B73 37T 7Th heo ory ry Manu Ma ua al
Bo oein ng 737N 7 7NG S stem Sys ms s Revis Re sion n dat te 05 5-08 8-15
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Forew word: This boo oklet describees systems p published in o our Faceboo ok pages: About This FB p page is to intteract througghout the B7 737 commun nity and has N NO direct link to any user companyy. THE CON NTENT SHALLL NOT BE USED FOR ACTUAL OPERATTION OF THE AIRCRAFT. The adm ministrator haas NO RESPO ONSIBILITY to o the content written on these pagess. Descripttion Administtrators: Ferdi Colijn: B737 7NG Type Rated
B Bert de Jong: Instructor Flight Engineer P‐3 3 Lockheed O Orion B737 7NG Ground School Instructor
B737Theory The goall of this FB page is to exp pand B737 th heoretical knowledge am mongst users and we try tto achieve that by expaanding the am mount of visitors aiming for interaction. There reest no copyrigght on our sttories but wee rather see you recomm mending us o on your privaate FB pages iso o sharing thee posts. Also feell free to "don nate" your eexperiences aand stories o on B737Theo ory and drop us a line by sending a messagge. We will eevaluate and d post them in time but b be aware that it must nott be a copy from any manual o or else we in nterfere with h copyrights. Thank yo ou
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Contents: Foreword: ................................................................................................................................................ 2 APU .......................................................................................................................................................... 8 Auto Slat System...................................................................................................................................... 9 Engine Electronic Control (EEC) ............................................................................................................. 10 When things go wrong and beyond basic systems knowledge ............................................................. 11 Engine fire detection ............................................................................................................................. 13 Feel Differential ..................................................................................................................................... 14 Fuel Scavenge Jet Pump ........................................................................................................................ 15 Fuel valves ............................................................................................................................................. 16 AC Generator ......................................................................................................................................... 17 Isolation valve ........................................................................................................................................ 19 Manual gear extension. ......................................................................................................................... 20 Mechanical pressure relief valves. ........................................................................................................ 21 Nitrogen Generating System ................................................................................................................. 22 Outflow valve. ....................................................................................................................................... 23 Flight Control “Breakaway” Devices ...................................................................................................... 24 Pack & pack control ............................................................................................................................... 25 Recirculation fans .................................................................................................................................. 26 Hydraulic Reservoirs .............................................................................................................................. 27 The APU Starter/Generator. .................................................................................................................. 28 Landing Gear Transfer Valve ................................................................................................................. 29 PTU ........................................................................................................................................................ 30 Wing Thermal Anti Ice (WTAI) ............................................................................................................... 31 B737 Yaw damping ................................................................................................................................ 32 Zone temperature control ..................................................................................................................... 33 Lavatory “fire protection”. .................................................................................................................... 34 Center tank boost pumps ...................................................................................................................... 35 Antiskid .................................................................................................................................................. 36 Leading Edge Flaps ................................................................................................................................ 37 Thrust Reverser ..................................................................................................................................... 39 Tail Skid .................................................................................................................................................. 41 Vortex generators.................................................................................................................................. 42 Window heating .................................................................................................................................... 43 Wing& Body Overheat ........................................................................................................................... 44 Horizontal Stabilizer Trim. ..................................................................................................................... 45
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Display Electronic Units. ........................................................................................................................ 46 Proximity Switch Electronic Unit ........................................................................................................... 47 Nose wheel steering lockout ................................................................................................................. 48 Weather radar ....................................................................................................................................... 49 Dissolved air .......................................................................................................................................... 51 Frangible fittings .................................................................................................................................... 52 Rudder(vertical stabilizer) load reduction ............................................................................................. 53 Rejected Takeoff – speed brakes relation. ............................................................................................ 54 Electrical Bus (bar) ................................................................................................................................. 55 Crew oxygen system .............................................................................................................................. 56 Main system hydraulic pumps, (corrected) ........................................................................................... 57 Cockpit Voice Recorder System ............................................................................................................. 58 Pressure control .................................................................................................................................... 59 Runway Awareness and Advisory System (RAAS) ................................................................................. 61 Electro Motor Driven Pumps Overheat ................................................................................................. 63 Cockpit panel “+” symbols. .................................................................................................................... 64 Overhead (P5) panel drains. .................................................................................................................. 65 Closed crossfeed valve on takeoff and landings? ................................................................................. 66 Amber AUTO BRAKE DISARM Light ....................................................................................................... 67 B737 Fire protection .............................................................................................................................. 68 Start switch functions. ........................................................................................................................... 69 Fuel nozzle “coking”. ............................................................................................................................. 71 Dual bleed light ..................................................................................................................................... 72 Air Cycle Machine operation ................................................................................................................. 73 Airstair ................................................................................................................................................... 74 Equipment Cooling ................................................................................................................................ 75 Overboard Exhaust Valve ...................................................................................................................... 76 Thermal electrical protections. ............................................................................................................. 77 Fuel temperature indication. ................................................................................................................ 78 Integrated Drive Generator (IDG).......................................................................................................... 79 Electrical Load Shedding ........................................................................................................................ 81 Common Display System (CDS) malfunctions. ...................................................................................... 82 Cargo Compartments air. ...................................................................................................................... 83 NiCad Battery operation. ....................................................................................................................... 84 Climb Thrust Reduction ......................................................................................................................... 86 The “white bug”. ................................................................................................................................... 87 Standby Hydraulic System operation. ................................................................................................... 88 Transformer Rectifier Units. (TRU) ........................................................................................................ 90 5
RAM AIR DUCT doors. ........................................................................................................................... 91 Standby Power. ..................................................................................................................................... 92 Fueling panel ......................................................................................................................................... 93 Brake accumulator ................................................................................................................................ 95 Control column shaker .......................................................................................................................... 96 Wheel thermal fuse plugs. .................................................................................................................... 98 Battery busses ....................................................................................................................................... 99 Electrical schematic ............................................................................................................................. 100 Fuel schematic ..................................................................................................................................... 101 Hydraulic schematic ............................................................................................................................ 102 Bleed schematic .................................................................................................................................. 103 Air condition schematic ....................................................................................................................... 104 Engine oil & fuel schematic ................................................................................................................. 105 Flight Mode Annunciations (FMA) ...................................................................................................... 106 Power Sources (NG) ............................................................................................................................ 108
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APU The APU U is a constant speed (± 49.000 RP PM) gas turbine engine that can su upply AC po ower and pressurizzed air. Thee starter/gen nerator is po owered from m either direcctly the main battery (28VDC) or transfer bus 1 (115V VAC) where eeither sourcee is converte ed into 270V VDC for starter operation n. At 95% starter o operation revverses to a 9 90 KVA geneerator, indicaated by the blue APU OFFF BUS light.. (90 KVA until 32..000 ft. and 66 KVA unttil 41.000 ftt.) Starter se equence is automatically a y determine ed by the Electronic Control Unit (ECU) thaat needs the battery swittch to be in tthe ON positiion to energize. The APU U can be useed for air and d AC power until 10.000 0 ft., just air to 17.000 fft. and just A AC power until 41..000 ft. Thatt is also the maximum sttarting altitu ude although h recommen nded at 25.000 ft. Air takes th he biggest performance from the APU A as it takes air from m the load ccompressor which is mounted d on a comm mon shaft wiith the comb bustion comp pressor. The more air taken in, the lower the performance of the APU. That is why theree is a restricction in altittude use, especially with h air and d by IGV’s tow ward the loaad compresso or. When when the demand iss large (high EGT), air usee is squeezed on suctio on feed the APU draws ffuel from tank #1 and w when operating for an exxtended time e select a fuel pum mp to pressure feed whicch extends th he lifetime off the APU. The ECU U protects the APU and sshuts down w with a low oil pressure, o overspeed or when a FA AULT light illuminattes. The lattter represen nts more thaan just the foregoing, including ECU failure, lo oss of DC power, A APU fire, oveertemp (during start), hiigh oil temp and many m more. The start limit is 2 2 minutes and a FA AULT light illuminates wh hen the start is aborted through a p protection orr when the ggenerator malfuncttions. A bluee MAINT ligght illuminates when oil quantity is low or a geenerator malfunction occurred d, the APU iss still allowed d to operate.. APU com mpartment and oil cooling is accom mplished byy exhaust air used as aan educator to draw outside aair into the ccompartmen nt from an inlet just abovve the exhaust. When th he APU is sttopped by placing p the switch s to OFFF, the ECU determines a cooling cycle of 1 minute b before the A APU actually stops. The ccooling cycle e closes the A APU BAV and trips the ggenerator OFF line. By doing so o it reliefs the APU from load and decreases the EGT preventting so called d cooking of the no ozzles. (resid dual fuel form ms carbon on n the hot nozzzles which ccan affect the flame patttern) Delay sw witching the Battery to O OFF to 2 min nutes after se electing the APU to OFF,, this allows the inlet door to close. The door closes when the APU A decelerrates to ± 30% to preveent the inlett duct to ute is by‐passsed when the APU shuts down throu ugh a malfunction, the Fire Switch collapse. The 1 minu is activatted or when the Battery Switch is sellected to OFF.
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Auto Slat System The Auto Slat system operates the LE slats automatically in flight when you’re approaching a stall under certain conditions just before the stick shaker becomes active. These conditions are when the flaps are at position 1 – 5 and hydraulic pressure is available through: • Hydraulic system B • PTU (extend & retract) • Standby hydraulic system (extend only) * With alternate Flap use, the Auto Slat function is not available. * With a short field performance configuration the Auto Slat operates with flap selections 1 – 25. At the flap position 1 – 5 the LE slats are in the intermediate (extend) position and the LE flaps at their only extended position . . . FULL. When the aircraft approaches the stall angle/speed region determined by the Stall Management and Yaw Damper (SMYD) computer, the Flaps/Slats Electronic Unit (FSEU) command the LE slats to the FULL extend position to prevent entering a stall condition. Another action by the FSEU is to delay the “transit lights” to operate for 12 seconds thereby preventing the LE devices transit lights to illuminate. When thrust is increased/stick force relaxed and the aircraft flies out of this condition (higher speed, lower AOA) the Auto Slat system drives the LE slats back to the intermediate extend position. Also here the transit lights will not illuminate. When the Auto Slat systems fails to operate or is not available by any cause, the AUTOSLAT FAIL indication illuminates on the flight control panel. When 1 SMYD computer fails the other will automatically take over and would go unnoticed unless you press RECAL during an Auto Slat condition.
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Engine Electronic Control (EEC) The EEC is mounted on the right top side of the fan duct and exists of two computers (channel 1 & 2), where one is active and the other standby although they’re both operating and cross linked during normal operation. The EEC receives numerous environmental and engine input signals to calculate fuel and control outputs to operate the engine and identifies the engines thrust rating by a pre‐ selected identification plug. Doing so it heats up and needs to be cooled which is achieved by tapping off, and directing fan air to the EEC. Normal power source of the EEC is an alternator mounted on front of the engine gearbox but is only valid when the gearbox (N2) reaches 15%. Before 15% N2, the EEC is powered by Transfer Bus 1 or 2 (Eng. 1 or 2) if available, and becomes energized when the Start Switch is placed to GRD or CONT or, when the Start Lever is moved to IDLE. A de‐energized EEC is indicated by blank engine indication boxes on the upper and lower DU’s even when the EEC button illuminates a white ON, just indicating that the EEC is selected to the normal mode. In this case the only indication visible directly from the sensors are N1, N2, Oil quantity and the vibration indicator, all others are blank. So . . . during a battery start (emergency power), indications of EGT, fuel flow, oil pressure and oil temperature remain blank until the alternator reaches 15%. On the aft overhead engine panel there are the two guarded EEC control buttons to select the EEC to the NORMAL mode of operation (white ON light), or the manual HARD ALTERNATE mode of operation (amber ALT light). An undispatchable failing EEC is indicated also on the engine panel by a ENG CONTROL light and will only illuminate when on the ground and the engine N2 >50%. A little teaser . . . . the last indication on the engine panel are two REVERSER lights . . . when and how long do they illuminate amber during normal operation?
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When things go wrong and beyond basic systems knowledge The next post is an actual situation that happened, losing a Transfer Bus in flight. I’ve tried to simplify the explanation but in fact it’s just an indicator of what CAN happen. At this point Non Normal Procedures, CRM and common sense is needed to fly out of these situations. It started with a MASTER CAUTION and a right SOURCE OFF, indicating that XFR bus 2 was not powered by its “last selected source” but by Transfer Bus 1. QRH tells us to select the GEN switch (affected side) ON what this time caused a TRANSFER BUS 2 OFF to illuminate with additional related indications. (DEU 2 and others, (check the power source booklet to find out) Next the APU was started and when attempted to connect the generator, a BATTERY DISCHARGE illuminated indicating an excessive discharge of a battery, with multiple additional indications. The crew decided to stop further procedures and investigation and used the system “as is”. To give you an idea, the Indications involved: battery discharge, master caution, right hand source off, right hand transfer bus off, Mach trim fail, auto slat fail, fuel pump 2 fwd., fuel pump 1 aft, electrical hydraulic pump #2, probe heat B, engine EEC alternate, zone temperature. After this ordeal the crew managed to land safely with this reduced electrical power condition and multiple caution indications. What actually has happened was that the Generator Control Unit (GCU) 2 had received an erratic signal through the Line Current Transformer (LCT) that IDG2 was connected to the transfer bus. This signal is then transferred to the Bus Power Control Unit (BPCU) who arranges switching in the electrical AC system to provide in the two major rules: • No paralleling of AC sources • An AC source connecting to a Transfer Bus disconnects the previous source (look at the first rule) This erroneous signal locked out the possibility to connect the APU or other AC sources like Transfer Bus 1 to Transfer Bus 2. However, as IDG 2 in fact was not connected, transfer bus 2 lost power. The erroneous indication must have originated at the GCB 2 (unit connecting IDG 2 to bus 2) itself, indicating the switch had closed although it had not moved.
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The BATTTERY DISCHA ARGE is prob bably caused d by the a (exxcessive) main battery diischarge by p powering the Batttery Bus as aalso the DC 2 2 system (TR R 2 & TR 3) were not po owered anym more and illu uminates when a b battery outp put condition ns exists of: • Current draw C w is more thaan 5 amps fo or 95 seconds • Current draw C w is more thaan 15 amps ffor 25 seconds • Current draw C w is more thaan 100 ampss for 1.2 seco onds. Mind yo ou, normally when Transsfer Bus 2 is de‐energize ed the Transffer 3 Relay w would switch h TR 3 to Transferr Bus 1 which h obviously d didn’t happen.
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Engine fire detection The engine fire detection system consist of a fire, and an overheat detection inside the nacelle which are only active when the engine is operating. Temperatures are guarded by 2 (A & B) detector loops which operate by expanding gas pressure inside the loop elements thereby activating an OVERHEAT, a FIRE or a FAULT (leaking loop tube) contact. The engine areas covered by the loops are inside the nacelles around the fan, and the “core” hot section so . . . a torch (see image) would go undetected as it occurs inside the engine. • OVERHEAT detection is indicated by an OVHT/DET, 2 MASTER CAUTION and respective ENG OVERHEAT indication. (± 170°C around the fan section and 340°C around the hot section) • FIRE detection would be indicated by 2 MASTER FIRE WARNING, the respective FIRE SWITCH, an OVHT/DET, 2 MASTER CAUTION and an audio FIRE BELL warning. (± 300°C around the fan and 450°C around the hot section) When either of the foregoing occurs the fire switch unlocks to allow it to be pulled up. A fire or overheat is detected when both loops exceed the mentioned limits and when one loop fails, it’ll go unnoticed and the detection system automatically switches to a single loop operation. One failing loop will only illuminate a FAULT during a test (also not on RECALL) and when both loops fail, the FAULT light illuminates but NOT the MASTER CAUTION. The detection tests on preflight are: • The OVHT/FIRE test which checks the operation of the engine & APU fire detection control module located in the E&E bay and not to forget the indications on the flight deck. • A FAULT/INOP test checks the FAULT detection circuits (loops and elements) and the flight deck indications by simulating a dual loop failure. Note that the APU fire detection also operates during the FIRE test and is visible/audible in the right main wheel well on the APU Ground Control Panel during pre‐flight.
