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  1. 03.01 AUTOMATIC FLIGHT, FLIGHT MANAGEMENT, NAVIGATION03.01.01 LOW ALTITUDE LEVEL OFFCapturing altitude in the TO and GA pitch modes can cause the B757/767 to be at an inappropriateairspeed at low altitude. Knowing when these altitude captures are likely to occur and recognizing
    their ADI autoflight mode annunciations are the key to dealing successfully with them.These altitude captures generally occur below 1000′ AGL before VNAV or FLCH is selected ontakeoff or before the command airspeed cursor is advanced to clean maneuvering on a go-around.Most often, these events happen when the mode control panel altitude window is set within 2000′of field elevation. At light weights, performance is such that altitudes 3000 to 4000′ above fieldelevation may be captured below 1000′ AGL. In these situations, pilots may want to include thepossibility of an altitude capture in their takeoff or approach brief.
  2. Altitude Capture in the Takeoff Pitch ModeAn altitude capture in the TO pitch mode engages
    the autothrottle in the EPR/N1 mode. Becausethrust remains at takeoff reference EPR while leveling off at the captured altitude, airspeedrapidly increases. This situation is annunciated to the pilot when A/T EPR and ALT CAP aredisplayed on the ADI. The Pilot Training Guide tells us to position the command airspeed bug tothe desired speed and select the autothrottle SPD mode. This allows the autothrottle to maintainthe commanded airspeed at the altitude capture. Selecting SPD also changes reference EPRfrom takeoff to climb.
  3. Flight Mode Annunciations of ALT CAP from the VNAV Pitch ModeOn some takeoffs, the autoflight may capture the Mode Control Panel (MCP)
    altitude window justas, or immediately after, VNAV is selected and the command airspeed bug may not advance allthe way to clean maneuvering. In this case, the autothrottle will engage in the SPD mode, but thecommand airspeed bug is not at the correct speed (i.e., below clean maneuvering speed). Toremedy this situation, advance the command airspeed bug to the appropriate airspeed.
  4. Altitude Capture in the Go-Around ModeCapturing altitude is a normal part of every go-around.
    But, if the capture occurs below 1000′ AGL,the command airspeed bug is still at landing target speed. The autothrottle engages in the SPDmode when altitude is captured in the GA mode, so airspeed will be too slow. To remedy thissituation, advance the command airspeed bug to clean maneuvering.
  5. 03.02 COMMUNICATIONS AND FLIGHT INSTRUMENTS03.02.01 SATELLITE COMMUNICATION SYSTEM (SATCOM)An Iridium Satellite Communication System (one SATCOM Radio with two channels) is currentlybeing installed on the B767 fleet and B757 tails N446, N447, N451-470UP. SATCOM equippedaircraft can be identified by a placard on the Center Instrument Panel above the Autoland StatusAnnunciator.
    • The SATCOM system can provide both data and voice services over greater ranges than VHFcommunications system and is more reliable over oceanic regions than HF communications system.The SATCOM system consists of an Iridium satellite-based communications system toprovide voice and data services. One channel of the SATCOM system is dedicated for voicecommunications and one channel is dedicated for ACARS data communications. The InertialReference Systems (IRSs) do not need to be aligned to make SATCOM calls.The Iridium “brand” system is a network comprised of three components:• Satellite Network• Ground Network• Aircraft SATCOM unit
    • The Iridium system has 66 satellites orbiting the Earth once every 100 minutes. The aircraft unitautomatically searches and uses the best satellites available. The system will switch betweensatellites automatically without dropping calls in progress. There is no initialization or network “logon” required like some other Satellite systems (e.g., INMARSAT). The Iridium system beginssearching for satellites automatically after power up.The SATCOM system is controlled by the ACARS Interactive Display Unit (IDU). On the newB767s (tails N335UP and above), the ACARS unit is upgraded to the Multi-function InteractiveDisplay Unit (MIDU). On all B767s, old and new, the SATCOM can also be accessed through theMulti-function Control Display Unit (MCDU) - (i.e., the FMC CDU).The ACARS priority for data mode selection is:1. VHF2. SATCOM3. HFTo access SATCOM through the ACARS unit, select Main Menu → SAT prompt.
  6. The SATCOM system uses the L (left) HF/SATCOM position on the Audio Selector Panel (ASP)to control voice transmission and receiver monitoring. The L HF switch on the ASP will provideHF communication until a voice SATCOM call is made or received,and then it will automaticallyswitch to SATCOM until the call is completed. Once the call is completed, the relay willautomatically select the L HF again. The default power on position of the automatic HF/SATCOMrelay is the HF position. The R HF position on the ASP is for HF communications only.SATCOM calls can be made one of two ways. The first method is by using the directory ofpre-programmed numbers in the SATCOM system. From the SATLINK menu, select the directoryprompt.
