Deck area conversation.
d) Records audio from the Captain’s, First Officer’s audio selector panels and Flight Deck
area conversation.
Records audio from the Captain’s, First Officer’s and Observer’s Audio Selector Panel and Flight Deck area conversation.
Cockpit Voice Recorder - OA-M Part B vol. 2 – 5.20.6
The cockpit voice recorder uses four independent channels to record flight deck audio for 120 minutes. Recordings older than 120 minutes are automatically erased. One channel records flight deck area conversations using the area microphone. The other channels record individual ACP output (headset) audio and transmissions for the pilots and observer.
- 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.
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.
- Which communication radio system can be operated from standby electrical power
a) VHF-1
b) VHF-2
c) Both VHF 1- and VHF-2
d) VHF-1 and HF-1
VHF-1
- Which of the statements is correct about the switch selection on ASP ( Audio Selector Panel)?
a) VHF – 1 is received only when the VHF-1 receiver switch is selected (pressed down)
b) The VHF-2 and Cabin/Service interphone receiver switches may be selected at the same
time.
c) VHF-2 is received only when the VHF-2 transmitter switch is selected.
d) Only one receiver switch can be selected at a time.
The VHF-2 and Cabin/Service interphone receiver switches may be selected at the same time.
- What is the normal source of power for TR3
a) Transfer bus 1
b) Transfer bus 2
c) AC standby bus
d) Main bus 2
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.
- What power the main battery charger?
a) DC bus 1.
b) DC bus 2.
c) AC ground service bus 2.
d) AC main bus 1.
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.
- The electrical system incorporates an automatic load shedding feature. What is the first bus that is shed?
a) Galleys on transfer bus 1 and shed first.
b) Galley on transfer bus 2 are shed first.
c) The AC ground service bus is shed first.
d) The AC standby bus is shed first
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.
- Either generator or the APU can power both transfer buses. In the event a power source fails, what is required for that transfer bus to be powered by the opposite transfer bus power source?
a) The generator switch must be OFF.
b) The battery switch must be ON.
c) The BUS TRANS switch must be in the AUTO position.
d) Transfer takes place as long as the AC electrical system is powered.
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.
- What will happen if one engine driven generator fails during cruise?
a) Only the associated GEN OFF BUS light and TRANSFER BUS OFF light illuminate.
b) Only the associated GEN OFF BUS light, SOURCE OFF light and TRANSFER BUS OFF light
illuminate.
c) The GEN OFF BUS light, SOURCE OFF light and TRANSFER BUS OFF light illuminate. Also
various other lights associated with systems which were powered by the transfer bus
illuminate.
d) Only the associated GEN OFF BUS light and SOURCE OFF light illuminate.
Only the associated GEN OFF BUS light and SOURCE OFF light illuminate
- What is the significance of an illuminated TR UNIT light while in flight?
a) Any one TRU has failed
b) Only TR1 has failed
c) Only TR3 has failed
d) TR 1or TR2 and TR3 has failed
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
- Which buses are isolated from each other when the CROSS BUS TIE RELAY opens?
a) Isolate transfer bus 1 and 2
b) Isolate DC bus 1 from DC bus 2
c) Disconnect TR1 and TR3
d) Isolate DC bus 1 from standby DC bus
Isolate DC bus 1 from DC bus 2
OA-M Part B vol. 2 – 6.20.7
- The standby power switch in AUTO position. When will the automatic transfer of standby power to alternate source occur?
a) AC transfer bus 2 or DC bus 1 loses power
b) AC transfer bus 1 or DC bus 2 loses power
c) AC transfer bus 2 or DC bus 2 loses power
d) AC transfer bus 1 or DC bus 1 loses power
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.
- Which statement about illumination of the GRD POWER AVAILABLE light is correct?
a) Ground Power has been plugged in and automatically powers both ground services busses.
b) Ground Power has been plugged in and meets airplane power quality standards.
c) Ground Power has been plugged in, however airplane power quality is not measured.
d) Ground Power has been plugged in, and automatically powers both transfer busses.