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Feel D Differen ntial The FEELL DIFF PRESSS indication o on the flight control pane el can illumin nate in the fo ollowing case es. s “aactual feel forces” f at the t control column fro om the hydraulically (The feeel system simulates supporteed elevator p panels) 1. The fiirst one is reelated to a differential d o A & B hyd of draulic presssure to the eelevator feel system. When either hydrau ulic system pressure p dro ops > 25% re elated to thee higher preessure, the FEEL F DIFF ol panel witth a 30 seco ond delay. TThe 30 second delay PRESS light illuminaates on the flight contro preventss the light from “flickerin ng” when prressure drop ps in either system by a h high demand d such as gear seleection. 2. The seecond is relaated to the dynamic air p pressure supply to the Eleevator Feel C Computer. Itt receives dynamicc pressure frrom the two pitot tubes mounted on either sidee of the verttical stabilize er. When the com mputer receiives an erraatic signal it’d be the same as th he pressure drop and the light illuminattes. (failed p probe heater and icing co onditions) 3. The third is relateed to the Stall Managem ment and Yaaw Damper (SMYD), ( and d a so called Elevator Feel Shiift module (EFS), ( which h creates a ±4 times higher h forwaard control column force when approaching the stalll region. This force uses a reduced syystem A presssure and wh hen this redu ucer fails, opening prematurelyy providing aa higher than n normal A system pressure to the feeel actuator, the FEEL DIFF PREESS also illum minates afterr 30 seconds. n the last syystem, it’s inhibited i <1 100 ft. RA and a AP selected, and w when the EFFS is not Note on operatio onal.
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Fuel Scaveng ge Jet Pu ump De fuel sscavenge jet pump scaveenges residuaal trapped fu uel from the center tank to tank #1. pped and Even at 0 Kgs indicaation there iss still some residual fuel in the centter tank. This fuel is trap cannot be b sucked up by the scaavenge line of o the cente er tank boosst pumps because of its elevated position. To be able to use this laast bit of fueel, a center ttank scavengge system is provided. To o activate the systeem, next co onditions neeed to be meet; the LEFT FWD pump operating aand tank #1 quantity lower th han half full. (< 1990 Kgss) When thee float type sshutoff valvee opens, it aallows LEFT FFWD fuel pump flo ow to create a negativee pressure in n the (non‐rrotating partts) eductor ttype scavenge pump which in n turn drawss fuel from the center taank relieving it in tank #1. Of coursee this will cre eate over time (the pump capacity is 100––200 Kgs/hr.. (AMM)) a rrelatively sm mall imbalancce between the main tanks. Th he book says that the syystem contin nues to run ffor the remaainder of thee flight (can’tt be shut off) but when you’ll remove thee controlled ccondition (LEEFT FWD fueel pump) also the jet pum mp stops ng. When thee center tank is depleted d, the scavenge pump d draws air from the cente er tank to operatin tank #1 w which obviously does no o harm to enggine #1 operration. Note: the “dissolved air” story off fuel. When on suction ffeed with a h high fuel tem mperature an nd a rapid decreasing pressure over the fuel, air bubbles (aeration) appear in tthe fuel possibly causingg engines to run errratic or even flame out w when sucked d up though the bypass vvalve. Note: when both ceenter tank fu uel pumps are a inoperative, fuel will be trapped d in the cen nter tank. There is no bypass vvalve provisio on for suctio on feed, also o the left main tank quan ntity has to b be below half full to even starrt the scaven nge jet pump p. Even so, th he scavenge rate is insuffficient to be e used for emptying the centerr tank. Under these cond ditions you’ll use main taank fuel befo ore the cente er tank is at requirred safe leveels and a possible overstrress of the w wing roots arises. (>453, th he main tankks have to bee full and >72 26, CONFIG)
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Fuel valves Let’s look at the most important valves in the fuel system, the Spar Fuel Valve and the Engine Fuel Valve a bit further than needed but still at an acceptable level. It will clarify what actually happens specifically with the Engine Valve. By all means just remember the easy way as the FCOM explains. The #1 most important fuel valve is the Spar Fuel Valve. This 28 VDC valve is mounted in the front wall “spar” of the main fuel tank supplying fuel to the fuel feed line of the engine. The DC power comes from the Hot Battery Bus and the valve even has an own recharging Battery Power Pack to be able to positively close the valve in case of an emergency such as a separated engine. The valve opens when the Start Lever is placed in the IDLE position and closes by CUTOFF of that Start Lever, or by pulling its Fire Switch. When the valve is closed it shows a dim blue light even with the Start Lever in CUTOFF as I always explain that any blue light is a “not standard flight condition light”, knowing that the book says it’s a status light. The Engine Fuel Valve is actually the High Pressure Shut Off Valve (HPSOV) and is integral with the Hydro Mechanical Unit (HMU) on the accessory gearbox. The valve opens and closes by the same controls as the Spar Fuel Valve but its actual opening is a bit more complicated. It relies on the so called Fuel Metering Valve (FMV) which is under control of the EEC. So, when conditions meet the requirements to open the HPSOV, the EEC signals the FMV to open up the HPSOV by servo fuel pressure. On the other hand the closing of the HPSOV is achieved by the Start Lever or Fire Switch, the EEC energizes the CLOSED SOLONOID of the HPSOV which uses 28VDC from the Battery Bus. During engine start this FMV is controlled by the EEC and when conditions dictate the HPSOV (Engine Fuel Valve) to close, the EEC commands the FMV and thereby the HPSOV to close in the following conditions: • A Hot Start occurs (>725°C) on the ground (exceedance protection) • If the engine decays after idle speed during start below 50% N2 speed and EGT exceeds the start limit • The EEC senses a “wet start” meaning no EGT rise within 15 seconds after the Start Lever is at Idle (YOU are the start limit for the EGT rise which is 10 seconds!!!) All of these conditions will be indicated by a bright ENG VALVE CLOSED light. Note that with an updated EEC software (7.B.Q and later) the EEC also provides a protection when approaching a Hot Start meaning a rapid increase in EGT. The 115/200 VAC, 400 Hz, 90 KVA Integrated Drive Generator.
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AC Generator I recently received a request from one of our followers to explain the operation of a brushless generator. I’ve send the explanation and thought on sharing this generic AC power generation info of an aircraft AC brushless generator. I’ve used the AC generator I’m familiar with and adjusted the image toward that generic explanation and added the 737 protection circuits in the GCU. The AC Generator is an assembly of three generators: • Permanent Magnet Generator (PMG) • Exciter Generator • Main Generator The most important Rotor components of the AC Generator are: • Permanent Magnet Generator rotor • Exciter Generator Rotor; which includes also the Rotating Rectifiers (3) and resistors (3) • Main Generator Rotor The most important Stator components of the AC Generator are: • PMG Stationary Armature; output: 39 VAC, 1 ø, 600 Hz • Exciter Generator Stationary Field; input: 28 VDC pulsating, 1,200 Hz • Main Generator Stationary Field; output: 115/200 VAC, 3 ø, 400 Hz Once the engine gearbox (N2) on which the generator has been installed has come on speed, voltage is excited in the PMG. This will be a 39 VAC, 600 Hz, 1 ø, at 100% revolutions of the IDG (± 12,000 RPM of the generator). This voltage is fed to the voltage regulator in the Generator Control Unit (GCU) through a DC Power Supply where it is converted into a pulsating direct voltage of 28 VDC, 1.200 Hz. The output of the voltage regulator is linked through the closed Generator Control Relay (GCR) to the Stator of the Exciter Generator which excites a 3 ø AC voltage in the Rotor. This AC voltage is than rectified by three rotating rectifiers and subsequently supplied to the Rotor of the Main Generator. The last step is that the Main Generator rotor field excites the required 115/200 VAC, 400 Hz, in the Main Generator Stator. The 115 VAC is the voltage taken from one phase and ground and the 200 VAC is the voltage between two phases (115 x √3) which explains the ra ng of what the generator can generate (115/200 VAC). The above shows that there is no need an external voltage source to ensure the generator is in operation, that’s why the system is also referred to as being "Self‐supported". OK the easy way is that the Permanent Magnet Generator (PMG) rotates by the IDG on the same shaft as the exciter‐, and Main rotors. The generated (39 VAC) is rectified to a pulsating DC in the control unit and send to the exciter stator. This DC power creates an alternate current in the exciter rotor and is rectified by the rotating rectifiers where after it finally creates an alternate current in the three main generator stator. This is the 115 VAC/400 Hz output of the generator and is monitored by the current transformers that relaxes or intensifies the DC power toward the exciter generator to the requested load of the electrical system.
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The in th he image shown protecttions in the CDU will de‐‐energize the GCR thereeby de‐energgizing the exciter field, f which h de‐energizes the geneerator. This de‐energizing GCR alsso occurs when w the generato or switch is sselected OFFF.
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Isolation vallve The isolaation valve seeparates thee left, from th he right side e of the bleed d manifold. It is powered d from AC Transsfer Bus 1 bu ut also can bee manually o opened/close ed by a control lever, acccessible in th he left air condition bay. Becau use it’s AC po ower* it will fail in the se elected position when po ower is remo oved. When th he Isolation sswitch is in th he AUTO possition the valve opening relies on thee so‐called “ccorner switch” positions. Th hey are the P Pack and Bleeed switches, when all theese switchess are NOT in tthe OFF position the isolation n valve is clo osed. On the other hand iif any cornerr switch is seelected to OFFF the Isolation n valve openss in the AUTO O selection. When a Pack switch is OFF, the Issolation valvve opens to ccreate equal performancce of the engines. o allow air frrom either siide of the maanifold When a Bleed is seleected OFF thee Isolation vaalve opens to to be useed for the offf side WTAI.. Note thee isolation vaalve logic is rrelated to sw witch position n so a tripped d pack or bleeed will not o open the Isolation n Valve when n in AUTO. Affter flight the Isolation valve should b be selected O OPEN just in case you need d to battery start engines when there is no APU o or external eelectrical pow wer available e. The ground aair connectio on is located on the rightt side of the m manifold clo ose to enginee #2. When N N2 >20% there is no personneel allowed in the vicinity o of the turnin ng engine so we have to sstart engine #1 first. When th his would be a battery staart you’ll neeed the isolation valve to be open, so when you re emoved AC poweer with the issolation valve switch OPEEN, the valve e is still in thee open posittion. * A geneeral rule for eelectrical pow wer is; “AC liies, DC dies”. This is a nice thing to o know also ffor analog in nstruments, aan AC powerred instrumeent stays whe ere it lost pow wer and a DC powered insstrument will drop off to o zero.
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Manu ual gearr extenssion. Let’s havve a look at tthis Non Norrmal procedu ure and its co omponents. he gear is UP P and the LG lever in the O OFF position n, hydraulic ssystem A preessure is removed When th from thee uplines to tthe actuators which causses the three e struts to “h hang” in theirr respective uplock. This is allso the prefeerred position of the LG leever during aa manual exttension attempt because e of the depressu urized hydraulic lines. he gear (all or any) does n not extend aafter a down selection, fo ollow the QR RH procedure e in an When th attempt to lower thee gear. Manu ual extension n of the gearr is accomplisshed by pulliing the three e “T” Access Door just behind the FO seat on the handles,, accessible tthrough the Manual Gear Extension A cockpit ffloor. d for this No on Normal prrocedure cou uld be caused d by: The need • Disrupted electrical signaal to the LG sselector valvve A hydraulic pressure available • No system A • LG lever stucck in the UP or OFF position When op pening the M Manual Gear Extension A Access Door, a “door open” micro swiitch comman nds the LG selecttor valve eleectrically dow wn regardless of the LG h handle position. This action activates the LG selector bypass valvee which conn nects the hyd draulic lines to return so the manual down selecttion does not hydrraulically restricts (locks) the actuators down cap pability. This also o prevents th he LG to retraact when thee door is not flush closed d after take‐o off and seleccted UP. This proccedure is covvered in the QRH by the LG disagree procedure w with the LG h handle UP an nd all red and greeen indicator lights illumin nated, tellingg you the geaar is down an nd locked bu ut not in the selected position. When yo ou’d pull anyy (or all) “T” h handle it sim mply releasess the uplock by cable actiion where affter the respectivve gear free‐‐falls down, ssupported by gravity (we eight) and airflow to the extend posittion. When th he gear is fully down, thee downlock “bungee” sprrings will hold d the downlo ock struts in an over centered d locked position. Normaally this is acccomplished by a downlock actuator b but with the absence of system m A pressuree, the springss enforce a m mechanical d downlock wh hich is indicatted by (6) do own and locked green lights. By the w way, there are 6 green ligghts as a redu undant indication. Neitheer gear is visible on the N NG and the doub ble green ligh hts for each strut will givve a backup ffor the down n indication.
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Mech hanical p pressurre relieff valves. There arre three mecchanical adjusted pressurre relief valves on the 737. Positive safety pressure relief is aaccomplisheed by 2 mech hanical adjustted pressuree relief valves, on each sidee of the outflow valve. Th hey are totally independeent of the prressurization system located o and prevvent the insid de/outside p pressure to eexceed +9.1 P PSID in the eevent of a preessurization system/o outflow valvve malfunctio on. (stuck clo osed outflow w valve) The fuseelage airfram me structure ccannot withsstand large n negative presssures and iss protected ffor that at a veryy low value. TThe negativee pressure reelief valve is llocated at th he right loweer side of the e fuselagee just fwd. of the outflow w valve. This sspring‐loaded door is also not depending on the pressurizzation system m and adjustted at just a –1.0 PSID vaalue. This will prevent thee aircraft to ccollapse when the inside/outside pressurre becomes n negative for example durring a (very) fast descentt.
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Nitro ogen Gen neratin ng System m Followin ng two Boeing 737CL explosion investtigations in A Asia (and oth hers includingg the B747 TTWA 800 midair explosion), a protection w was developeed by Boeingg to minimizee explosive vvapors in the center tank. Thee 737 explossions were caaused by trapped fuel higgh temperattures due to radiant heatt from the Packks under the tank which fformed highly explosive vapors. The fuel was ignited by the ccenter tank fuel pumps which were still running with an empty center tank. Early days center tank fu uel pumps d did not had aan automaticc shut off witth LOW PRESSSURE as thee newer mod dified ones th hat shut down aftter ±15 seco onds of LOW PRESSURE. TThis is also th he reason that someone has to be on n the flight deck when a ceenter tank pump is runniing as by the e FCOM, the book does n not cover exp plicit modificaations to each aircraft. This prottective devicce (NGS) divides Nitrogen n from Oxyge en by a separation modu ule and leave es Nitrogen n enriched aiir (NEA) in th he center tan nk to a level w which will no ot support co ombustion. TThe oxygen level is decreeased by the NGS to ±12% % which is su ufficient to p prevent ignitiion. The NGSS has only an indication aavailable in th he right main n wheel welll next to the APU fire con ntrol panel, so o it has no visible clew fo or crews of itts operation during flightt. ons are: Indicatio • OPERATIONA O AL (green) • DEGRADED ((blue) • INOPERATIV VE (amber) The nitro ogen generation system gets bleed aair from the left side of th he pneumatic manifold w where after its cooled, driveen through tthe separatio on module and directed tto a flow valve into the ccenter tank. Thee NGS operaates automattically only in n flight and sshuts down in the next co onditions: • Either enginee is shut dow wn in flight partment • Fire or smokke detection in any comp • Left Pack oveerheat
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Outfllow valv ve. To stay in line with the previous post, let us llook at this p pressurization componen nt of the 73. The outfflow valve reestricts/regullates the flow w of conditio oned air overrboard, thereeby creatingg a pressurizzed environm ment in the aaircraft. The valve is locaated at the afft lower sidee of the fuselage and has rakeed edges for noise reducttion purposees. The valve is moved b by a common n actuator w which can be operated byy either of the three outflow valve eleectro motorss. Two motorrs are operatted by the prressure systeem controlleers and one iss directly operated d by a switch h when in Maanual operattion. Automattic control is accomplisheed by meanss of 2 Cabin P Pressure Con ntrollers (CPC C’s) which alter control eeach flight orr when a maalfunction occurs on the o operating co ontroller. A th hird way of controlling the outflo ow valve is b by a manual ttoggle switch h on the pressurization p panel. The sw witch is spring lo oaded to neu utral and has three positions, CLOSE – – Neutral – O OPEN. The outfflow valve indicator show ws the actuall position of the outflow valve in all m modes of ope eration provided d the Batteryy Bus is poweered through h the PRESS C CONT IND C//B. Electricaal power to the three electro motors is provided b by: 1 electrical power to thee auto electrro motor 1 iss supplied byy the 28 VDC Bus 1 A • AUTO mode t through CPC C 1. (PRESS CONT AUTO 1 1 C/B) • AUTO mode A 2 electrical power to thee auto electrro motor 2 iss supplied byy the 28 VDC Bus 2 t through CPC C 2. (PRESS CONT AUTO 2 2 C/B) • MANUAL mo ode electricaal power to the manual e electro motor is supplied directly by tthe 28 V VDC Battery Bus. (PRESS CONT MAN C/B) A mode selector is used to deterrmine the op peration of th he outflow vaalve, either A AUTO, ALT(e ernate) or MAN((ual). The outfflow valve reeceives a clossed signal wh hen the cabin altitude reaches 14.500 feet in the AUTO mode off operation so it is not afffected throu ugh the MAN NUAL mode. de and a presssure loss, yo ou’d have to o close the ou utflow Just for tthe “mind seet” when at aa high altitud valve to increase preessure in the aircraft which results in lowering cabin altitude. Aircraft ccontrol overrride devices.