  7. 03.03 ELECTRICAL03.03.01 THE APPROACH MODE (ELECTRICAL EMPHASIS)Selecting the approach mode on the MCP starts a series of electrical isolation events.
    When thethree autopilots are armed for approach and no faults exists, the center busses transfer from theleft main busses to the standby system. The standby electrical system is used as an independentpower source to supply the center autopilot during autoland operations. This provides threeindependent power sources for the three autopilots. They are the two AC/DC busses, and thebattery charger/battery. If the battery charger fails, the battery provides an instantaneous backupfrom the hot battery bus. The battery switch must be ON to enable relays to isolate the center buswhen approach is selected. The APU generator (if operating) will pick up a bus if an engine drivengenerator fails. Both AC bus tie breakers have to be in AUTO for electrical isolation to occur.Let’s say you are performing an autoland, and the left AC bus tie is locked open due to a shorton the bus (ISLN). Will the Captain lose his flight instruments?The answer is no. When approach mode is selected, each flight instrument transfer bus is receivingpower from the respective AC bus. When power is lost on the transfer bus, it will automaticallyswitch to the other power source, regardless of altitude. It is designed this way so flight instrumentsare not lost during an autoland.
  8. Some Other General Considerations• The busses go into isolation mode when approach is selected,
    but the other two autopilotsdo not engage until 1500′.• Fail operational (LAND 3) requires triple redundancy of electrical power sources, engagedFCCs, ILS receivers, IRSs, radio altimeters and hydraulic power to the autopilot servos.
  9. Generator Loss Above 200′ RA With The APU Generator Operating
    • APU generator picks up the failed bus.• LAND 3 will still be displayed on the Autoland Status Annunciator (ASA).
  10. Generator Loss Above 200′ RA With The APU Generator NOT Operating
    • Electrical system reverts back to normal operation.• The bus tie breakers closed to keep both main AC busses powered from the operatinggenerator.• The Flight Control Computers (FCCs) recognize that the isolation has been removed, resultingin center busses switching back to left main AC and DC busses.• LAND 2/NO LAND 3 (i.e., fail passive) will be displayed on the ASA.
  11. Below 200′ RA General Considerations
    • The ASA is inhibited from changing except to indicate NO AUTOLAND.• The Bus Tie Breakers (BTB) go into an isolation mode. In the isolation mode, both BTBscannot close simultaneously to power the two main busses from one source.
  12. Generator Loss Below 200′ RA With The APU Generator Operating
    • The APU can pick up only one failed bus, as long as that bus is not isolated for an electricalfault.
  13. Generator Loss Below 200′ RA With The APU Generator NOT Operating
    • The respective AC bus and autopilot will NOT be powered.• ASA will not change unless both generators fail.
  14. Bus Isolation Condition Is Removed In Three Ways
    1. Above 200′ RA and a fault has been detected which results in a NO LAND 3 condition.2. After initiation of an autopilot go-around when the aircraft is above 100′ radio altitude and apositive rate of climb has been established.3. When all autopilots are disconnected.
  15. 03.04 ANTI-ICE, RAIN03.04.01 PITOT PROBE ANTI-ICINGThe pitot probes are electrically heated in the air by 115 volt AC
    volt AC power through individual circuits toeach probe. Current sensing relays in each circuit provide inputs to the Engine Indicating and CrewAlerting System (EICAS) computer and lights on the Overhead Panel for individual probe heatmalfunction. The EICAS computer displays an advisory message for the malfunctioning probe(s).If high heating occurs on the ground, a maintenance message is recorded by the EICAS computer.On the ground, with either or both engines operating, the two probes on each side will be in LOWHEAT. With both engines shut down, the probes are not heated.A ground test switch allows the probes to be heated with the engines shut down. The probelights will extinguish when the system is tested.
  16. Probe Anti-ice Heat General Operation
    Application of anti-ice heat to the various probes is automatic. Air/ground status and engineoperation determines the mode of heat application. With the aircraft on the ground and no enginesrunning, the probes are unheated.
  17. Pitot Probes
    The four pitot probes are heated in an identical manner. When either engine is running and theaircraft is on the ground, the pitot probes are heated to LOW HEAT (half phase 115 volt AC).Whenever the aircraft is in the air mode, the pitot probes are heated to HIGH HEAT (full phase 115volt AC). When either heat is applied, the indicator light and EICAS message will extinguish.
  18. Angle-Of-Attack ProbesThe Angle-Of-Attack (AOA)
    probes are heated anytime the engines are running or the aircraft is inflight. Application of heat will extinguish the indicator light and EICAS message.