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.
- What happens if the crew inadvertently take off with the APU powering both TRANSFER BUSSES?
a) During climb the GALLEY BUSSES may become inoperative.
b) Both MAIN BUSSES may become inoperative above 400ft. RA or after 20 seconds from
lift-off.
c) Both IDG’s will come on line automatically if the APU fails or is shut down.
d) A and C are both correct answers.
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.
- What will cause the DRIVE light for the respective IDG to illuminate?
a) Low oil pressure in the IDG
b) Low Generator frequency
c) Low Generator voltage
d) High Generator voltage
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.
- At what speed does starter cutout occur?
a) At 56% N1
b) At 65% N1
c) At 56% N2
d) At 65% N2
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.
- When can the thrust reversers be deployed?
a) Either radio altimeter sense below 10 feet.
b) Both radio altimeter sense below 10 feet.
c) Either radio altimeter sense below 25 feet.
d) Both radio altimeter sense below 25 feet.
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.
- If a wet start is detected, when will the EEC automatically turn off ignition and shutoff fuel to the engine?
a) 10 seconds after the start lever is moved to idle during ground starts.
b) 15 seconds after the start lever is moved to idle during ground starts.
c) 30 seconds after the start switch is moved to GRD during ground starts.
d) 20 seconds after the start switch is moved to GRD during ground starts.
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.
- If a crossbleed start is required during inflight starting, where will a ‘’X-BLD’’ indication be displayed?
a) On the ENG OUT page.
b) Above the N2 dial on the CDS.
c) On the ‘’SYS’’ page of the CDS.
d) Above the N1 dial.
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.
- When does the EEC monitor engine parameters to detect impending hot starts, EGT start limit exceedance, and wet starts?
a) Ground starts only.
b) In-flight starts only.
c) Ground starts and in-flight starts.
d) Ground starts and in-flight if cross bleed is used.
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.
- Which engine parameter does the EEC use to control thrust in normal mode?
a) N1
b) N2
c) EGT
d) Fuel Flow
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.
- What is the primary source of power for the thrust reverser system?
a) The pneumatic system
b) The standby hydraulic system
c) Either electrical system
d) Hydraulic system A for the no. 1 thrust reverser, hydraulic system B for the no. 2 thrust
reverser.
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.
- What is the power source for the right ignition system?
a) DC bus2
b) AC transfer bus 2
c) AC standby bus
d) The battery
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.
- What are the conditions that will automatically close both the Spar Fuel Shutoff Valve and the Engine Fuel Shutoff Valve?
a) Respective Start Lever is placed to CUTOFF or Respective Engine Fire Handle is pulling.
b) ECC detects contaminated fuel.
c) Engine Flame out.
d) Engine start switch is place to off.
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.
- Which statement is true regarding engine protection provided by EEC in flight?
a) The EEC provides redline exceedance protection for N1, N2 and EGT.
b) The EEC provides flameout protection and red line exceedance protection for N1, N2 and
EGT.
c) The EEC provides flameout protection and red line protection for N1 and N2. The EEC
does not provide EGT redline exceedance protection.
d) The EEC does not provide any protection it optimizes the fuel consumption.
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.
- What happen if you move the APU switch to OFF without operating it in the ‘’NO LOAD’’ mode for the recommended period?
a) The APU will shutdown immediately.
b) The APU will operate in the no load mode for 1 minute, then shutdown.
c) A FAULT warning will occur.
d) APU will not shutdown.
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.
- What is the electrical power source for the APU starter motor?
a) Always the battery.
b) The battery or AC power from transfer bus 1.
c) Always AC power.
d) The battery or TR2.
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.