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Flight Control “Breakaway” Devices There are two devices that allow you to control the aircraft in case of a malfunctioning or jammed control system. One concerns roll control. When one of the yoke cables (or aileron PCU/spoilers) becomes jammed or moves freely, the opposite control is still available to roll the aircraft. The two yokes are interconnected at the base of the co‐pilots control column by the Aileron Transfer Mechanism through torsion spring friction and a “lost motion device”. If the FO control jams, the spring force can be overcome by the Captain thereby controlling the aileron PCU through cables. If the Captain control jams, the FO can control roll by use of the flight spoilers. Note that this only happens when the yoke has been turned ± 12° which engages a so called “lost motion device” which in turn operates the flight spoilers. The second is related to pitch control. When one of the control columns becomes jammed, the crew can override (breakout) the failing control. The control columns are interconnected below the cockpit floor by a torque tube with a device that enables the controls to be separated from each other. The Elevator Breakout Mechanism connects both control columns by two springs which will separate the columns when ± 30Lbf/13Kgf is used to overcome them. When applied, the control columns are mechanically separated from each other. Note that deflection of the elevators is significantly reduced and a higher force is needed to move the elevators. (even higher than with manual reversion)
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Pack & pack k contro ol There arre two Packs activated byy an AUTO/H HIGH selectio on that indiviidually has tw wo airflow direction ns, one that ggoes through h a three staage cooling cycle (2 air to o air heat excchangers and d an expansio on turbine) aand one that bypasses the cooling maachine and itts componen nts. The two flow direction ns are mixed at the output of the exp pansion turbine of the co ooling machine. Air that e enters the Packks through th he Pack Flow w Control and d Shutoff valvve is at ± 212 2°C and is co onditioned an nd cooled to a mixed minimum Packk output of ±± 18°C as set the lowest o on the zone temperature e control selectorss. (auto zonee temperaturre range is 18°C – 30°C)W When these sselectors aree all in the OFF position,, the left Pacck puts out a fixed 24°C aand the rightt Pack 18°C. There arre two combined Zone/Pack controlleers that conttrol the required output temperature e of each Pack. These two Pack Controllerss have an auto “on side”,, and a stand dby “off side” control, the latter takes ovver if an auto o controller faails. In this ccase a PACK OFF light illu uminates on recall togeth her with a Master Caution light. When bo oth Pack Conttrollers fail, aa Pack OFF liight illuminates with a M Master Caution light, the packs will still o operate until a temperatture exceedaance occur. When a Pack becomes overloadeed by the demand of coo ol air, a PACK K trip off lightt illuminatess with a Master C Caution lightt and the Pacck Flow Conttrol and Shuttoff valve clo oses shuttingg down that P Pack. When th he Pack coolss down and tthe light extinguishes, the Pack can b be reset by th he reset buttton on the Bleed panel. To p prevent this condition fro om re‐occurring select a higher temp perature to ““unload” that Pack by demand ding less cold d air from the cooling maachine bypasssing it. A Pack automaticallyy provides a high airflow when the otther Pack is sselected to O OFF provided d the aircraft iis in the air w with flaps up. The other cconditions re equire enginee performan nce and inhib bits the automattic high flow.. Note: the image is ju ust a simplifieed flow and pack component, and co ontroller imaage to illustraate the flow thro ough the pacck and the co omponents in both contrrollers.
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Recirrculatio on fans The recirrculation fan ns are located d under the cabin floor o on the forward cargo com mpartment’ss aft bulkhead d. The purpo ose of these ffans is to re‐‐use air draw wn from the ccabin and disstribution comparttment back into the mix manifold. Do oing so there e is no need ffor air from tthe Packs, th hereby relievingg the Packs frrom producing condition ned (cool) airr improving eengine perfo ormance. The e left recirculaation fan circculates air baack into the m mix manifold d from the diistribution co ompartmentt underneeath the cabin floor (mix manifold/fan area), the right recirculation fan fro om the passe enger comparttment. When a higher amou unt of fresh aair is needed d from the paacks, the recirculation fans are autom matically shut dow wn under sevveral conditions with thee recirculatio on fans selectted to AUTO O, and the iso olation valve sellected to AUTO or OPEN: On the gground usingg engine bleeed air: Left RECIRC FAN shuts down wheen both Packks are selecte ed to high flo ow On the gground usingg APU bleed aair: Left RECIRC FAN shuts down regaardless of Paack selection In flight using enginee bleed air: Left RECIRC FAN shuts down wheen either Pacck is selected d to high flow w Both REC CIRC FANS sh hut down wh hen both Paccks are selected to high fflow In flight using APU bleed air: Both REC CIRC FANS sh hut down reggardless of P Pack selection Reading the first parrt it makes seense that thee left fan (disstribution co ompartment)) shuts down n first as this areaa heats up byy the several operating components. (my personaal point of viiew)
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Hydrraulic Re eservoirs The 3 hyydraulic fluid reservoirs aare located in n the front of the main w wheel well. They are presssurized from thee bleed maniifold to supp ply positive fluid to the pu umps, preventing cavitattion and foam ming. The stan ndby system reservoir is p pressurized tthrough the B reservoir. These pressures (45 – 50 0 PSI) can onlyy be checked on 2 gages m mounted on the forward d main wheel well bulkheead. Quantityy of the A & B reservoirs is diisplayed direectly through h gages on th he reservoir b by a float typ pe transmitter which also send ds a signal to o the DEU’s ffor display on the lower DU. The stan ndby system reservoir on nly has a low quan ntity switch, which displaays the STAN NDBY HYD LO OW QUANTITTY light on th he flight conttrol panel wh hen < 50%. pipe to preseerve fluid to the EMDP w when a leak o occurs at the EDP. The A reeservoir has aa 20% standp The EDP is more likely to malfunction becausse of the enggine gearboxx mounted heeavy design and higher caapacity it puts out. (±4x) The B reservoir has aa common sttandpipe for both system m B pumps so o when a leaak occurs, fluid will drain thee entire B reservoir until a 0% indicattion. In this ccase the B syystem cannott be pressuriized anymoree but the rem maining 1.3 U USG can be u used for the PTU to operate the LE lifft devices. A second standpip pe at 72% preeserves fluid d to this level for both B ssystem pump p operation, in case a leaak occurs while using the stand dby hydraulic system. Minimum m quantity fo or the A & B reservoirs iss 76% which triggers a white RF (refill) indication on the lower DU U when on th he ground an nd TE flaps aare up, or no engines are operating. Besides tthat, when eequipped witth an updatee pin function n to the loweer DU on sysstems, there can also be a red dial indication when A o or B quantitiees decrease tto 0%, or inccreases to 10 06%. The pum mps heated (ccase drain) ccooling fluid return to the e reservoirs, is routed through oil‐to‐‐fuel heat excchangers mounted on thee bottom of the main tan nks. To achieeve enough ccooling for on the ground o operation, th here should b be at least 760 Kg of fuell in the tankss each.
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The A APU Sta arter/Ge enerato or. The APU U is started th hrough a starter/generattor and when n on speed transfers to aan AC generaator. The startt sequence o of the APU sttarter/generrator is deterrmined by th he Generatorr Control Uniit (GCU) which reeceives poweer from the SSwitched Hott Battery Buss. That is thee reason whyy the Batteryy Switch must be in the ON po osition (switched hot batttery bus ene ergized) to o operate the A APU. When sswitched OFF, thee Switched Hot Battery Bus and ECU b become de‐e energized wh hich in turn sshuts down tthe APU immediaately withoutt the regularr 1 minute co ooling cycle. (trips the geenerator off lline and closes the APU bleeed valve to u unload/cool tthe APU prio or shutdown) Strangely enough po ower to the sstarter is pro ovided by eith her the Batteery (28 VDC)), or Transferr Bus 1 (115 VAC C). Both voltages are first changed/boosted to a w whopping 27 70 VDC by th he Start Power Unit (SPU), w where after a Start Converter Unit (SC CU) creates the 270 VAC w which is neeeded to drive e the starter/ggenerator in the start mo ode. This sign nal lasts untiil 70% RPM w where the SP PU becomes de‐ energizeed and the AP PU becomess self‐sustaining and acce elerates furth her to its opeerating RPM. When th he APU RPM reaches ±95 5% the ECU ccommands th he blue APU GEN OFF BU US light to illu uminate as a sign nal that the A APU generato or can assum me the electrrical load. The AC ggenerator co onsists of thee same parts as the “regu ular” AC geneerator as desscribed in an earlier post and d can supply 90 KVA belo ow 32,000 feet and 66 KV VA at 41,000 because of A APU load cap pabilities with low w air densities.
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Land ding Gea ar Transsfer Valv ve The Land ding Gear Traansfer Valvee has two ways of operation. The simp plest is to traansfer the no ose wheel steeering operaation from itss normal hyd draulic system A, to the alterrnate hydrau ulic system B on the grou und (only), byy a switch on n the left front (Capt) pan nel. The seco ond way of o operation (in flight) is a bit more complex as it has 3 condition ns that needss to be met befo ore the LG trransfer valvee moves from m its normal hydraulic sysstem A operaation for geaar retractio on to the alteernate hydraaulic system B. N2 below 50% % 1. Engine #1, N 2. Landing Gear Handle in U UP A OT in the UP aand locked p position 3. Any gear NO The PSEU U is triggered d by those co onditions and moves the e LG transfer valve to system B. Note that the PSEU light is inhibiteed from T/O tthrust until 3 30 seconds aafter landing but DOES gu uard and ope erate the 737’s sysstems. Losing engine #1 stops the ED DP (hydraulicc system A) o output so thee only way to o pressurizze the A systtem is by meeans of the Electric Hydraaulic Pump w which puts ou ut 4 times less volume tthan the EDP P. This would d result in 4 ttimes slowerr movement of its compo onents includ ding a gear retrraction which is an unwaanted situatio on just after takeoff or o on a go‐aroun nd with N‐1 conditions when you u need to cleean that conffiguration as fast as possible to decreease the masssive drag by aany extended gear. In that case the rretraction is transferred ffrom the A, tto the B systtem so a normal ffast retractio on of the geaar is achieved d. The Pow wer Transfer Unit (PTU) iss a backup to o the LE lift d devices if the hydraulic syystem B EDP fails or has low output. It su upports the B B system elecctric hydraulic pump to o operate the llift devices in n a higher sp peed as it wo ould be 4 tim mes slower w with just the EMDP. The P PTU can also operate the e lift devices w when system m B fluid is lo ost to a 0% in ndication, still holding ± 1 1.3 USG resid dual fluid in tthe reservoir to be used by the PTU.
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PTU The PTU operates when the nextt conditions are met: 1. A Airborne and d, 2. S System B ED ow (< 2350 P PSI) and, P pressure lo 3. T TE flaps less than 15° butt not UP. If this occcurs the PTU U control valve opens, allowing system A pressure to operatee the PTU hyd draulic motor. TThe motor drrives a hydraaulic pump th hrough a com mmon shaft and uses thee 1.3 USG fro om below th he standpipee on the botttom of the B reservoir to operate thee selected liftt devices. Of course there aree return linees back to thee B reservoirr from the PTTU hydro mo otor and used d devices wh hich are not visib ble on common simplified d (FCOM) sch hematics. Note thaat the PTU do oes NOT tran nsfer fluid fro om A to B, and that the sselected devvices can be e extended AND retrracted by use of the PTU U but will opeerate according the used pumps. (EM MDP + PTU orr PTU only)
Teaser . . . .how CAN N you transfeer hydraulic ffluid from A→ →B or B→A??? A →B 1. EMDP's OFF. 2. Release parkking brakes, deplete accu umulator (<1 1800 PSI) 3. EMDP A, ON N and apply p parking brakees. 4. EMDP A, OFF and depresssurize by co ontrol column movementt. 5. EMDP B, ON and releasee parking brakes. (Sends tthe fluid bacck to system B) A →B 1. EMDP's ON. 2. EMDP B, OFFF and depresssurize by co ontrol column movementt. 3. EMDP A, ON N and apply p parking brakees. (Uses fluid from systeem A) 4. EMDP B, ON and releasee parking brakes. (Sends tthe fluid bacck to system B) B →A 1. EMDP's OFF 2. Either FLT CO ONTROL to SSTBY RUD. 3. No1 thrust reverser OUTT (uses stand dby hyd sys) 4. FLT CONTRO OL to ON. 5. EMDP A, ON N. 6. S Stow No 1 th hrust reverseer (using sys A)
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Wing g Therm mal Anti Ice (WT TAI) Wing anti‐ice is provvided for the inner three LE slats onlyy and is prefeerably used aas a DE‐icer. ANTI‐ ice would constantlyy heat up thee LE thereby melting the ice crystals immediately, creating waater “runbackk” over the w wing and posssibly freezin ng up on fligh ht controls. B Besides that it would havve a negativee effect on en ngine performance and ffuel consump ption. Note thaat use above FL 350 may cause a dual bleed trip o off by the req quest of the amount of aair also note thaat (ENG) anti‐‐ice is not reequired when n < ‐40°C SATT. not de‐iced b because the n narrow outer slat cannott hold the haardware need ded such The outeer slats are n as, a bleed manifold,, telescopic ttube and sprray tubes. Th he wing is acttually not pro oducing mucch lift in me ice accretion on that p part of the wing would no ot hurt that areaa anyway and they realizzed that som too much. Eventuallyy some drag and increase in stall spe eed occurs, not to forget that in case you use WTAI the stall warning computer remains set with increaased speed lo ogic. Where there is little cooling airflow over the LE on the grround, they aare protected against overheating. First the engine bleeed air is extrra cooled thrrough the pre‐cooler whiich allows tapped off fan air to o extra cool tthe engine b bleed air for m maximum LEE cooling on the ground. Second therre is an overheat sensor (± 1 125 °C) which h closes both h WTAI valve es when exceeeded and op pening up aggain at a predeterrmined valuee. During the design/teest phase it turned out th hat ice does not accumulate on the eempennage, mainly due to itts position in n relation to tthe engines causing hot air from the engines striking the emp pennage. Although h some ice can build up iin that area, it doesn’t haave any adveerse consequ uences (the sstabilizer regularlyy changes the AOA and eeventually sh hedding ice u under the new w conditionss). The milittary version of the Boein ng 737, the P8 Poseidon, does have a so‐called eleectro‐mechaanical expulsion de‐icing syystems (EMEEDS) installed d on the lead ding edges off the raked w wingtips, horizontal and verttical stabilizeers. The systeem is specially designed ffor the aggreessive slow aand low levell cold weatherr mission assignments of this aircraft and does baasically the saame as a de‐‐icing boot b but deformss the LE self b by using low electrical cu urrent (28VDC and 25 Am mps).