  19. Total Air Temperature ProbeThe Total Air Temperature (TAT)probe, located on the right side of the fuselage just aft of thecockpit, is electrically heated. The TAT probe light on the Overhead Panel will extinguish when anengine is running. However, power to the heater is not applied until the air/ground system senses“AIR.” Bleed air is used to induce airflow across the probe. The TAT probe light and EICAS messageindicates either: Probe is not being heated inflight or neither engine is running on the ground.
  20. 03.04.02 ENGINE NACELLE ANTI-ICING (COWL HEAT)The nacelle anti-icing system receives power from the 28 volt DC battery bus.
    When selected, atwo second time delay on the cowl anti-ice valve control relay creates a disagreement allowingviewing of the valve lights on the anti-ice Control Panel. As soon as the relay energizes, the cowlvalve opens and the limit switch movement turns off the light. EICAS displays a green TAI cueabove the selected engine N1 anytime the engine Anti-Ice switch is on. If one engine bleed switchis off, the N1 indicator for the engine with the operating bleed and anti-ice systems will also displaya minimum required N1 Bug. If the actual N1 on the supplying engine decays below the minimumcommanded value, the TAI Bug changes from green to yellow.The “valve” light in the engine Anti-Ice switch indicates disagreement between actual andcommanded position of the cowl anti-ice valve. Momentary illumination of the valve light indicatesa normal “in transit” condition. If the valve does not reach the commanded position within a giventime frame, an EICAS advisory message is displayed.When the respective Engine Start switch is in AUTO, an engine Anti-Ice switch on will alsocommand continuous ignition on the selected ignitor(s). In a situation where a cowl valve willnot open, QRH procedure will have us leave the switch on to provide ignition while avoiding icingconditions.On the other hand, if a cowl valve will not close, QRH procedure will have us push the switch backon, this time to continue
  21. 03.04.03 ENGINE NACELLE ANTI-ICE COMPONENTS LOCATION
    The nacelle anti-ice valve is a pressure regulating and shutoff valve capable of regulatingdownstream air to 20 to 28 psi. The air is distributed to the nacelle inlet lip through a distributionduct. The air is eventually vented overboard through slots in the bottom of the inlet lip. The fanspinner is always heated, whenever the engine is running, by the engines own internal bleed air.
  22. OperationThe valve is electrically
    signaled and pneumatically controlled. When the engine Anti-Ice switch ispressed, a two second time delay activates to allow disagreement light viewing.
  23. 03.04.04 WING THERMAL ANTI-ICEElectrical Functional DescriptionWhen in the air and the control switch on
    ch on the Pilot’s Overhead Panel is selected ON, a circuit iscompleted through the air/ground relay to the valve solenoids. When the air/ground relay is in theground position, the system can only be operated by the test switch.The white “ON” symbol in the switch indicates that the system is selected on. The amber “VALVE”light illuminates for valve position disagreement. Valve disagreement also generates an advisorymessage on EICAS. When selected ON, the system will display a “VALVE” light until the valveopens. The amber “VALVE” light will then extinguish.Pneumatic Functional DescriptionBleed air entering the wing anti-ice system is regulated by the wing anti-ice valves.The telescoping ducts at the No. 4 and No. 7 slat connect the supply ducts with the spray tubes inthe slats.Spray tubes extend along the cavity of the leading edge slats to spray the heated air against theinside of the leading edge.Flight Compartment Window Anti-icingAll flight compartment windows are heated electrically using two controllers, each with BITE. Thecontrols for the system are on the Pilot’s Overhead Panel.
  24. 03.05 AIR SYSTEMS03.05.01 B757 PNEUMATIC SYSTEM
    On turbine engine powered aircraft, the engine compressor section is a readily available source ofpressurized air (commonly referred to as bleed air) for cabin air conditioning, engine inlet cowl andwing TAI systems, pneumatic drives, etc. In general, the lowest operating penalties result whenthe extracted bleed air characteristics exactly match the requirements of the operating systems.Consistent with other design considerations such as: installation space, equipment weight, controlfunctions, engine compressor stages available for bleed air extraction, etc., also play significantroles in influencing the optimum system design. Current engine bleed air systems use two engine compressor stages, (commonly referred to as lowand high) and an automatic control system to select one of the two stages.
  25. 03.05.02 ENGINE BLEED AIR SYSTEMThe B757 engine bleed air system for the Pratt and Whitney powered engines extract bleed airfrom either the low or high stage of the compressor spool.