- When will an APU high air flow occur?
a) The aircraft is on the ground, APU bleed air switch is on and either or both pack switches are in High position.
b) The aircraft is in flight, APU bleed air switch is on and either or both pack switches are in High position.
c) The aircraft is on the ground, APU bleed air switch is on and both pack switches are in Auto position.
The aircraft is in flight, APU bleed air switch is on and either or both pack switches are in High position.
- What is the recommended ‘’no load’’ cooling down period before shutting down the APU?
a) 15 seconds.
b) 30 seconds.
c) 45 seconds.
d) 60 seconds.
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.
- How does the APU get its fuel when all AC pumps are not operating?
a) It suction feeds.
b) By a DC driven pump.
c) By the APU fuel pump.
d) A and B is correct.
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.
- When can APU generator supply electrical power to both transfer busses?
a) On the ground or inflight.
b) On the ground only.
c) In-flight only.
d) On the ground and only during climb in-flight.
On the ground or inflight
- Pulling the APU Fire Warning Switch UP accomplish which of the following?
a) Close the fuel APU shutoff valve, APU bleed Air Valve, and APU inlet door.
b) Disarms the APU extinguish circuit.
c) Trips the generator controls relay but the breaker is kept close.
d) Deactivates APU start switch.
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.
- What does Engine fire protection consist of?
a) Engine overheat detection powered by the DC standby bus, engine fire detection powered by the DC standby bus and engine fire extinguishing powered by the hot battery bus.
b) Engine overheat detection powered by the battery bus, engine fire detection powered by the switched hot battery bus and engine fire extinguishing powered by the DC stand by bus.
c) 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.
d) Engine overheat detection powered by the battery bus, engine fire detection powered by the battery bus and engine fire extinguishing powered by the DC stand by bus.
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.
- What is an indication that both loops have failed?
a) There is no cockpit indication.
b) An illuminated FAULT light.
c) Simultaneous overheat and warning light illuminated.
d) The fire warning bell will sound without illumination of the red light.
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
- If the DETECTOR fault lights illuminate during a Cargo Fire Test, what is being indicated?
a) This is normal indication during the Cargo Fire Test.
b) One or more detectors in the loops have failed.
c) The system to check the loops has failed.
d) All the detectors have failed.
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.
- What happens when the engine fire handle is pulled?
a) Close the fuel, hydraulic and bleed air valves.
b) Only trips the GCR (Generator Control Relay) but not GB (Generator Brakers).
c) Enables the thrust reverser.
d) Trips the corresponding air conditioning pack.
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.
- Which of the following fire detection system uses a dual loop configuration?
a) APU.
b) Engine.
c) Main wheel well.
d) Wing body overheat system.
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.
- Fire protection system (detection and extinguishing) is provided for:
a) Engine, APU, and wheel well.
b) Engine, APU, wheel well, and the toilets.
c) Engine and APU only.
d) Engines, APU, cargo compartment, and toilets.
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.
- The FAULT light on OVERHEAT/FIRE PROTECTION PANEL SWITCHES monitors the detector loops of:
a) Engine no.1 and engine no.2 and the APU.
b) Engine no.1, engine no.2 and the wheel well.
c) The APU and the wheel well.
d) Engine no.1, and engine no.2.
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.
- What does the ALTERNATE FLAP master switch do?
a) Opens a flap bypass valve to prevent hydraulic lock of the flap drive unit.
b) Arms the alternate flaps position switch.
c) Energizes the standby Rudder.
d) Closes spoilers shut off valve.
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.
- Which of the following is true concerning the Flight spoilers?
a) There are four (4) flight spoilers located on the upper surface of each wing. Each spoiler is powered by both system A and B to provide isolation and maintain system operation in the event of hydraulic system failure.
b) There are six (6) 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,
c) 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.
d) There are six (6) flight spoilers located on the upper surface of each wing. Each spoiler is powered by both system A and B to provide isolation and maintain system operation in the event of hydraulic system failure.
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.