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B737 Yaw damping Airplanes with continued Dutch Roll tendencies usually are equipped with gyro stabilized yaw dampers. The Boeing 737 has two yaw dampers, a primary– and a standby yaw damper that keeps the airplane stable around the vertical axis when selected ON and with the respective hydraulic system pressurized through minimum SMYD generated rudder inputs. When engaged in NORMAL OPERATION, the primary yaw damper provides input to the main Rudder Power Control Unit (PCU) solenoid valve and is controlled by the Stall Management and Yaw Damper Computer 1 (SMYD 1). The input solenoid valve uses hydraulic system B to move the yaw damper actuator which ads in the mechanical rudder input. The yaw damper itself does not feedback motion back to the rudder pedals. The yaw damper input to rudder movement is limited to 2° with flaps up, and 3° with flaps down. To engage the primary yaw damper select: • Hydraulic system B ON, • FLT CONTROL B switch ON and • YAW DAMPER switch ON o Engage light extinguishes When engaged during MANUAL REVERSION, the standby yaw damper uses the standby Rudder PCU and is controlled by SMYD 2 which operates with standby hydraulic system pressure. During manual reversion the so‐called “Wheel To Rudder Interconnect System (WTRIS) supports standby rudder operation through SMYD 2 which receives an input signal from the Captains control wheel for coordinated turns during manual reversion. To engage WTRIS and standby yaw damping select: • Both FLT CONTROL switches OFF • At least one FLT CONTROL switch to STBY RUD • YAW DAMPER switch ON o Engage light extinguishes Both FLT CONTROL A and B switches must be OFF to enable SMYD 2, and one or both switches must then be in the STBY RUD position to provide standby hydraulic pressure. WTRIS only operates at < M 0.4 and yaw damper input to the standby rudder PCU movements are limited to 2° with flaps up, and 2.5° with flaps down. Both yaw damper systems are selected by a common “engage switch” on the Flight Control panel. When selected ON and the YAW DAMPER light extinguished, it only tells you the respective yaw damper is engaged regardless of operating by hydraulic pressure. During preflight the switch holds and the light extinguishes even without hydraulic system B pressure. The other way, if you’d lose system B pressure, the switch still holds with no light illuminated but primary yaw damping is lost. The switch only kicks OFF when the FLT CONTROL B switch is deselected from the ON position. To regain yaw damping you would have to transfer to manual reversion to operate the standby yaw damper with the standby hydraulic system which you (of course) will not do.
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Zone temperrature ccontrol Temperaature control is achieved d by mixing cool Pack air w with hot Pacck by‐pass air. The normaal temperaature range sselection is frrom 18°C – 3 30°C through h mixing cold d air from thee Packs with trim air for each individual co ompartmentt. The left Paack provides 20% to the C Control Cabin and 80% to o the mix manifold d where the right Pack prrovides 100% % to the mix manifold. Th he Zone/Pack controllerss hold the vario ous control eelectronics fo or the Cont C Cabin and Paassenger zones. The Contt Cabin has tw wo controlleers, a primarry and a backkup where th he Passengerr Cabin has o only one controller for eaach area from either Zone/Pack controllerr. (see previo ous image) Iff both Cont C Cabin contro ollers fail you u’d get a bin ZONE tem mp light with h a Master Caution, if one fails they illuminate on n recall. If a Cont Cab Passengeer Cabin Con ntroller fails tthe ZONE temperature liight and Masster Caution illuminates on recall and the two cabin reequirement w will be averaged. A ZONEE temperaturre light also iilluminate w when there is aan exceedan nce of duct teemperature,, the respective trim valvve will close w which can be e reset by the reeset button w when cooled d down. (seleect colder on n that area) In the no ormal mode the Packs prroduce a tem mperature acccording the selection of the lowest temperaature, the rem maining zonees use trim (hot) air requ uired for their selected teemperature. Unbalanced mode (C Control Cabin n trim air maalfunction) The left Pack producces the selectted Control C Cabin tempe erature and tthe right Pacck puts out th he lowest P Passenger Cabin selected d temperaturre, the Passe enger zone trrim valves stiill operate. Unbalanced averagee mode (any Passenger Cabin trim airr malfunction n) The left Pack producces the selectted temperaature but the e trim air valvve still operaates and the right Pack putts out an aveerage of both h Passenger C Cabin selecte ed temperattures. Single Paack operation and Trim O ON results in normal tem mp control, w with Trim swittch OFF all trrim valves close and the Pack averagges the threee compartme ent requirem ments. Trim swiitch OFF, all ttrim modulaating valves aare close and d the left pacck produces tthe selected Control Cabin temperature w where the rigght pack produces an ave erage of the Passenger C Cabin selecte ed temperaatures. output from the left and 18°C from th he right Packk. Temp seelectors OFF will create a fixed 24°C o
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Lavattory “firre prote ection”. I noted aalso B737 cabin crew “Likkes” to our FFB page, so I’’ll try to aim a couple of ssubjects in th hat direction n, of course aalso “need to o knows” forr flight crewss. Let’s start with Boein ng’s approacch of “fire pro otection”, off course we’re discussingg fire detection & extinguisshing NOT protection ;‐) The lavatory smoke d detection syystem needs 28 VDC from m DC Bus # 1 to operate. The lavatory is equip pped with a ssmoke detecction system and a fire exxtinguishing system. In so ome 73’s you still find a “SMO OKE” annunciator light at the P5 forward overhead panel but mostly there e is no indicatio on on the fligght deck. h the next co omponents: In the caabin we find smoke detecction indications through 1. Smoke Dete S ctor Unit As the name says, it’’s a smoke deetection and d the unit is m mounted agaainst the ceilling of the lavatory. It has a greeen (power) light and a rred (smoke d detected) ligh ht, also an alarm horn wiill sound whe en smoke iss evident for > 8 secondss. 2. Lavatory Call Light Located above the laavatory and iis a Call/Reseet Light that flashes amb ber when smoke is detectted. 3. Master Lavatory Call Ligh ht At each EXIT locator light there aare three indicator lights where a flasshing amber Master Call Light indicatess there is sm moke detected in the lavaatory in that respective area (fwd. or aft). A ontrol Panels (fwd.& aft)) 4. Attendant Co On thesee panels therre are more options than n just smoke detection as you can tesst the system m here and deteect FAULTS. W When smokee is detected d a red light fflashes togetther with a fllashing locatter light that iden ntifies the arrea where the smoke is d detected and d an intermittent horn is sound throu ugh the panel. Th he switches and lights on n the panel aare self‐explaanatory, wheen a FAULT iss detected during a test the failing detecctor is indicatted through the location n indicator. 5. Passenger Address (PA) ssystem The PA ssounds a repetitive high cchime when smoke is de etected.
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Center tank boost pumps There are two boost pumps located in the center tank that feed fuel into the engine supply fuel manifold at a rate of ± 10.000 Kgs per hour. The valves are mounted on either side of the crossfeed valve so with a closed crossfeed valve the pumps provide pressurized fuel to their respective side, the left center boost pump is than needed to supply positive fuel feed to the APU. Electrical power to operate the pumps are left, AC transfer Bus 1 and right, AC Transfer Bus 2. The design is such that there is no backflow possible through the pumps, meaning a check valve prevents fuel transfer through the engine feed manifold. These pumps also do not have a by‐pass valve which is needed for suction feed as with the main tank fuel pumps so, fuel in the center tank is trapped when both center tank pumps are OFF or producing no pressure. (the fuel scavenge jet pump (100 – 200 Kgs/hr.) is not considered a transfer flow) The center tank boost pumps are of a higher pressure then the main tank pumps thereby causing the center tank to empty first to prevent wing root stress when this would not be the case. The FCOM limit states that the wing tanks have to be full when there is more than 453 Kgs of fuel in the center tank. The second limit is related to that, i.e. when there is more than 453 Kgs in the center tank the boost pumps must be ON. I posted the C‐130 video where wing root stress caused the wings to shear off, the wing tanks were not full and the aircraft uploaded water and chemicals in a huge tank inside the aircraft every time to fight forest fires. About the same happens when the 453 limit is not honored with a possible exceedance of the MZFW. There are updated center tank boost pumps that automatically switch OFF when LOW PRESSURE (<22 PSI) is detected for >15 seconds. As these newer type pumps modifications are not covered by the FCOM the NOTE still exists to be on the flight deck when a center tank pump is operating. The 2 LOW PRESS lights on the fuel panel are extinguished when the pumps are OFF where the main tank pumps show LOW PRESS with their switch OFF. I call that “Recall Logic” as this would be a normal condition when the center tank is empty and the pumps OFF, preventing the MCS to illuminate FUEL at the Captain side Annunciator Panel (Recall) when pushed with the center tank empty and the switches selected OFF. The LOW PRESS circuit is checked when the pumps are selected ON for a short moment until the 22 PSI is reached.
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Antisskid The 73 iss equipped w with a system m that prevent wheels fro om skidding (deceleratin ng), thereby optimizing braking capabilities on any runwaay surface condition. An antiskid condition n releases brrake pressuree to the affected wheel(ss) which stop ps the skid co ondition when: • Uncommand d deceleratio on. (Antiskid protection) • Wheel(s) sto W ops instantan neously. (Loccked wheel p protection) • Landing with h (parking) brakes ON. (TTouchdown p protection) • Hydroplaning To detecct a wheel un ncommanded deceleratio on, an electrrical so‐called d transducerr is mounted d underneeath the hubccap of each w wheel and iss monitored by the Antiskid/Autobrake Control U Unit (AACU). This signal iss compared tto informatio on from both h Air Data Inertial Refere ence Unit’s (ADIRU’s) A and is also used for aauto brake syystem wheel speed functions. The AAC CU controls the anti‐skid system and monitors forr malfunction ns which aree indicated on the flight deck by an Anttiskid Inoperaative Light. A An additional signal to th he AACU com mes from the parking ns (releases) hydraulic flu uid through the brake syystem becausse the normaal antiskid syystem return parking brake valve. When the parking brakee valve has a disagree witth the lever (switch) the antiskid inoperattive light also o illuminatess. Antiskid is provided during operaating normall (system B), alternate (syystem A), an nd operation of the brakes w with residual accumulator pressure. W When in norm mal operatio on, antiskid iss provided th hrough 4 antiskid valves for eaach wheel seeparately and d during alte ernate or emergency (acccumulator) o operation through 2 antiskid vaalves whereb by the wheels are proteccted in pairs. To allow w retract brakkes to operatte (Alternatee brake presssure, system A) the antisskid system iss de‐ energizeed when the gear retracts. Be aware that the an ntiskid system m releases b brake pressurre, also durin ng emergenccy (accumulaator) operatio on which wou uld reduce eemergency brrake applicattions when sstepping on tthe brakes to oo hard.
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Leading Edge Flaps High‐lift devices on each wing are 2 LE Krueger Type Flap Panels and 4 LE Slats. The LE flaps have 1 extend position, Full Extend where the LE Slats have 2 positions, Extend and Full Extend, indicated on the aft overhead panel. On the center instrument panel just below the (TE) Flaps Indicator there is also an amber LE FLAPS TRANSIT, and a green LE FLAP EXTEND light. In NORMAL operation, the LE Flaps move by system B pressure to extend when the TE Flaps travel away from the UP position. They move in sequence after the TE Flaps travel to their selected position as commanded mechanically by a follow‐up cable system of the TE Flaps system. The extend time from UP to EXTEND takes ± 7 seconds and from EXTEND to UP ± 7.5 seconds. When the B system pressure is low, a so‐called priority valve gives operation priority to the LE Flaps over the TE Flaps. It reduces the flow rate to the TE Flaps, so the LE Flaps move relatively faster to their extend position. When the B system EDP pressure is low, the PTU supports LE Flap extend & retract movement. Refer for PTU operation elsewhere on this B737Theory FB page. In ALTERNATE operation, the LE Flaps uses standby hydraulic pressure and can only extend the LE Flaps. (Red guarded switch indicates an irreversible action) In this case the command is electrically through the Alternate Flap switches on the Flight Control Panel and the extend time from UP to EXTEND is ± 32 seconds. During cruise, pressure is removed from the LE Flap hydraulic system which creates a hydraulic lock of the LE Flaps. This prevents LE Flap extension at high speeds/altitudes which is accomplished by command of the Flaps and Slats Electronic Unit (FSEU). This condition exists when the next condition is met for >5 seconds: • Air born, • Flap Lever UP, • LE Flaps (and Slats) UP The LE uncommanded motion (UCM) detection function stops the LE normal operation if two or more LE flaps (or slats) move away from their commanded position. Different than the LE Slats, the LE Flaps do not have an internal actuator locking device so when residual system B pressure has leaked away during extended parking, the panels can droop off by their weight and gravity forces. This will de‐activate the Stall Warning Test capability. Rudder (vertical stabilizer) load reduction As on most large aircraft the vertical stabilizer is one of the most fragile structural parts. It cannot withstand large loads caused by full rudder deflection at higher speeds and therefore is protected against those high forces. The 737 rudder main PCU receives input from the pedals through input levers and a feel and centering unit which moves the rudder panel by hydraulic system A & B pressure. Pressures will be at normal values (± 3000 PSI) when flying < 137 Kts, above 137 Kts a load limiter reduces system A pressure to 1450 PSI resulting in a ± 25% reduction of the total load on the rudder. The result of this reduction protects the vertical stabilizer against high forces at a higher speed, leaving full pressure and deflection available when needed, at takeoffs and landings for directional control. An example of the vertical stabilizer “weak point” is an attempt in 2001 to recover an A300 after being struck by wake turbulence and aggressive maximum rudder inputs which sheared of the vertical stabilizer. Also note that the vertical stabilizer was the only intact part of the Air France 447 incident over the Atlantic. 37
In the paast of “my fieeld of experience” I saw a vertical staabilizer of a P P3 Orion totaally being sheared off like itt was removed with a chain saw wheen it struck a wash rack w when the airccraft has bee en swapped d around by a twister at NAS Jackson nville and wh hen a P3 hits a power cab ble at Pago Pago Hawaii. Be aware of the structural design n of your airccraft!!
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Thrust Reverser Each engine is equipped with a thrust reverser system that reduces stopping distance and brake disc wear. The T/R’s reverse the fan airflow forward through blocker doors, cascades and translating sleeves The left T/R uses hydraulic system A and the right, system B where they both are able to receive standby hydraulic pressure when their respective hydraulic system is unserviceable. Note that T/R use with standby pressure is of a lesser rate so, losing one main hydraulic system will operate that side slower than with main system pressure creating a possible swerve during reverser action. The T/R’s are controlled by the T/R levers on the thrust levers and operate when < 10 ft. RA or on the ground. The T/R operates when the thrust lever is at the Idle position and the T/R handle is lifted to the interlock position when the isolation valve positions to deploy the “translating sleeves”. The EEC’s determine through a Linear variable differential transformer (LVDT) a 60% opening of the two sleeves on each T/R, where after the mechanical interlock releases and the levers can be lifted further to the Detent 1, 2 or MAX position. When the sleeves move, the CDS shows the next message on the Upper DU. • Amber REV when deploying or stowing • Green REV when fully deployed When stowing the T/R’s, the stow command is initiated by passing the 1 Detent position which commands the T/R sleeves to stow. When the T/R lever is full down and the sleeves at the 0% (closed) position, the isolation valve closes and the locks engage. During normal operation the amber REVERSER light on the engine control panel illuminates for 10 seconds without a MASTER CAUTION during a T/R stow operation and extinguishes when the locks are engaged. The light will stay illuminated if the T/R does not stow in 10 seconds, indicating a malfunction. When the light illuminates for more than 12 seconds a malfunction is detected and the ENG annunciator and MASTER CAUTION light illuminates. When the down motion of the T/R levers is delayed for more than 18 seconds, the ENG annunciator and MASTER CAUTION light illuminate and the locks will engage, preventing further movement of the sleeves. To clear this situation you can cycle the levers to the interlock position and back down. When a serious malfunction or disagree exists between the LVDT’s, the ENGINE CONTROL light illuminates on the engine control panel together with a MASTER CAUTION. When illuminated, it could mean a serious engine (EEC) malfunction or an LVDT malfunction/disagree, when illuminated do not dispatch the aircraft.
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Each T/R R translating sleeve has two deactiva ation points, installing i two o pins at thesse points pre events T/R deplloyment. Follow the “thru ust reverser deactivation d for f flight disp patch proced dure” from the e current (AMM) manual m to operate the aircraft with deactivated T/R R’s.
An auto––restow circuit compares actual reveerser sleeve p position to the command ded position. When it determ mines an inco omplete stow wage or unco ommand mo ovement of tthe sleeves to the deployyed position,, the circuit ccommands to stow the TT/R. When acctivated, the isolation valve remains open and the control valvee is held in th he stowed position until the thrust reeverser is co ommanded to o deploy.