    The stage selection is fully automaticand requires no crew action. The system selects the required stage based on the available enginecase pressure which is continuously sensed by the high pressure shut off valve’s pilot. When thesensed pressure exceeds the automatically selected switching pressure threshold, the valve pilotcauses the high pressure shutoff valve to close and permit the low stage to service the operatingsystems. High stage bleed air is used for all conditions when the engine case pressure is lowerthan the selected switching pressure threshold. Note, that while the switching pressure thresholdis a constant over an altitude range, for an engine, the EPR at which switch over occurs is afunction of both altitude and Mach number. In cruise, switch over from the low stage to the highstage is followed by an increase in manifold pressure indication. The high pressure shutoff valveis a pressure regulating valve and maintains the downstream pressure at 55 ± psig. The valvealso has overpressure and over temperature shutoff features. The high pressure shutoff valveautomatically closes and stays closed if system malfunctions cause pressure downstream of thevalve to exceed 130 psig. This event is annunciated by the HI Stage light. Maintenance action isrequired to reset the valve.
  26. The low stage bleed duct has a check valve to prevent high stage flow ingestion into the low stage.
    Bleed air from the low or the high stage passes through an air to air heat exchanger, referred to asa precooler and a Pressure Regulating and Shutoff Valve (PRSOV) before entering the manifold.Inlet cowl TAI, engine start, air conditioning and the wing TAI systems are the main engine bleed airusers. Each of these systems is equipped with a shutoff valve. The cowl TAI system is serviced byprecooled air. This method of servicing the cowl TAI system is the reason for the two switchingpressure thresholds. The PRSOV is a solenoid controlled, pneumatically actuated valve that canbe commanded open for the normal regulation by the control panel ENGINE BLEED AIR switch. Itserves as an ON/OFF, pressure regulating, over temperature shut off valve. The valve is wired tothe Fire switch and serves as the firewall shut off valve. It also checks reverse flow. The PRSOVmaintains the downstream pressure at 45 psig whenever the inlet bleed air pressures are higher.Cooling airflow for the precooler is extracted from the engine fan air duct and is controlled by afan air modulating valve. This valve modulates fan air flow through the precooler to maintainprecooler discharge bleed air temperature.An over temperature sensor (thermostat) located downstream of the precooler and connected tothe PRSOV, is provided to limit the bleed air temperature leaving the precooler. When activated,the thermostat causes the PRSOV to close and reduce flow to stabilize the discharged bleed airtemperature in the range of 450-490°F. This also causes a reduction in the supply or manifoldpressure.
  27. An Overheat switch, with a nominal setting of
    495°F is provided as a back up protective device tothe thermostat.
  28. On actuation, the switch automatically closes the
    PRSOV and the high stage valve, and illuminatesthe BLEED annunciator and the ENGINE BLEED AIR OFF light. Crew action is required to resetthe system.A reverse flow controller automatically closes the high stage valve and the PRSOV on detection ofreverse flow through the precooler. The reverse flow controller is a differential pressure actuatedmicroswitch which on closure causes the PRSOV to go closed. This feature provides additionalprotection against reverse flow during the engine start operation.Temperature sensors are provided along the bleed air ducting to detect bleed air leakage. TheDUCT LEAK lights (one per side) annunciate the overheat condition. The duct leak detectionsystem can be checked by the DUCT LEAK switch located on the accessory panel.Two pressure transmitters are located in the pneumatic manifold, one on each side of the isolationvalve. The sensed bleed air pressures are displayed on the control panel dual pressure gauge withan accuracy of ± 3 psig. These pressures are also displayed on the ECS page of the EICAS.A precooler outlet temperature sensor and a manifold pressure sensor are provided on each sideof the isolation valve for use by the EICAS computer for maintenance messages. The flight deckpanel annunciations are also displayed on the EICAS.
  29. Split Pressure IndicationWith identical engine bleed air systems on either side of the normally closed isolation valve, it isreasonable to expect essentially the same pressure indications on both sides of the isolation valve,when operating at the same EPR. However, different pressure indications are often noticed duringhigh altitude cruise, at the end of cruise or while cruising at a low gross weight. A split pressureindication during above mentioned operating conditions is not a positive indication of a malfunction.It can occur, in a normal system, due to one or more of the following:
    • Tolerance of system control valves. The PRSOV and the high valve each have a pressureregulation tolerance of ± 5 psig.• Tolerance of the duct pressure transmitter and gauge indication system.• Engine bleed pressure variations between engines due to deterioration or variations in enginebuild tolerances, etc.• One engine bleed air system operating on the low and the other on the high stage bleeddue to tolerances in the switching pressures.The first three items above, can cause large differential pressure indications if the controlcomponents of the two engine bleed air systems are operating at, or close to, the extremes of theallowable tolerance band. These items cause split pressure indications when the system valvesare in control, that is, during moderate to high power operations. During low power operations, when the control valves are essentially wide open, the pressure differential generally reduces. Aconsistent differential pressure greater than 10 psig during climb power, though not an indication ofa malfunction, should be noted and may be cause for maintenance action.The last item above, can cause differential pressure indications of 20-30 psig. These generallyoccur at high altitude and a low engine thrust when the engines are operating near the switchingpressure threshold or switchover EPRs.From the above discussion, it would be apparent that split pressure observation is not a positiveindication of a fault or premature selection of the high stage bleed air. Switching from low tohigh stage occurs with decreasing EPR and is accompanied by a sudden increase in manifoldpressure. Switching from high to low occurs with increasing EPR and is accompanied by adecrease in manifold pressure.