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Tail S Skid To proteect the aft low wer fuselagee from over rrotation dam mage the 737 7NG is equipp ped with a TTail Skid. It consistts of a sort‐o of shock abso orber cartrid dge, a skid fairing and a skid “shoe”, w where the last two parts aree outside thee fuselage wh here all otheer parts are inside. A light to ouch to the rrunway causes the shoe to wear off, indicated byy 4 dimples o on the shoe indicatin ng the amoun nt of wear an nd is an indiccation when the shoe neeeds to be replaced. he shoe hits aan object or an uneven p part of the ru unway duringg the skid, th he lower part of the When th shoe sheears off as on n the left imaage to indicaate a tail dragg but does not damage the skid fairin ng. A firm to ouch crushess the cartridgge pushing th he skid fairin ng inside the aft fuselage. The higher the force thee further thee skid disappears indicateed by colored d decals. If th he green deccal is still visiible the skid is sttill “serviceab ble” but if the green decaal disappears inside the ffuselage, thee red decal in ndicates that the skid must bee replaced. When th he “kiss to th he runway” iss more than firm, the skid disappearss totally insid de the aft fusselage and a safety pin (fuse pin) allowss the cartridgge to pivot in nside (other tthan crushin ng) thereby ng the aircraaft structure against massive loads. protectin There is also an optio on for a retraactable tail sskid that exte ends on takee offs and lan ndings which h is under control o of the Suppleemental Proxximity Sensin ng Electronicc Unit (SPSEU U) and operaates with hyd draulic system A A pressure. The SPSEEU command ds the tail skkid to extend if: • In the air forr 2 minutes aand, • Landing gearr lever is DN and, • Either enginee is running The SPSEEU command ds the tail skkid to retract if: • On the groun O nd 5 secondss or, • Landing gearr lever is nott DN or, • No engines o operating
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Vorte ex generrators. The 737 is equipped with several boundary laayer control devices or, vvortex generrators (VG’s). They aree mounted on the next lo ocations: • the wings t • the tail cone t e gine nacelle • the inner en t • the nose t • the APU inle t et door The vorttex created b by the VG afffects the bou undary layer on the respeective surfacce behind the e device by “pulling” air from outside, into the boundary layer. It creates an air swirl that draws air fro om above th he boundary layer into th his layer intensifying it an nd making it more compaact. VG’s are e mounted d to slow, co ontrol or even prevent bo oundary laye er separation n. VG’s aree used on thee 737 wings tto improve h high Mach pitch characteeristics beyon nd initial bufffet and to lowerr stall speedss in the landing configuraation. The (back swept) w wing design ccreates a relaative weak bo oundary layer where the outboard wiings are morre sensitive to initial flow w separation. The purpose of the wing VG’s is to strengthen thee boundary llayer (especiially with higgh AOA’s) and d direct the airflo ow on the su urface contro ols. On the tail cone, VG’’s are mountted to separaate the flow field from th he horizontal tail therebyy reducingg drag, impro oving perform mance and reducing elevvator vibratio ons. A Vortexx Control Devvice (or naceelle chine) is installed on the inboard side of the n nacelles. The e engines are mounted relativeely close to tthe wing whiich results in n air disturbance at high aangles of attack. To control tthe air flow aat high AOA’s and slow speed, a Boeiing invented VCD is mounted on the inner side of the engine naacelle. The crreated nacelle vortex is d delayed with h high AOA’s to support tthe airflow o over the wingg, increasingg lift in thosee conditions. There arre a number of VG’s mou unted on the nose of the aircraft just before the w windows. The general purpose is to o reduce airfflow noise byy ± 3 – 4 Dbss. on the fligh ht deck, direccting the airfflow om sharp edgges and corn ners of the w windows. away fro On the A APU inlet doo or, there is a VG installed d to improve e high altitude starting off the APU. W When the inlet doo or is opened during flightt, the VG imp proves inlet ram recovery and thereb by the pressu ure differencce across thee APU even tto assist (electrical) starting.
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Wind dow hea ating Window w heat is provvided to improve impact resistance (b bird hit), ice buildup and defogging aand should b be selected O ON at least 10 0 minutes beefore take‐offf. Each frontal window ((L & R SIDE & & FWD) have an own Window w Heat Control Unit (WH HCU) which rreceive poweer from theirr respective 1 115 VAC Transferr Bus. This heating is accom mplished by llaminated glass and vinyyl window layyers with in‐b between a conductive layer that allows elecctrical curren nt to flow thrrough it wheen heating is selected. Th his “graduall increasing” current flow w creates heaat by resistance in the layer towards a target temperaature of ±43°°C. The WHCU adjusts heeating curren nt in its operaating range tto prevent a thermal shock an nd reduces th he current flow at higherr ranges to p prevent an ovver temperatture. Window w heat becomess active when selected to o ON and thee window temperature iss < 37°C, indicated by a ggreen ON light (or extinguished d OFF light) m meaning theere is currentt flow througgh the condu uctive layer. W When dow reaches the target teemperature the WHCU interrupts eleectrical poweer and extingguishes the wind the ON light (or illum minates the O OFF light) wh hich could alrready be thee case when parked into the sun on a hott day. To makee sure the sysstem operates when neeeded, there is a POWER TTEST switch tthat by‐passe es the thermisttor and sends electrical ccurrent throu ugh the wind dows. Be awaare this actio on will bypasss the control u unit temperaature regulattion so when n activated to oo long it could cause an n overtemp ccondition in that w window. Anotherr test functio on is the OVH HT TEST switch that simu ulates an oveerheat condittion with the e window heat swiitches in ON,, indicated byy amber OVH HT lights on the control p panel and exxtinguished O ON lights (or illum minated OFF lights). The simulated overheat must be reset by selecting thee window he eat switchess to OFF than n back to ON N. When a window “overtemps” at values higheer than 62°C, the WHCU interrupts power to the affected windshieeld and illum minates an OV VERHEAT ind dication toge ether with an n extinguisheed ON light (or Illuminatted OFF lightt). This condition can be reset by sele ecting the affected windo ow to OFF an nd back to ON when allowed d sufficient co ooling first (2 2 – 5 minutes according tthe QRH). When w window heat is inoperativve prevent sp peeds above 250 Kts at altitudes belo ow 10.000 ft. to minimizee the effect w when a bird strike occurss on the window.
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Wing g& Body y Overhe eat On the P P5 overhead bleed panel,, there are 3 caution ligh hts mounted that warns yyou for an ovverheat condition in the bleeed system. Tw wo of them aare INSIDE th he manifold and considered “safe” ass they n after the overheat cond dition is corrrected, the B BLEED TRIP O OFF and PACK K light. have a reeset function The third d light is the WING‐BODYY OVERHEATT light that in ndicates an o over temperaature OUTSID DE the manifold d and is conssidered “not safe”. This in ndication determines an overheat in the area wh here the duct is lo ocated indicaating a duct leak or worsse, a duct rup pture. dication are: The areaas covered byy the left ind • Left engine sstrut (154°C)) • Left inner leaading edge (154°C) • Left air cond dition compartment (124 4°C) • Keel beam area (124°C) • APU bleed d uct area (124°C) A The rightt indication ccovers the engine strut, leading edge e and air con ndition on the right side. ondition (leaak) exists, use e the current QRH to dettermine the location When a wing & bodyy overheat co and isolaate the leak b by selecting a combinatio on of pneum matic system related swittches to OFF. When located aand isolated, the temperrature drops and extingu uishes the ind dication know wing that the overheat has disappeared but no ot the cause of the indicaation. When th he correct QR RH procedurre was follow wed the overhead conditiion should clear as the so ource has been n removed so omewhere d during the prrocedure. If n not . . . the Q QRH doesn’t suggests ste eps beyond tthe procedure so use common sensee to fly out of this conditiion. P5 bleed pan nel to test th he continuityy of the sensing The system has a tesst switch locaated on the P loops. Th he test startss when the b button is pusshed > 5 seco onds and indicates the saame as in an overheat condition b by the next aamber indications: W Y OVERHEAT • WING‐BODY • AIR COND an A nnunciator • MASTER CAU UTION
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Horizzontal S Stabilize er Trim m. One of the importan nt features reelated to pitcch and load b balancing is tthe movablee horizontal sstabilizer trim con ntrol (stab trim). A contro ol jackscrew moves the le eading edge of the horizo ontal stabilizzer as a trim in o order to achieeve this goall and can be operated: • Manually byy two trim wh heels which o operate the stabilizers geearbox and jjackscrew through c cables and c able drums. • Electrically eeither througgh yoke trim switches or Auto Pilot co ommand to tthe stabilizer e electrical trim m actuator. o AC p power – AC TTransfer Bus 2 o DC ccontrol – DC Bus 2 Electricaal movementt of the trim actuator by either the yo oke switch‐ o or the Auto P Pilot will bacckdrive the trim wheels on the control sttand. When the handle o on the wheells are extend ded during electrical operatio on, this can in njure the opeerators leg/kknee. 2.9°. Indicatio on in “Units”” is Extremee UP of the leeading edge is restricted at 4.2°, and DOWN at 12 mechaniically provideed on the co ontrol stand tthrough a fle exible cable tthat is driven n off the trim m control mechaniism underneeath the fligh ht deck floor.. As referencce, the 0° neu utral position n equals 4 un nits on the trim position scaale. Stabilizeer Trim Cutou ut switches aare located o on the contro ol stand in orrder to interrrupt either ccontrol column sswitch–, or A AP electrical power towaard the trim m motor when an uncomm manded move ement or “runawaay” trim occu urs. A Stabilizer Trim Oveerride switch h is located o on the aft ele ectronic paneel in case a counter move ement of the trim is required o opposite of tthe control ccolumn move ement. When not in OVEERRIDE, a me echanical control ccolumn actuaated stabilizeer trim cutou ut switch willl interrupt electric poweer to the trim m motor when atttempting to trim oppositte of control column or A AP command ded force. (co olumn nose DOWN vs. trim UP or vv) The overrride switch can also be u used to by‐p pass the conttrol column aactuated stab bilizer trim ccutout switchess in the event both (yokee switch or AP) fail in the open positio on, to be able to operate e the stab trim m. Electricaal movementt by the yokee switches caan vary betw ween high speeed (0.4 unitt/sec) when tthe flaps are NOTT UP and low speed (0.2 u unit/sec) when the flaps are fully UP.. When the trim is under AP control h high speed iss 0.27 unit/seec while low w speed with the flaps UP P is 0.09 unit//sec.
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Displlay Elecctronic U Units. There arre two DEU’ss in the Comm mon Display System (CDSS) which are located in th he E & E bayy that receive d data from the B737’s airccraft systems and avioniccs. This data is converted d into a video o signal that is seend to the six Display Units (DU’s) on n the flight de eck. Both DEEU’s provide data toward d all six DU’s butt are differen ntly powered d. DEU 1 is powered thro ough the 28 V VDC Standbyy Bus and DEEU 2 through the 28 VDC Bus 2. They both have a “hold‐up po ower” from the Hot Batteery Bus which is used to supplyy power to the DEU’s during power ssurges of maaximum 2 secconds or elsee the DEU po owers down. Both DEU’s “ccrosstalk” to compare criitical data an nd when therre is a differeence, this could create an amber CDSS FAULT indication as desscribed below. The same “split” is m made for pow wering the co omponents tthat distinguish DU operaation when p powering up the aircraft. I meaan when the Battery Switch is selecte ed ON, DEU 1 is powered d through the DC Standby Bus but also o both Captains–, and Up pper DU’s, ass well as the Captains EFIIS control panel. Note thaat it takes ± 9 90 seconds to get displayy because the DEU has to o become op perational. W When the DC Bus 2 2 becomes powered the same applies for the Firsst Officers sid de. An (undiispatchable) amber CDS FAULT displaays on both P PFD’s when tthere is a DEEU operation nal failure on the ground and o one engine operating. Wh hen both enggines are opeerating or when in the aiir the ULT message is replaced b by a DISPLAYY SOURCE me essage. The DISPLAY SOU URCE also sh hows CDS FAU when on ne DEU is selected (ALL O ON 1 or 2) to provide all ssix DU’s with data. Note: Sw witching too fast between SOURCE seelections can n create a po ossible incorrrect data disp play, use a 1 – 2 second intervval when switching between the SOU URCES. A (dispattchable) white CDS MAIN NT indication n tells us thatt there is a p partial data in nput malfunction on a DEU w when on the gground and o one engines is not operating.
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Proximity Switch Electronic Unit The PSEU is located on the right side in the E & E bay and receives input from the six strut compression sensors (2 on each strut). These ground/air signals are used by the PSEU for several aircraft systems and/or indications such as: • Landing gear transfer valve • Landing gear position indicating and warning • Speedbrake deployed indication • Takeoff warning • Door warning • Air/ground relays • See image . . . The PSEU also serves as a FAULT detection regarding several aircraft systems when on the ground and the thrust levers <53°, or after landing when on the ground for >30 seconds and the thrust levers <53°. Generally speaking, the PSEU light is inhibited in flight but it does monitor systems and records any FAULT to be annunciated 30 seconds after landing. When a system status FAULT is detected or an overwing exit flight lock fails before take‐off, the PSEU light illuminates together with the OVERHEAD annunciator and a MASTER CAUTION light. An undispatchable FAULT is evident when the PSEU light illuminates after landing when on the ground >30 seconds and the thrust levers <53°, in this case the light can only be reset by a BITE check of the PSEU or when the FAULT is corrected by maintenance. A dispatchable FAULT exists when the PSEU light illuminates after landing when on the ground for longer than 30 seconds and the thrust levers <53° and the light extinguishes when the parking brake is set or the engines are shutdown. A dispatch fault will not cause a recall of the Master Caution annunciator light but just illuminates the PSEU indication. A dispatchable (simple) FAULT occurs if the PSEU light illuminates when pressing RECALL and resets by pressing MASTER CAUTION. The SPSEU light or Supplemental Proximity Sensor Electronic Unit is provided that uses the Landing Gear DOWN signal to extend the two position tail skid and/or to determine a failure of the flight locks on the additional 2 mid exits. (– 900’s) Note that some indications are only valid for certain models/serials.
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Nose wheel ssteering g lockou ut Yesterdaay I answered a question n to one of ou ur followers and want to o share this (o obvious) info o. owing or pushing the airccraft with a ttug and tow bar, system A A pressure h has to be rem moved When to from thee nose wheel steering sysstem. It prevvents unwan nted dangero ous movemeents of the to ow bar injuring people or daamaging the nose wheel steering systtem. It is donee either by switching botth A pumps tto OFF or by use of the to owing lever aand lockout pin to depressu urize the nosse wheel steering system m. Moving tthe lever to the towing (fwd.) positio on moves the e towing shu utoff valve to o such a posittion that it shuts o off A hydraulic pressure tto the steering valve and d connects both sides of the steering valve to each oth her, and retu urn. This allow ws the nose gear to rotate freely to aa maximum o of 78° indicaated by red strip pes on the low wer side of tthe fuselage.. Moving the e nose wheel beyond the stripes requ uires disconneecting the to orsion links or even the taaxi light wirin ng. Most companies wants you to always switch the A pump ps OFF even iif the lockout pin is installed as a safety m measure. Thiss is done to p prevent misccommunication with the multi‐cultural ground ob bservers related tto pin installaation. nking about tthe system as an engineeer . . . . . reme ember the N NOSE WHEELL STEERING sw witch. Just thin When it is selected in Alternate o on the groun nd, it moves the Landing Gear Transfer valve to select B system p pressure to n nose wheel ssteering!!!
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Weather radar The on‐board weather radar can provide the following information: • Weather • Windshear • Terrain The WXR switch on either EFIS panel activates the weather radar and displays the weather radar data on the ND in the MAP, MAP CTR, VOR & APP modes. (not in the plan mode) The radar covers 180° in front of the aircraft by receiving transmitted radio frequency echo pulses on the ND’s. When selected on the EFIS control panel in a correct display mode, the DEU’s send an analog discrete to the weather radar control panel which sends it to the weather radar transceiver and switches it ON. When the aircraft is equipped with a predictive wind shear system (PWS), it’ll be available below 2300ft. The weather radar does not need to be switched ON for the PWS to work, it switches ON automatically when take‐off thrust (PL > 53°) is set. PWS information is available after the WXR switch on either EFIS control panel is pushed and a 12 sec warm up period, where after Alerts become available. Alert activation regions for TAKE‐OFF are: • Warnings and Cautions are enabled from 0 knots until the aircraft reaches 80 knots. • From 80 knots until the aircraft passes 400 feet, only Warnings are enabled. • From 400 feet through 1,200 feet, Warnings and Cautions are enabled. • All alerts are disabled from the time the aircraft passes 100 knots until it reaches 50 feet. Alert activation regions for APPROACH are: • PWS switches automatically ON when the airplane descends below 2300 feet RA. • PWS switches automatically OFF when one of the next conditions occur: o Aircraft speed is less than 60 knots. o Aircraft climbs above 2300 feet RA. If PWS is ON and WXR is not selected on the EFIS panel, all antenna sweeps search for wind shear. If WXR is selected, the antenna uses one sweep to search for wind shear and the other sweep to search for normal weather returns. PWS operation does not affect the WXR mode or range selected by the flight crew. Alert activation regions for LANDINGare: • Warnings and Cautions are enabled from the time the aircraft passes 1,200 feet until 400 feet. • From 400 feet until 50 feet, only Warnings are enabled. • From 50 feet until touchdown (0 feet), all alerts are disabled. • No display Wind shear alerts are active in the cockpit below 1,200 feet AGL.
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The weaather radar actually enterrs the wind sshear scanning mode at 2 2,300 feet AG GL to provide time for the ssystem to power up (if neecessary) and d update the e displays before the airccraft reachess the 1,200 feet AGL level.. TEST During the test: • The R/T tran T nsmits a few pulses to lett the BITE mo onitor for correct operation • The R/T mak T kes a test patttern and sen nds it to the DEU to show w on the ND’’s • The R/T send T ds test messages, mode, gain and tiltt information n to the DEUs to show on n the ND’s • WXR test pa W ttern shows on ND’s. • The test patt T tern shows u until anotherr mode on th he WXR paneel or EFIS pan nel is selecte ed.
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Disso olved air A fuel ph henomena w within high alltitude aviatiion is called ““dissolved aiir” or “aeratiion”. on of fuel cau used by rapid dly decreasin ng tank presssure It is a ressult from thee highly aerated conditio during climb, allowin ng entrapped d air in the fu uel to expand d. Reduced aair pressure above the fu uel surface p promotes the release of dissolved airr from the fu uel. Air releassed from thee fuel can have degradin ng effects on n the perform mance and saafe operation n of a fuel syystem. mp LOW PREESSURE’s durring rapid clim mbout can cause thrust d deterioration n or 2 Main ttank fuel pum even an engine flameout on the affected enggine at highe er altitudes (>> ± 13,000 feeet). UEL PUMP LO OW PRESSUR RE) (QRH, FU This altittude varies w with the prevvailing fuel teemperature in the tank (the higher th he fuel temp perature, the loweer the altitud de at which the gradual p power loss occcurs). Once preessure has sttabilized and d excess air h has escaped ffrom the fueel, loss of botth fuel boostt pumps has no eeffect on enggine operatio on with maximum power settings at aaltitudes up tto above 30,,000 feet. Thee time requirred to stabiliize the fuel from this high hly aerated ccondition cannot be dete ermined exactly, since it is a ffunction of b both rate‐of‐cclimb and fuel temperatu ure. Solution to the problem is level o off and let th he engines sttabilize at alttitude or pressure feed th he engines as suction feeed increases the aeratio on effect and d ads in the p possibility to ingest aeratted fuel into the engine feed line. Fuel stab bilization sho ould occur affter a few minutes of stabilized cruisiing operation or back on n pressuree feed.
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Frang gible fitttings Frangiblee fittings aree mounted in n the rim of tthe main wheel wells to p prevent a rotating blown n tire to enter the wheel welll. If it shears,, only that side will freefall back dow wn by relievin ng Landing Gear Actuatorr up pressuree overboard. (4 green an nd two red in ndications) Note thaat we do havve retract braakes through h the alternate brake systtem (hydrau ulic system A) but when a ttire blows th here is a good d chance thaat the brake lines will be substantiallyy damaged causing the retraact brakes no ot to work. Retract b brakes (and nose wheel ssnubbers) arre mounted tto stop the w wheels from rotating, hanging in their upllocks. A high speed rotatting wheel w would cause ttremendous precession fforces to the structuree during a tu urn that’s wh hy they are sttopped afterr retraction.
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Rudd der(verttical sta abilizer)) load reductio on As on mo ost large airccraft the verttical stabilizeer is one of the most fraggile structuraal parts. It caannot withstan nd large load ds caused by full rudder d deflection at higher speeds and thereefore is prote ected against tthose high fo orces. The 73 37 rudder maain PCU rece eives input frrom the pedaals through input levers an nd a feel and d centering u unit which moves the rud dder panel byy hydraulic ssystem A & B B pressuree. Pressures w will be at normal values (± 3000 PSI) when flying < 137 Kts, above 137 Ktss a load limiter reeduces systeem A pressurre to 1450 PSSI resulting in n a ± 25% reduction of th he total load on the rudder. TThe result off this reductiion protects the vertical stabilizer agaainst high fo orces at a higher speed, leeaving full prressure and d deflection avvailable when needed, att takeoffs an nd landings fo or direction nal control. An exam mple of the veertical stabilizer “weak p point” is an attempt in 20 001 to recoveer an A300 aafter being strruck by wakee turbulencee and aggresssive maximum rudder inp puts which ssheared of th he vertical sstabilizer. Alsso note that the vertical stabilizer waas the only in ntact part off the Air Fran nce 447 incident over the Atllantic. In the paast of “my fieeld of experience” I saw a vertical staabilizer of a P P3 Orion totaally being sheared off like itt was removed with a chain saw wheen it struck a wash rack w when the airccraft has bee en swapped d around by a twister at NAS Jackson nville and wh hen a P3 hits a power cab ble at Pago Pago Hawaii. Be aware of the structural design n of your airccraft!!
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Rejeccted Tak keoff – sspeed b brakes rrelation. Let me sstart to tell th hat the speed brakes alw ways refer to extension of both, groun nd and flightt spoilers. As we all kn now, when laanding, the gground spoile ers are triggeered by the rright main strut compresssion and thee flight spoileers when anyy strut comp presses. During taakeoff, the sspeed brakess do not need d to be arme ed but actually they are . . . . by a so ccalled "speed b brake refused d take off (RTO) switch". The switch is activated w when you reeject the take eoff and lift the th hrust reverseer levers up with the thrust levers at IDLE. In turn n they’ll activvate the RTO O switch by a reveerser cam which will dep ploy the speeed brakes thrrough an autto speed brake actuator. When acccording the QRH the speeed brake lever is raised by the Captaain, the RTO O switch (auto o speed brake acctuator) doessn't operate anymore bu ut the speed brakes are n now manually deployed. Everybody can help me out with triggering su ubjects as Sh hrikant does,, and I invite anybody to do so. Also exp periences “ou ut the field” improves the knowledge e of our FB page followerrs.
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Electtrical Bu us (bar)) We all have seen circcuit breaker panels and kknow there aare Electrical Busses behind them bu ut what are theyy? A Bus Baar functions aas a central tterminal in the aircraft electrical systtem to conneect main elecctrical system p power to varrious electriccal components. It simplifies the wirin ng system an nd provides aa common n point from which voltagge can be disstributed thrroughout thee system. Alsso using Bus Bars(com mmonly located in the flight station o or galley) savves weight ass the copperr wiring is takken from that poin nt in the airccraft instead of from the source for each electricaal componen nt, which would require aa multitude o of distant wiiring . . . .and d weight. A Bus co onsists comm monly out of high capacity Copper strrips to which the several users are co onnected. The Bussses are poweer fed from ttheir electrical sources (G Generator, Exxternal cart, Battery, TR or other Busses) tthrough currrent limiters that protectt them and itts attached ssystems for h high currentss that in turn cou uld cause a fire hazard orr damage equ uipment. Fro om the Bus ittself individu ual electrical components are con nnected through a thermal protection n, a circuit breaker, whicch vary in cap pacity visible on top of them m. The bus componentssidentify in D DC, and AC busses as they differ in th he amount off copper strip ps. The DC bus b bar is formed d by two strip ps, the +, and the – wherre the latter connects to the aircraft structuree or Ground Bus. The AC C bus bar is fo ormed by thrree Copper sstrips, 3 AC p phases which h of course aare isolated b by Teflon divviders. On top p of most Bu usses there iss an isolation n strip mounted to prevent a short to th he respectivee Bus would something cconductive sttrikes it.
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Crew w oxygen n system m First of aall I’ll tell you u that oxygen n is dangerou us!! It canno ot mix with grease in anyw way causing an explosion. That is thee reason when maintaineers charge O O², they wearr white coverralls to see iff there is any greaase visible on n their clothing. O yeah so . . . do NOTT use lip chap sticks wheen using O²!! A green oxygen cylin nder is moun nted in the E& &E bay and ccan be eitherr 114/115 ft³³ or 76 ft³. Th he minimum m charge is ccompany relaated but mostly a rounde ed up value regarding the highest temperaature and maaximum fligh ht deck occup pation (see yyour PI sectio on in FCOM 1 1). Maximum m allowed indicated prressure is 185 50 PSI but th he overboard d discharge vvalve is set att a whoppingg 2600 ndicator discc located justt behind the E&E bay acccess blows‐o out either by PSI weree the green in overcharrging or by thermal expaansion. ure, one direcct reading gaage on the bo ottle and thee second is o on the aft There arre 2 indicatorrs for pressu overhead panel whicch gets an electrical signaal from a pick‐up in the m manifold and d is powered by the Battery B Bus. From thee bottle the pressure is rreduced to ± 60 – 85 PSI and has a prrotection at 1 100 PSI to prrevent a too high pressure toward the reggulators. The crew w regulators have the next options: Normal –– in this posiition you havve to inhale tto get a dilutted oxygen flow meaningg it is a mix o of environm mental (cockkpit) air and o oxygen. (no w wise during fumes/smokke) 100% ‐ yyou’ll inhale 100% oxygen n on demand d. EMERGEENCY – in thiss position yo ou get pressu urized 100% oxygen through the massk. There is a test lever o on each regu ulator to testt the oxygen system sepaarately. The first test is ju ust to slide thee RESET/TESTT lever backw wards were it releases re esidual pressure in the mask with a sh hort rush of air and indicating the yellow X X flow indicattor. The seco ond test need ds to be perfformed in th he EMERGENCY position o of the regulaator were yo ou first have to iidentify the crew pressure (in the maanifold) than n push the bu utton to test and slide simultan neously the R RESET/TEST lever backwaards for 5 secconds. This sshould resultt in a constan nt flow of air witth the X flow w indication. There should be no morre decrease tthan 100 PSI,, a sharp dro op‐off or slow increase of presssure on thee indicator. W When any of the previouss occurs the valve on the e bottle is either closed or not completely o open. After tthis don’t forrget to rotatee the test bu utton back to o 100%. The test can be perfo ormed togetther with thee mask micro ophone test ((SP 1 in FCOM M 1) through h selectingg MASK, FLT INT and SPK KR on the aud dio elect pan nel so you’ll h hear the rush h of air throu ugh the cockpit sspeaker wheen simultaneously pushin ng the INT sw witch during tthe O² test. Not all companies perform thiss test but rememb ber it is emeergency equiipment so it can saafe your day during No on‐Normal procedures as you h have to estab blish crew commun nication som mewhere!
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Main n system m hydrau ulic pum mps, (co orrected d) The 737NG is equipp ped with two o Engine Drivven Pumps (EEDP) mounteed on the geaarboxes of th he engines, and two Eleectric Motor Driven Pumps (EMDP) w which are located in the m main wheel w well. The biggest N NEED TO KNO OW between n the EDP an nd the EMDP P is that the EEDP’s deliverr hydraulic fluid at a ± 6 times higher capacity than th he EMDP’s. TThis means th hat when an EDP does no ot turn, everything operated d through thaat hydraulic ssystem (A orr B) travels 6 times sloweer!!! with jusst the EMD DP. Boeing eq quipped certtain hydraulic operated ssystems with h a back‐up to improve th heir operatin ng speed as n needed such as the PTU aand the LGTU U. The hydrraulic pumpss are axial vaariable displaacement pum mps that varyy the demand by anglingg a so called sw wash plate. W When little demand need ded, the plate angle is low w and when a high demaand is needed it is high delivering high volume of fluid. Pressure es supplied b by the pumps depend on n the demand where a nom minal 3000 P PSI exists witth a 2800 minimum as peer FCOM and d a 3450 PSI maximum. Pressuree control is accomplished d by a pressu ure module that receives fluid from th he EDP and EEMDP where affter it supplies the respeective hydrau ulic system, it also holds tthe overpresssure protection. Selectingg the EMDP tto OFF, shuts off AC elecctrical powerr to the pump p, AC XFR Bu us 1 power fo or the system B B EMDP, and d AC XFR Bus 2 power forr the system A EMDP. This X‐powerin ng the EMDP’s prevent a complete main hydrau ulic system lo oss in the eve ent of an enggine failure ttogether with a bus transfer problem maaking the system more reedundant. The EMD DP’s have an overheat indication if th he pump ove erheats, it deepends on tyype whether it is the cooling ffluid or the eelectric moto or that overh heats. Also it depends on type if the m motor shuts down automattically, or only the light illuminates w when an overheat exists rrequiring crew action to sshut that overheated pump do own. pressurizatio on solenoid vvalve closes, blocking thee output presssure to With thee EDP switch in OFF a dep the respective system m. This solen noid remainss energized in the OFF po osition degraading the life etime of that soleenoid giving us the reaso on to leave th he EDP switch in the ON position until closing beccomes necessarry. When th he fire switch h is pulled, th he fluid supp ply shutoff vaalve to the reespective ED DP is closed and the LOW PRESSURE indiccation is disaarmed. Hydrauliic pump outputs: 36 GPM M EDP EMDP 5.7 GPM M STDB 3.7 GPM M PTU 11.6 GP PM
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Cockp pit Voicce Recorrder Sysstem The Voicce Recorder Unit (VRU) iss water, shocck, and heatp proof and is located in th he aft cargo comparttment. The C CVR controls are on the fw wd overhead d panel. The VRU U receives po ower from th he AC XFR Bus 2 (VOICE R RCDR on P18‐‐2) and is equipped with an Underwaater Locator Device whicch activates .. . . submerged in water tto a depth of ± 20.000 ft. The 37.5 khzz, one second d pulse tone is detectablee for > 30 daays within a rrange of maxximum ± 3.5 Km. The CVR R records aud dio from the three ACP’s,, and cockpitt area on a 4 channel solid state conttinuous loop datta tape for 12 20 seconds, w where after the tape eraases the first recording an nd stores the e current audio. There co ould (depend ding on type)) also be provisions to sto ore ACARS data link messsages or clocck inputs from either the Capttains, or Firstt Officers clo ock on the CV VR tape. The cockkpit voice reccorder contrrol switch is located on th he overhead panel and h has the next functions: • Controls VRU C U switching p power which h comes from m DC Bus 2 (V VOICE RCDR RELAY on P1 18‐2). • ON, O the CVR R receives power for maaintenance or o pre‐flight testing,the switch automatically positions to AUTO when either engin ne reaches id dle RPM. • AUTO, the C A CVR receives power wheen either enggine reachess idle RPM and remains powered until 5 minuttes after thee last engine has been shut down. ere is AC On certaain 737’s the CVR becomes activated any time the power on the aircrafft so there is no control sswitch on the e ovhd panel. (cc/b in) On the C CVR recorderr panel are lo ocated: • Area microp A hone • Channel mon C nitor indicator and/or staatus indicato or light • Test button T o Creaates a test to one toward the 4 channels (Captain, FFirst Officer, Observer an nd area) o Thesse tones are indicated byy a deflection into the grreen area of the channell monitor indiccator. Can bee heard when plugged in n to the head dset jack. o Wheen a fault is detected, the t audio to one stops an nd the indicaator stays in n the red areaa, or extingu uishes the status light. When W no fau ults are deteected the staatus light illum minates mom mentarily. • Erase button n o Can only be used d when on th he ground (P PSEU determines that) with the parking brakke. o Activvates when h holding the b button for > 2 seconds (aalso can be >> 5 seconds) • Headset jackk o For ttest tone transmission or recording p playback. The CVR R circuit breaker should b be in all the time and can be pulled (ccompany policy) when tim me allows affter an emerrgency evacu uation or when the Captaain deems th his necessaryy when he waants to save valu uable inform mation of thee last 120 min nutes prior p pulling the c//b after flightt. 58
Presssure con ntrol To survivve at altitudee the 737 is eequipped witth an autom matic altitude control systtem by mean ns of pumpingg air in, and rrestricting air out. There are two iden ntical digital Cabin Pressure Controlle ers (CPC) that alteer each flightt, and back each other up p in the even nt of a failingg CPC. Pressurizzation start w with an input on the oveerhead pressurization con ntrol panel o of a cruise–, aand landing aaltitude wheere after all p pressure events happen aautomatically. This startss by moving the throttless up until N1 on both enggines reaches 60% > 1.5 sseconds, or N N2 reaches 8 89% >1.5 secconds. At that mom ment the outflow valve m moves towarrds close, briinging the caabin altitude to ± 200 ft b below field elevvation by raising the differential presssure to 0.1 P PSID. This prrevents unco omfortable pressure surges w when rotation n creates a n negative presssure outside e of the outfflow valve byy its position. mb mode, inccreasing the diff/press to oward the firrst limit After liftt off the conttrol changes into the clim of 7,45 P PSID at 28.00 00 ft keepingg the cabin att the departu ure field elevvation until aapproximate ely 18.500 ft. After climb bing through h 28.000 ft th he diff/presss increases to o 7.80 PSID u until 37.000 fft where after it in ncreases to tthe maximum m automaticc limit, 8.35 P PSID. The AM MM also men ntions a maximum diff/presss when deviiating from aaltitude of 8.45 PSID but this is not byy FCOM wheere the maxim mum cabin alttitude is 8.00 00 ft.When the cabin altittude exceeds 10.000 ft, aa cabin ALTITTUDE ALERT sounds togetherr with the illu umination off Cabin Altitu ude Warningg lights on bo oth forward p panels. (whe en installed d) At a cabin altitude of 1 14.500 ft the outflow valvve receives aa close signall overriding tthe Cabin Pressuree Controller. 0.1 PSID) is 3 350 fpm and cabin ROC d during the cliimb is The maxximum cabin ROD during takeoff (to 0 600 fpm (can be 750 0 fpm). When ap pproaching tthe set FLT A ALT by 500 ft (0.25 PSI), th he pressure controller en nters the cru uise mode off operation aand maintain ns a constantt cabin altitude. The otheer way aroun nd, when desscending further tthan 0.25 PSI from the seet FLT ALT, th he controllerr changes intto the descen nd mode of operatio on and pressu urizes the caabin with 350 0 fpm (can be 500 fpm or 750 fpm). TThe controlle er automattically changees to a higheer pressurizin ng rate of 75 50 fpm when a cargo fire is detected.