  30. Pressure CyclingSome operators have
    encountered pressure cycling on one or both sides of the isolation valveduring some cruise conditions. This generally occurs at thrust levels where the engine casepressures are close to the switching pressure threshold, or the operating EPR close to theswitchover EPR. Pressure oscillations are accompanied by cabin air flow oscillations as the airconditioning pack flow control valves attempt to follow the delivered pressures. The pressurecycling occurs due to switching between the low and high stages. This condition has also beenobserved on aircraft where sensing lines leading to the PRSOV were loose.From the above discussion, it seems apparent that if the operating EPR is close to the bleed stageswitching EPR range and small cyclic changes in the EPR occur, a normally operating system mayoscillate between the low and high stage bleeds delivering cyclic pressure to the manifold. Thesmall engine EPR changes may stem from interactions and activities of the flight managementsystem, thrust management system, engine bleed (surge control) valves, component hysteresis,tolerances, etc. Some operators have advised that changing the operating EPR up or down,stops pressure fluctuations.Technically, selection of a different EPR moves system operation away from the switching pressurethreshold. Increasing the operating EPR increases engine case pressure and causes the enginebleed air system to preferentially select the low stage bleed. Decrease in the operating EPRcauses the system to select high stage bleed.
  31. SummaryThe B757 engine bleed air system has been designed to minimize fuel burn penalties imposedby the pneumatic systems.
    It has defined engine case pressure thresholds for switching from thelow to the high stage and vice versa for low (<FL310) and high (>FL310) altitude operation. OnPratt and Whitney powered aircraft, it is also dependent on the cowl TAI system selection. It usespneumo-mechanical controls for bleed stage selection and bleed air pressure and temperatureregulation. Individual components have industry accepted tolerances. However, these tolerances inconjunction with the automatic bleed stage selection technique used, sometimes provide indicationwhich can be interpreted as faults. The probability of this occurring is high at EPRs close to thebleed stage switching EPRs. Such an operation can cause split pressure indications and/orpressure cycling. These in turn can cause secondary effects such as, cabin inflow noise, cabinaltitude fluctuations, etc. The Boeing Company continues to make improvements in the overallsystem. In the interim, the best advice is to avoid operation in the region of bleed stage switching.
  32. 03.06 PERFORMANCEWhen does the 15 degree bank limit apply when following a special engine out departureprocedure (i.e., pink page)?
    The special engine out departure procedure protects for obstacle clearance up to 1500′ AGL orOCA (if published), whichever is higher. If the procedure calls for a turn before reaching 1500′ AGLor OCA, the turn is based on a maximum bank angle of 15 degrees. (The AFM does not have datafor climb gradient reductions for bank angles greater than 15 degrees.)Once the aircraft attains 1500′ AGL or OCA, whichever is higher, the 15 degree bank limit no longerapplies. Above this altitude, the pilot assumes responsibility for assuring terrain clearance.As a general rule on any takeoff, obstacles are analyzed up to 1500′ or OCA. Once the aircraftattains 1500′ and is clear of obstacles, the pilot assumes responsibility for assuring safe terrain andobstacle clearance.Let us review the original purpose of:• Engine-out departure procedures.• Approach/missed approach procedures.• Climb limit weight for landing AAM data.
  33. 03.06.01 ENGINE-OUT DEPARTURE PROCEDUREThe engine-out departure procedure (whether straight out or special) is based on an enginefailure just prior to V1.
    It ensures terrain clearance by a “net” height of 35′ for all obstacles up toat least 1500′ AFE. The data for this analysis comes straight from the AFM. The AAM data wasnot designed to be used for an engine-out go-around. However, in many cases, it is a “good”source for evaluating lateral guidance.Furthermore, the AAM data (in conjunction with the engine-out departure procedure) makes manyassumptions that are unlikely to exist during the engine-out go-around. Here are just a few:The engine-out departure procedure (whether straight out or special) assumes that the enginefails just prior to V1. Furthermore, it is assumed that the aircraft is at maximum takeoff grossweight for that runway. The procedure also assumes that the engine-out takeoff profile is precisely followed. As you can imagine, the engine-out takeoff profile is significantly different than theengine-out go-around procedure.Finally, there is no readily available FAA approved data which could be used to fully analyzeobstacle clearance during an engine-out go-around. The engine-out departure procedure is oftenthe best source for approximating the engine-out go-around flight path.