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After landing the controller maintains a diff/press of 0.15 PSID equivalent to 300 ft below landing field elevation until both engines spool down to N1 < 50%, or N2 < 84% for at least 1.5 seconds where after the outflow valve is commanded open. An OFF SCHED DESCENT indicates together with a Master Caution and an AIR COND annunciator when descending before the set FLT ALT is reached. When this occurs, the CPC resets the landing altitude automatically to the (stored) departure field elevation so you don’t have to reset when an immediate return to that field is required. Anytime when the FLT ALT is changed during flight, the destination field elevation data is lost. If one CPC fails the backup CPC takes over, indicated by an amber AUTO FAIL (Master caution and AIR COND annunciator) and a green ALT light on the control panel. When ALT is selected on the mode selector, the AUTO FAIL light extinguishes but the ALT light remains illuminated. If both CPC’s fail it will be indicated by an AUTO FAIL and Mater Caution together with FLT & LAND ALT flashing dashes in the windows. (no ALT light as this function is unavailable)
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Runw way Awa areness and Ad dvisory System (RAAS)) Sometim mes referred to as “SMAR RT LANDING FEATURE”. The RAA AS is an optio on on the Boeeing 737NG to the EGPW WS, which speecifies alertss or notifications regardin ng your position, and a ru unway positio on. It is highly SA improvving, preventting runway incursion ns by visuallyy (on the ND D) and verbal warnings, iff an incorrectt aircraft possition versus runway exists. The system uses the aircrafts GP PS position in n conjunction n with EGPW WS stored airp port and run nway data which are comp pared, wheree after an aleert is passed onto the cockpit system ms. When RA AAS is enableed the system operates w without any action of thee crew. The ccallouts can be stopped by selection n of the Runw way Inhibit SSwitch on the e EGPWS con ntrol panel in ndicated by tthe RUNWAYY INOP lightt. The light also illuminates when input data (GPSS, Airport datta) to the RA AAS operatio on is incorrecct or not available. ROUTINE ADVISORIEES 1. Approaching A g Runwaypro ovides in‐the e‐air awaren ness of which h runway thee aircraft is lined up w with on appr roach. 2. Approaching A g Runway prrovides on‐th he‐ground aw wareness of approximate runway ed dge being a approached by the aircraaft during taxi operations. 3. On Runway O provides aw wareness of w which runwayy the aircraftt is lined‐up with. 4. Distance Rem maining provides awareness of aircraft along‐traack position rrelative to th he runway end..
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NON‐RO OUTINE ADVISORIES 1. Runway End d improves awareness off the aircraft position relaative to the rrunway end during low visibilityy conditions. 2. Approaching A g Short Runw way providess in‐the‐air aawareness off which runw way the aircrraft is lined‐up with h, and that the runway leength available may be m marginal for normal landing o operations. 3. Insufficient Runway Len ngth providess on‐the‐ gro ound awaren ness of which h runway the e aircraft is lined‐up w with, and thatt the runwayy length avaiilable for takkeoff is less than the defined minimum takeoff runway length. olding on Ru unway adviso ory provides crew awareness of an exxtended hold ding 4. Extended Ho period on the runway. T ke‐Off provid des awareness of excessive taxi speed ds or an inad dvertent take e‐off on 5. Taxiway Tak a a taxiway. 6. Rejected Takkeoff / Distaance Remain ning providess position aw wareness durring a Rejecte ed Take O Off (RTO). 7. Taxiway Lan T nding providees awareness that the aircraft is not lined up with h a runway aat low a altitudes.
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Electtro Moto or Drive en Pump ps Overrheat I receiveed a question n why the EM MDP’s have aan OVHT protection and tthe EDP’s no ot. Looking at the imagee it is self exp planatory, th he (EMDP) hyydraulic pum mp is operateed by an elecctro motor th hat creates h heat, where tthe engine driven hydrau ulic pump is directly driveen from the engine gear boxx. The impeller type first stage pump,, pumps cooling fluid to tthe casing arround the electro motor an nd pre‐presssurizes fluid ttoward the n nine piston vvariable displlacement pump. The EMD DP’s have an overheat indication if th he pump ove erheats, it deepends on tyype whether it is the cooling ffluid, or the eelectric moto or that overh heats. Also itt depends on n type if the motor shutss down automattically, or only the light illuminates w when an overheat exists rrequiring crew action to sshut that overheated pump do own. The elecctro motor teemperature sswitch bringss ON the OVERHEAT light when the ttemperature e is 113 °C or mo ore, and reseets at temperratures betw ween 85 °C an nd 102 °C. The seco ond possibilitty is that AC power is rem moved autom matically when the electro motor temperaature reaches 124 °C or m more stoppin ng the EMDP P, and resets at 60 °C to 7 71 °C. The casee drain fluid ((also from th he EDP) is routed through h the oil‐to‐ffuel heat excchangers on tthe bottom o of the main w wing tanks, tto cool the h hydraulic fluid before returning back into the rese ervoir. This is do one for cooliing purposess (of course) and to preve ent foaming. Rememberr that for gro ound operatio ons there is aat least 760 K Kg’s of fuel needed, to acccomplish en nough cooling for sufficie ent heat exchanger (cooling) operation. Switchin ng the EDP to o OFF only closes the dep pressurizatio on solenoid vvalve downsttream of the pump, stoppingg the output to the system. The EDP supply shutoff valve upsstream of the pump stayys open until the FIRE SW WITCH is pulle ed.
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Cockp pit pane el “+” sy ymbols.. (red circled on the im mage) I used to o look around d, challenge (and being cchallenged) p pilots and fligght engineerrs during “lon ng haul” flights, ggaining know wledge on sysstems and fliight deck surrroundings not to forget situational awareneess. Did you eever wonderr what the raandom “+” syymbols on th he cockpit/avvionics panels represen nt? Probablyy not, but herre is the short explanatio on on them aas a nice to kknow subjectt. Underneeath those “++” symbols aare the wiringg connectorss (Cannon plugs) located for backligh hting, or electricaal components (gauges) o on that paneel. If any of th he electrical componentss / backlightiing fail or flickers, gently tap th hat location o on the panel to possibly correct the o occurring problem. Of co ourse, after lan nding call outt maintenancce to have a closer look aand fix it.
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Overh head (P P5) pane el drain ns. I’ve receeived a questtion from one of our follo owers related to the prevvious post, m meaning whaat is the purpose of specific (and sometim mes unnoticeed) compone ents around yyou on the flight deck. Q: Whatt is the purpo ose of the veertical tubes against each h side of the fwd (1L and 1R) window w frames coming ffrom the oveerhead panel? hat it Well . . . . it appears that the fibeerglass insulaation blanketts of the oveerhead panel are such, th creates ccondensation above the overhead paanels. The vaariation of warm cockpit air and cold aircraft skin tem mperature pro oduces a mo oist environm ment behind the P5 overh head panels,, which could d cause electricaal problems. Boeing mounted a plastiic “drip pan”” to collect th hat condensaation moiste er and drain it tthrough thosse tubes to the aircrafts ffuselage drain system.
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Close ed crosssfeed va alve on ttakeoff and lan ndings? n the NG to asssure that thee fuel crossfee ed valve is clossed for takeofffs and landinggs. I’ve There is aa limitation on found thiis rule in somee Company M Manuals and trry to give a decent answer tto the questio on of one of our followerss . . . . WHY?? do NOT know if this limit stiill applies Note thatt it is NOT in tthe up‐to‐datee manuals I haave in my possession so, I d to your current Compaany operation so . . . . be caareful. (although the rule caan do no harm m if all is norm mal) urce for the crrossfeed valvee is the Batteryy Bus. By the waay, power sou Original (older) AFM teext: Fuel Limitations: • Fuel Crossfeed valve must be closed for takeoff and laanding. he reasons I caan think of, is that when yo ou’re close to an imbalance condition (45 53 Kg/1000 Lb bs), you One of th could end d up with an aactual imbalan nce when the valve is open. u, Boeing philo osophy inhibitted the imbalaance warning on the groun nd so it’s only active in the aair. Don’t Mind you ask me w why, just pay aattention on th he wing fuel lo oad before yo ou rotate prevventing unwan nted roll move ements dependin ng on the amo ount of differeence between the wing tanks fuel load. er) deliver an equal pressurre, so a pump p with Further eexplanation is that the fuel pumps do nott always (neve higher ou utput pressuree will feed botth engines witth an open cro ossfeed valve resulting in aa possible imb balance during higgh thrust setttings. This wou uld be the casse at takeoff–,, or go‐around d thrust, so exxactly the mom ments that you’re not paying attention to tthe fuel load. WD pump pro oduces the higghest pressure e thereby feed ding both enggines with fuel from In the image the left FW tank #1. h a massive strructural probllem (crash) th he valve Another rreason could be safety wisee, if anything happens with separatess both sides o of the fuel man nifold reducin ng hazardous cconditions.
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Ambe er AUTO O BRAK KE DISAR RM Ligh ht The next conditions are related to th he illuminatio on of the AUTO O BRAKE DISA ARM light. BRAKE DISARM M light will illu uminate First of alll it starts wheen RTO is selected (on the gground), the aamber AUTO B for 1 – 2 sseconds,indicating a self‐teest of the system and when n successful it will extinguish. On the other hand when thee test was not successful, th he light remains illuminated d. 2, 3 or MAX) iss madefor landing and the system self‐teest fails, the A AUTO When any auto brake sselection (1, 2 d auto brakingg is inhibited. BRAKE DIISARM light illluminates and n was not manually deseleccted before laanding, the AU UTO BRAKE DIISARM light w will When an RTO selection 1.4) after toucch down sense ed through thee PSEU, and n no auto brake occurs. illuminatee ± 2 seconds (AMM says 1 O BRAKE DISA ARM amber ligght comes on when autobraakes are seleccted and any o of the next conditions The AUTO occur: • a malfunction a ning autobrakee system • a malfunction a ning(normal) aantiskid system m • manually overriding the au utobrake systeem by stepping on the brakkes nitiated, the A AUTO BRAKE D DISARM lightillluminates wh hen: When aftter touchdown (or RTO)autto braking is in • moving the sp peed brake levver down/dettent a e thrust leverss (not within 3 3 seconds afte er touch down n) • advancing the • manual brakin ng is applied depth technical activation o of the light is rrelated to the e Hydraulic sysstem B pressu ure to the systtem and More in d causes th he light to illum minate when: • RTO autobrakke is command ded to apply, and the auto brake solenoid valve presssure is low • the autobrake t e selector is in n the OFF position andthe ssolenoid valvee pressure increases more tthan 1000 PSI ntrol unit (AAC CU) from the A ADIRU The last ccause is an invvalid input to tthe antiskid/aautobrake con on’t forget the light test an nd pushing thee light ;‐) O yeah do
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B737 7 Fire prrotectio on
GENERALL NOTES: • FIRE EXTINGU UISHING is pow wered by the HOT BAT BUSS, so dischargee is possible even with all electrical power OFF. • APU & ENG D A DETECTION sysstems are pow wered by the B BAT BUS. • WHEEL WELL W FIRE DETECTIION system usses AC pwr (A AC transfer buss 1), so this AC C bus must be e or operation o of this system.. powered by aany AC source (APU or Eng. Gen, Ext pwr) energized fo • OVHT & FIRE O indications arre shown on the flight deckk on; o OVHT&FIRE protection panel o Cargo fire panel o Mastter caution paanel. ENGINES • Dual loop oveerheat/fire dettection system m o when n 1 loop fails, there will be NO flight deck indication b but the fault detection syste em automatically discconnects the d defective loop p). or loops on thee same engine e have faults, the FAULT ligght will illumin nate. o If both (2) detecto MASTER CAUTTION. o Therre will be NO M • 2 Extinguishe 2 r bottles for 2 2 engines. h can be used for an “on‐sid de & off‐side” engine fire. o Both APU S re detection. ((no overheat) • Single loop fir • 1 Extinguisher bottle. (can also be contro olled from AP PU ground con ntrol panel) W U FIRE is deteccted, the APU will automatiically shut dow wn but NOT automatically • When an APU e extinguish. COMPARTMEN NT CARGO C • Uses (dual loo op) SMOKE deetection system o 4 dettectors in FWD CARGO COM MP (all 737NG G's) o 4 dettectors (737‐6 600) o 6 dettectors in AFTT CARGO COM MP (737‐7/8/900's) o Poweered by DC BU US 1 & 2. • Fire EXTINGUISHING by; GO COMP (alll 737NG's). o 2 nozzzles (HALON)) in the ceilingg in FWD CARG o 2 (73 37‐600) or 3 (7 737‐7/8/900'ss) in AFT CARG GO COMP. o (EXTINGUISHERS, so powered b by HOT BAT BU US) WHEEL W WELL • 1 Loop for FIR RE detection in n main landing gear wheel well • EXTINGUISHIN NG by lowerin ng the landingg gear (below max LG exten nsion speed)
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Startt switch functio ons. ons on the staart switch represent several important fu unctions towaard starting en ngines and pro otecting All positio the mightty CFM’s against unwanted d or uncommaanded decay in N2, and/or flame‐outs. The start switcch position iis monitored b by the EEC (th hrough the DEEU’s) to activate the several functions and related actions of the switch, together w with igniter selection.
OFF • • • GRD • • • •
•
• • CONT • • •
Neither igniteer is activated when the staart lever is in the cutoff position. O On ground, w when N2 dropss below 57% u until 50%, both igniters will activate wheen the start levver is in t the Idle positi ion. In flight when n N2 drops below Flight Idlee RPM until 5% %, both igniteers will activate when the sttart lever is at Idle.
TThe engine bleed valve (solenoid) is com mmanded to close the valvee. T The starter va alve opens. O On ground, th he selected ign niter(s) will acctivate. o Poweered by Xfr bu us 1 for enginee #1 left ignite er, AC stdb bu us for the right igniter. o Poweered by Xfr bu us 2 for enginee #2 left ignite er, AC stdb bu us for the right igniter.. In flight, both igniters will aactivate. do not want tto be bothered d by a failed sselected igniteer when you n need to start aan engine o You d in‐flight. There mu ust be a reaso on why you waant to start a sshut‐down en ngine like a gre eater blem on the (o only) operating engine. prob T The EEC is pow wered from itts XFR bus (1 ffor eng #1, 2 ffor eng #2) beelow 15% N2 w where after an n AC a alternator mo ounted on thee gear box takes over; o > 15% % N2 (gearbox RPM) it’ll prroduce enough AC power to o operate thee EEC. o On a battery start you won’t see EGT, FF, Oil press & oil teemp until the EEC becomes powered % N2. afterr reaching 15% ECU (APU) recceives a signal to open the APU IGV’s. o To provide maxim mum air capaciity and pressu ure for startingg. A At 56% N2, th he starter swittch is comman nded to move e to the OFF po osition. (AMM M says 55%) o This is the AUTO p position with n newer switch features.