  34. 03.06.02 PUBLISHED APPROACH/MISSED APPROACH PROCEDUREThe standard published
    approach and missed approach procedures are based on TerminalInstrument Procedures (TERPS). It is simply intended to offer obstacle protection for an aircraft thatcan meet the standard or non-standard minima climb gradients incorporated in the TERPS analysis.
  35. 03.06.03 CLIMB LIMIT WEIGHTS FOR LANDINGThe AAM climb limit weights for landing are based
    on the lesser of APPROACH CLIMB LIMIT(engine-out, approach flaps, gear up) and LANDING CLIMB LIMIT (all engines operating, landingflaps, gear down) weights. For two engine aircraft, the approach climb limit weight is alwaysmore limiting.03.06.04 PUTTING
  36. 03.06.04 PUTTING IT ALL TOGETHERPoint one, there is no data specifically intended to ensure terrain clearance during an engine-outmissed approach. The Captain must use good judgment, terrain awareness and all availableinformation during the emergency.Point two, both the engine-out departure
    procedure and “standard published” missed approachprocedures may offer a safe course to fly during an engine-out missed approach. A non-standardmissed approach (i.e., requiring a higher than standard climb gradient) may exceed the engine-outclimb gradient, and therefore should normally be avoided during an engine-out go-around.Point three, all of this engineering talk assumes very specific conditions, such as the precise pointof engine failure, flap retraction procedures, etc., that will never occur during an actual go around.Remember, there is no data for this. Therefore, Captain’s judgment is paramount. Consequently,understanding the limitations of the data (engine-out departure procedure and the standard missedapproach procedure) is key to making a sound decision. For this reason, I would not fault a Captainfor electing to do either the special engine departure procedure or standard missed approachprocedure during a engine-out go-around.Unfortunately, the lack of data make it difficult to identify the “one right solution.”There is no right answer. You have a piece of information that indicates the special engine-outmay be the best course of action.A significant point worth mentioning: Many crews have a habit of requesting a “straight out” missedapproach during a single engine approach. Presumably, this is because the crew assumes it is easier to fly runway heading rather than the published missed approach during the singleengine missed. From an obstacle clearance standpoint, this choice definitely puts the crew in theproverbial “No Man’s Land
  37. 03.07 ENGINES, AUXILIARY POWER UNIT (APU)03.07.01 APU OIL SYSTEM
    GeneralThe APU oil system provides oil under pressure to lubricate and cool the main bearings and thegearbox. The APU gearbox (accessory drive) serves as the oil reservoir and has a capacity of 6.2quarts. The APU oil pump is an integral part of the APU gearbox. The APU control unit will initiatea protective shutdown when low oil pressure or high oil temperature is sensed. This illuminates theAPU fault light as with other protective shut downs of the APU.Maintenance checks the APU oil quantity by viewing a sight glass on the APU (the B757 doesnot have oil quantity indication in the cockpit). If the oil level reaches the “ADD” level (about twoquarts low), it will be filled to the full level. EICAS will generate an “APU OIL QUANTITY” statusmessage when the oil quantity reaches this (two quarts low) level.Referring to the MEL “EICAS” section, this message means that with normal oil consumption, theAPU can continue to operate for an additional 30 hours after the message appears. If this messageappears on the ground, it will not erase using the switches on the accessory panel and must beentered in the aircraft logbook. If the message appears after dispatch, write it up and send amessage via ACARS. An APU shut down is not required.The B767-300 freighter has an APU oil quantity indication on the status page. The possibledisplays are FULL, .75, .50, .25 or ADD. A display of .50 does not mean half full or half empty, itmeans one quart low. “ADD” means two quarts low and like the B757, two quarts low generatesthe “APU OIL QUANTITY” status message.This means we have 30 hours of run time remaining.APU oil quantity checks and servicing are more frequent on the B767 to comply with ETOPScertification requirements. Operating under the ETOPS program requires us to make periodicinflight starts of the APU to prove inflight start reliability. This program collects data via flightcrew reporting on ACARS.Both the B757 and B767 use the same model APU which is manufactured by Garrett. TheAPU generator can provide its full load capability to the aircraft’s service ceiling and bleed air isavailable to approximately 17,000′.As with the B757, the B767 APU can be started and operated at altitudes up to the aircraft’sservice ceiling.
  38. APU Fuel SystemThe APU receives fuel from the aircraft’s fuel tanks via a shrouded line which runs through thefuselage. If the line were to leak, the shroud would catch the fuel and vent it overboard through adrain mast located on the right side of the tail cone (see Figure 2).
    Due to the aircraft’s fuel system design, the APU normally feeds from the left main tank (seeFigure 3). With no AC power available (battery start of the APU), a dedicated DC fuel boost pumppowered by the battery bus will draw fuel directly from the left main tank and deliver it to the APU.When AC power is established and the left forward boost pump produces pressure, the DC pumpshuts down automatically. The left forward boost pump, which is powered automatically duringAPU operation, feeds the APU via the left engine fuel feed manifold. Therefore, any boost pumpcould potentially feed the APU if the left forward pump was not operational.Fuel must then pass through the main APU fuel shutoff valve and then through the shrouded lineback to the APU fuel control.
  39. The APU fuel control includes another SHUTOFF valve (this one solenoid operated) through whichfuel must pass before reaching the nozzles in the combustion chamber
    The main APU fuel shutoff valve, at the tank, is monitored by the APU fault light on the OverheadPanel. If this valve is not in the commanded position, the fault light illuminates. The light is normalduring valve “in-transit” operation. However, if the valve fails to reach the commanded positionwithin time limits, it will generate the “APU FUEL VAL” advisory message. With this messagedisplayed, the QRH states, DO NOT START THE APU.In the event of an APU fire, the auto fire shutdown system will close both fuel valves to “flame-out”the APU. The APU fire switch, when pulled, also commands both fuel valves to close.The following fuel system question comes up quite often, “What about the limitation that says youcannot operate with more than 2000 lbs. of fuel in the center tank, with less than full mains, and theAPU has reduced my left tank down below full?” (Refer to AOM “Fuel Loading” limitation). Theanswer, no problem, as long as the FUEL CONFIG light is not on, and the right tank is full. You canhave all the fuel you want in the center tank, as long as the imbalance between the full left tank andthe full right tank is less than 1800 lbs. due to APU fuel burn. This brings up another question:“What is full?” The “NOMINAL” full value of the main tanks based on standard fuel density (6.7 Ibs.1 U.S. gallon) is 14,600 lbs. Our fuelers generally use 14,500 as full for the mains and put theadditional required fuel in the center tank.
  40. This technique of using 14.5 is just one more safeguard
    to prevent overfilling and allow for expansion. Occasionally, you may see more than 14.6 in themains, and this to, would be considered full.On the B767, 22,050 lbs. of fuel can be loaded in the center tank with less than full mains. TheB767 APU also feeds from the left tank. The imbalance limitation is, however, slightly morecomplicated. With full mains on the B767, (40.7 each), the total main tank fuel load exceeds79,800 lbs., and therefore, the imbalance limit is 1500 lbs. (See AOM Max Fuel Imbalance in AOMlimitations). A greater imbalance is allowed as the total main tank load is reduced.If the APU on a B767 reduces the left tank below full, as long as the imbalance between themains does not exceed 1500 lbs. (and the right tank is full), we can dispatch with any amount offuel in the center tank.If both main tanks are less than full, a maximum of 2000 lbs. (B757) or 22,050 lbs. (B767) may beloaded in the center tank as long as the sum of the center tank plus the actual zero fuel weightdoes not exceed the maximum zero fuel weight of the particular aircraft.
  41. 03.08 ENGINES03.08.01 ENGINE OIL SYSTEM
    GeneralThe purpose of the oil system is to lubricate and cool the engine bearings and the accessory drive.The engine bearings support the N1, N2 and N3 (RR) rotor shafts. The accessory drive (gearbox) islocated on the underside of the engine and is driven by the N2 (N3 RR) spool. The Integrated DriveGenerator (IDG) has its own independent oil system which does not mix with the engine oil.
  42. Engine Oil Pressure System And IndicationOil is supplied under pressure to the engine bearings by the oil pump, which is driven by theaccessory drive.
    The low oil pressure alerts provided to the flight crew consist of an EICAS display color change(amber or red) and a discrete OIL PRESS light for each engine. The source of these oil pressurealerts come from separate sensors.The EICAS oil pressure indicator receives signals from a pressure transmitter while the ENG OILPRESS light and message is signaled by a separate pressure switch. This redundancy not onlyprovides a backup indication but also gives the crew the ability to troubleshoot a faulty or erroneouslow oil pressure indication.If an engine oil pressure indication on EICAS turns amber or red, the OIL PRESSURE procedurein the QRH must be accomplished. This procedure varies slightly by engine type, but all three(PW, RR, GE) prescribe accomplishment of the ENGINE FAILURE/SHUTDOWN checklist if oil pressure is at or below the red line limit. The color change on the EICAS display may or may not beaccompanied by the respective ENG OIL PRESS light. If the light is not illuminated in conjunctionwith the color change on EICAS, it could be due to the following:• Amber band and red line values are variable and based on engine RPM while the low oilpressure switch is activated by a fixed low pressure value.• A faulty or erroneous indication.Recall that when the EICAS computer selector is in the AUTO position, the left EICAS computeris utilized. Automatic selection of the right EICAS computer will only occur if the left computerfails. A fault occurring in an individual indication will not result in an automatic switching to theright computer.If the crew suspects a possible EICAS indication fault, the right EICAS computer can be manuallyselected for comparison purposes. As with any engine oil system non-normal, the other oilindicators (in this case, temperature and quantity) can also be checked to help verify the apparentabnormality.The ENG OIL PRESS light will illuminate when the oil pressure switch senses less than 70 psi(PW), 18 psi (RR) or 10 psi (GE). As mentioned earlier, this is a fixed value and does not vary withengine RPM. This will also generate a L or R ENG OIL PRESS advisory message. If this alert isnot accompanied by a color change in the EICAS oil pressure display, the switch could be faulty. Ifthe EICAS oil pressure indication is normal, the B757 QRH states, “operate engine normally.” TheB767 QRH currently does not address this situation.
  43. Oil FilterEngine oil passes through the main oil filter and then, prior to entering the bearings, through“last chance filters.”If the main oil filter is cloggingwith contaminants, a differential pressure will develop across thefilter eventually generating the OIL FILTER advisory message. This message was designed toalert the crew/maintenance that the oil filter needs to be checked. The OIL FILTER procedure inthe QRH must be accomplished if this message is displayed.If additional contaminants further block the filter, differential pressure will rise to a value which willopen a bypass, sending contaminated oil downstream to the last chance filters. The last chancefilters are designed to stop the larger particles from reaching the bearings.NOTE: The main oil filter on the RR engine does not have bypass capability.On PW and GE engines, if the OIL FILTER message cannot be extinguished by reducing thrust,the QRH prescribes accomplishment of the ENGINE FAILURE/SHUTDOWN checklist. Thismessage on RR powered B757s is not addressed in the QRH other than a crew awareness typecondition statement.
    On PW powered B757s, the OIL FILTER checklist allows operation on the ground with thecorresponding message displayed if oil temperature is 35°C or less. This is due to the possibilitythat cold (thick) oil may be the cause. The QRH directs the crew to wait until oil temperatureincreases above 35°C. If the message remains with oil temperature above 35°C, the engineshould be shutdown.
  44. Engine Oil Temperature Indication and Control
    Engine oil is cooled by a fuel/oil heat exchanger and also by a fan air/oil cooler. The primarypurpose of the fuel/oil heat exchanger is to provide fuel heat. Fuel temperature is regulatedautomatically by controlling the amount of hot oil which flows through the heat exchanger.A probe senses the temperature of the oil as it leaves the engine on its way back to the tank. If oiltemperature reaches the upper amber band or red line limit, the EICAS display will turn amber orred. There is no associated EICAS message, but there is a QRH procedure (OIL TEMPERATURE)which must be accomplished.Only the Rolls-Royce engine has a low oil temperature amber band. When oil temperature is 0°Cor less, the display color changes from white to amber. As per the AFM limitation, as long as oiltemperature is -40°C or warmer the engine can be started. The display will remain amber until theoil warms up above 0°C, at which time the thrust levers can be advanced
  45. Engine Oil Quantity
    After engine shut down, a mechanic will typically observe the oil quantity on EICAS, and then on awalk-around, check each tank. Although there is no minimum oil quantity for dispatch, mechanicswill typically add oil to the 19-20 quart level to avoid over servicing.If a flight crew observes an abnormally low oil quantity when accomplishing the cockpit set-up, itcould be due to oil migration to the gearbox, or the result of a short ground run-up of the enginewhich has left oil in the bearing compartments. It could, of course, be due to an actual low oilquantity situation. If you are not comfortable with an indicated oil quantity, write it up and callmaintenance.During engine start, the oil pump fills the bearing compartments and gearbox with oil, whichcan result in a substantial indicated quantity decrease. This “gulping” is normal, and a quantitydecrease of up to 9 quarts (PW) below the pre-start value may be observed (less on RR and GE). Ifa large quantity of oil has migrated to the gearbox, it is also possible to see a quantity increaseafter the initial “gulp” as the scavenge pumps return oil to the tank.The oil quantity indicator on the lower EICAS display can indicate from 0 to 23 quarts of oil.Maximum indication on the B767 is 22 quarts. There is no red line, no amber band and thereforeno associated color changes with this indicator. The white “low quantity” awareness band indicatesfour or less quarts, but does not generate any status or alert level messages or lights. There is norequired crew action for low engine oil quantity unless accompanied by low oil pressure and/orhigh oil temperature.
  46. 03.09 FIRE PROTECTION03.09.01 GENERAL

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