On ground, acctivates the seelected igniter(s) when the start lever is at Idle. O In flight, activvate both ignitters when N2 drops below idle also with the start leveer at the Idle p position. S Selected; o Takee‐off. o Land ding. o Before TAI is seleccted.
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FLT • •
Activate both igniters. Selected; o Adverse weather.
AUTO (when installed) • Ignition is OFF. • Both igniters activate when engine start lever is in IDLE and: o An uncommanded rapid decrease in N2 occurs o On ground, N2 is between 57% and 50% or, o In flight, N2 is between idle and 5%. • Activates selected igniters when: o Below 18000 feet altitude and flaps extended. o TAI is selected.
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Fuel nozzle “coking”. I’d like to emphasize a procedure in the FCOM toward shutting down the engines related to something that’s known as coking. First of all what is “coke” besides cola, it is fuel derived from refined petroleum with a high percentage of carbon. Next the procedure described in the FCOM 1’s Normal Procedures; Operate the engines at or near idle thrust for a minimum of three minutes before shutdown to thermally stabilize the engines and reduce under cowl soak‐back temperatures. Routine cool down times of less than three minutes before engine shutdown can cause engine degradation. Finally the explanation why these 3 minutes of cool down before shutdown; Temperatures in the combustion chamber run up to ± 1700 °C which goes unnoted by the aircrew, as they get the EGT in the exhaust presented on the Upper DU, generally around 400°C at Idle. When shutting down a jet engine at higher than Idle RPM, the temperature in the combustion chamber is substantially higher than at Idle. This results in a higher than normal residual fuel nozzle temperature at shut down, which causes carbon in the residual combustion chamber fuel vapor to settle (coking) on the nozzles. This settled carbon on the nozzles can disturb the normal spray pattern of the fuel on subsequent operation of the engine resulting in a disturbed flame pattern, negatively affecting engine performance and/or damage the combustion liner as in the image. Also there is an example image (not CFM56) of carbon settled on a jet engine fuel nozzle. When Idle RPM is used for ± 3 minutes, the nozzle temperature is that low, that fuel nozzle “coking” is diminished resulting in a safer, more economic, longer, less maintenance engine life reducing costs.
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Dual bleed light The DUAL BLEED light illuminates when there is a “possibility” of two bleed sources to the left side of the bleed manifold where APU bleed connects. These two sources are related to one of the engines and the APU, so not related to two engines or external air. It are the ENGINE BLEED VALVE switch positions, in combination with the ISOLATION VALVE, together with an open APU BLEED VALVE that makes the light to illuminate. So during non engine operation (pre‐flight) with the ENGINE BLEED VALVE switches in ON together with an operating APU and the APU BLEED VALVE open, the light illuminates. The APU needs to operate to open the APU BLEED VALVE by pressure, with the switch in ON and the APU shut down, the valve closes. The APU puts out a nominal pressure of ±36 PSI to the bleed manifold where the engines in Idle (9th stage air) ±32 PSI (or ±34 PSI from the 5th stage) which means that the APU is supplying air to the bleed manifold as this is higher pressure. In this case, the ENGINE BLEED VALVE (and HIGH STAGE VALVE) are closed as it senses a higher downstream backpressure. When the DUAL BLEED light is illuminated according the above explanation, you need to stay at Idle thrust to prevent a possible backpressure to the APU as what the books say. When you move the thrust levers up to above Idle, bleed pressure from the 5th stage increases above 34 PSI closing the HIGH STAGE VALVE and 36 PSI preventing the APU to supply pressure to the manifold, so the engine bleed takes over. Actually the books “tells” the crew that an engine (or engines) and the APU both supply pressure to the pneumatic manifold at the same time which should reminds you to use Idle thrust only. This is just to be sure, that the APU BLEED VALVE is closed when the light is extinguished. Looking at the image it is mechanically impossible (except with a failing check valve) that the engine bleed backpressures the APU because there is a check valve, preventing the APU BLEED VALVE receiving a higher backpressure. There are (a lot) more Boeing design related questions which I receive on a weekly bases which remain questions, so is this light and failing check valve issue but . . . stick to the procedure and restrict thrust to Idle when the light is illuminated.
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Air Cycle Machine operation The two air condition packs provide “conditioned” air (temp and moisture) to the air condition system. They are supplied with bleed air from the bleed manifold, normally from the APU or the high pressure compressor of the engines. Hot air flowing into the packs has a temperature of ± 199 – 227°C, controlled by the engine BASOV and the pre‐cooler. To get to the normal temperature range of 18 – 30°C of the AC system, it needs to be cooled which is accomplished by the two packs. Let’s look at the flow of air through a pack explaining the components the air passes. Air enters the pack through the PACK FLOW CONTROL VALVE and can go in two directions, into the cooling circuit or it by‐passes the so called ACM (air cycle machine) circuitry. The cooling cycle starts at the “primary” (air–to–air) heat exchangers that cools the air. The heat exchanger works like car radiator but exchanges hot bleed air to the ram air duct airflow. On the ground by the way, duct airflow is created by a fan which is driven off the ACM. Next the air “hits” the compressor of the ACM which turns the compressor and turbine resulting in an increase of pressure and temperature. Another “secondary” heat exchanger cools the air again, where after the following components in the manifold are to “create” and extract water from the air. A “secondary” water extractor drains water from the manifold, a re‐heater pre cools the air before it enters the condenser and warms the air (from the 2nd water extractor) before it enters the turbine to increase efficiency of that turbine operation. The condenser creates water droplets in the air where after the “primary” water extractor removes water by creating a swirling motion, “centrifuging” that water to the outer collector wall where it is collected and relieved in the ram air duct adding in cooling at the heat exchangers. The last component of the pack is the expansion turbine which can cool the air to below the freezing level as a function of very fast expansion (extracting energy) of the air. Finally the air is mixed with by‐pass air representing the pack control requested output temperature. The pack is protected against a request of a too high demand of cool air, thereby overloading that pack. This is accomplished by several temperature sensors detecting over temperatures in the cooling cycle causing the pack to trip off line. If this happens, it closes the pack flow control valve indicated by the PACK OFF indication on the bleed panel together with a MASTER CAUTION light (AIR COND). The solution to this problem is to let the pack cool down and selecting a warmer temperature before resetting and “unloading” the tripped pack or else the overload/overtemp will occur again. Water is extracted from the air to protect the avionics in the aircraft against moisture, to prevent mold to form and against oxidation of metal components.
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Airstair The (optional) most common airstair is located below the left forward main entry and is controlled from the forward attendant panel, or from outside the aircraft. Another option is an airstair below the left aft entry door which is not discussed in this post. Inside control: The interior control panel has two modes of operation, NORMAL and STANDBY. • NORMAL operation is powered from the 115V AC standby bus. o either the extend or retract switches are depressed momentarily to operate the stairs. • STANDBY operation is powered from the switched hot battery bus, so the BATTERY SWITCH must be positioned ON. o both, the standby switch AND the retract or extend switch must be pushed and hold to operate the airstair. The forward entry door to has be partially open before electrical power is available to operate the airstair. When the stairs reaches full extension, electrical power is automatically shut off to the motor and the tread lights are turned on. (when rotary switch is in AUTO) NORMAL operation is interlocked by handrail switches to prevent the stair from being retracted with the handrail extended but the STANDBY system bypasses these switches so caution has to be exercised to prevent damage. Outside control: When operating the airstair from the outside, the forward entry door does NOT to be open for airstair operation as the exterior control switch by‐passes the door open requirement. The power selection switch provides NORMAL and STANDBY operation of the airstair and is spring loaded to NORMAL. In NORMAL, the 115 VAC AC standby bus powers the airstair electrical motor so the BATTERY SWITCH needs to be ON. The STANDBY position provides DC power from the 28 VDC switched hot battery bus for airstair operation where this (external STANDBY) switch energizes the switched hot battery bus regardless of BATTERY SWITCH position. Both NORMAL and STANDBY operation are interlocked by handrail switches to prevent the stair from being retracted with the handrail extended. Caution must be exercised when using the maintenance switch located under the airstair. If the upper handrail extensions are not properly stowed before retraction, damage to the airplane structure or damage to the airstair handrail may result. An amber AIRSTAIR light, located on the overhead door caution annunciator panel, illuminates (provided DC bus 1 is powered) when the airstair pressure door is unlocked, also illuminating the AIRSTAIR light and the DOORS annunciator light together with the MASTER CAUTION lights
.
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Equipment Cooling Electronic and Electric equipment heats up substantially when used and requires cooling to operate without overheating. The equipment cooling system consists of a supply‐, and an exhaust duct with respective NORMAL and ALTERNATE fans. (4 fans) The two separate systems cool their own respective electronic components either by applying cool air, or removing warm air. The result of a failed or shut down individual system, affects specific components which are described next; The “supply system” pushes cool air from the Cabin Compartment to, and affecting: • Captains DU’s • Lower DU • Captains CDU • Aft electronic panel • Equipment racks in the E&E compartment The “exhaust system” pulls warm air and relieving it into the FWD Cargo Compartment Liner though, and affecting: • First Officers DU’s • Upper DU • First Officers CDU • P6 circuit breaker panel • Overhead panel • Equipment racks in the E&E compartment. Flow sensors in the supply and exhaust duct indicate a lack of airflow which results in the illumination of the related EQUIPMENT COOLING OFF light , the OVERHEAD annunciator and a MASTER CAUTION light. Selecting the alternate fan should restore airflow and extinguish the OFF light within approximately 5 seconds. If an overtemperature occurs on the ground, a crew call horn in the nose wheel well sounds. Additional cooling flow is created by the open overboard exhaust valve on the ground, and at low altitudes as the valve closes at 1 PSID. (± 3000 ft). Executing an in‐flight forward cargo smoke alarm procedure, power to the normal and alternate exhaust fans is interrupted and the exhaust low flow detector is inhibited for the remainder of the flight. (no indications) The stopped exhaust fan(s) prevent smoke from entering the occupied compartments.
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Overb board E Exhaustt Valve The previious post expllained Equipm ment Cooling, one of the components meentioned in th hat (exhaust) ssystem was the O Overboard Exh haust Valve (O OEV). Let’s takke a closer loo ok at that valvve on its purpo ose and operaation. nd Electrics Co ompartment and reliefs exhaust air overrboard in The OEV is located in tthe aft of the EEquipment an open possition behind tthe E&E accesss hatch. The vvalve has two functions, it ccontrols equip pment coolingg exhaust air that flows overboarrd, and it has a function during a smoke removal proccedure. ment cooling eexhaust air ovverboard when the airplanee is on the gro ound and at a lower The OEV reliefs equipm altitude tto improve cooling. When tthe OEV is clossed in flight, tthe Equipment Cooling exhaust air is routed into the forwaard cargo com mpartment lineer as a meanss of heating. N Note that the C Cargo Compartments are n not supplied with fresh waarm air duringg flight as theyy are closed arrea’s for fire eextinguishing rreasons to me eet Class C comparrtment requirrements. board exhaust valve has three modes off operation. The overb Normal 0 ft. The norm mal mode applies with The OEV is open on thee ground and closes when 1 PSID is reached at ± 3000 wing bleed panel switch po ositions: the follow • Left and right pack switch –– AUTO/OFF N switch – AUTTO • R RECIRC FAN w High flow The high flow mode im mproves ventillation by an in ncreased air fllow and depends on the deegree of opening of the Valve (OFV) and the next bleed panel sw witch selection ns: Outflow V • L or R PACK sw witch – HIGH N switch – AUTTO • R RECIRC FAN he OEV is com mmanded open n when the OFV > 3.5° but remains closeed when the O OFV < 2°. In this configuration th Smoke reemoval. The smokke removal mo ode opens thee overboard eexhaust valve from full open to 54° open n, to remove smoke from the flight deck an nd E/E compartment. The smoke removaal mode is activated with th he following sw witch positions on the bleed panel: • L PACK or R PA ACK switch – HIGH N switch – OFFF • R RECIRC FAN
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Thermal ele ectrical protecttions. Electrical and avionics equipment iss protected wh hen overloading or shortingg during operaation by several mmon is the direct accessible and resettable circuit breeaker (CB). The less know and devices, tthe most com hidden, aare un‐resettable current lim miters and meelt fuses which protect heaavy user comp ponents and th he bus bars. Let’s focu us on the circu uit breakers ass they are meentioned in sevveral procedu ures and direcctly accessible on the flight decck and in galleeys. operational information. First NEED TO KNOW o onent behind a “tripped” CB needs to bee checked firstt by maintenaance On the grround, the eleectrical compo before it is allowed to reset. why the CB triipped and thee risk is presen nt that you initiate a In flight itt is different aas there must be a reason w fire hazarrd by resettingg without kno owing what caaused it to trip p. The generall (QRH) rule iss to allow a “cool down” tim me of about 2 2 minutes befo ore resetting but be aware WHAT you’ree resetting. If tthe component is non essential for the safetyy of flight, leavve it alone. If it is, it has to be closely loo oked at as there are cases in n the past n resetting thee CB only oncee so . . . . “sit o on your handss” and think iff it is really ne ecessary that caussed a fire even to push the CB back in. It is the soun nd judgment o of the crew with the respon nsibility of thee Captain to determine mpletion of a fllight. The QRH H also guides tthe flight crew w to reset or p pull CB’s if a reset is needed forr the safe com on‐normal pro ocedures but sspecial attenttion is needed d as described. during no neral rule in‐flight bearing in mind that o often a DC CB only is contro ol (switching) p power, I personaally used a gen and AC iss operating po ower as shown n in the imagee. The rule was to reset esseential DC CB’ss once after ± 2 minutes w when that sysstem is really n needed, but b be specifically cautious with h AC CB’s and NEVER RESETT FUEL RELATED CB’S. Also neever use a CB aas a switch beecause that’s n not where theey are designeed for and you u’ll peration of the device. I’ve seen CB’s pop pping almost when you onlly look at them m as in degrade tthe correct op our (P‐3 LLockheed Orio on) operation we had CB’s pulled and pu ushed every flight which weere later replaaced by switches.. on this subjectt, circuit breakers are found throughout the aircraft aand are normaally heat Ok enouggh “warned” o triggered. Newer type CB’s are electtro‐mechanicaal operated and are actuallly relays with a coil. When tthe ntacts of the rrelay, stoppingg Voltage flow w to that respective current fllow becomes too high the ccoil opens con electrical component. heats up and curves up (as in the With therrmal CB’s, in ccase of an oveerload or shorrt, a bi‐metal in the device h image) w which by overccenter spring fforce action “pops” the shaaft out of its base showing aa typical white e shaft underneaath its top. This interrupts tthe control orr power Voltagge flow througgh the CB to tthat compone ent, protectin ng it against itss malfunction and possible overheat con ndition. presents the lo oad it is protecting on, in teerms of Amperes. The Circuit brreakers have aa number on ttop which rep higher the Amps, the m more cautiouss you need to be to reset th he CB as they present heavyy load equipm ment with a higher rrisk of fire hazzard.
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Fuel ttemperrature in ndicatio on. Yesterdayy a B737 instrructor made m me curious on why fuel tem mperature is m measured in taank 1? It appears to be a “lefttover” from B7 737‐200’s where system A was hydrauliccally pressurizzed by two ED DP’s and P’s. As explain ned in an earlier post, the EMDP electro motors heat u up significantlly and are system B by two EMDP uid. This heateed up (case drrain) fluid is ro outed through h an oil‐to‐fueel heat exchan nger on cooled byy hydraulic flu the botto om of tank 2 b before returning to the reseervoir, increassing the fuel teemperature in n that tank. In n that case, fuel in tank 1 would be much ccolder than in n tank 2 which h limits (‐43°C or 3° above freezing point)) need to be monitored more than warm fuel (dissolved airr). It was for this design phiilosophy that fuel temperattures are d from a FUELL TEMPERATU URE BULB in taank 1 to stay informed on th he coldest fueel temperaturres. measured make a differen nce as both syystems have the same pressurization feaatures (1 These dayys on the NG’s it doesn’t m EDP & 1 EEMDP) so tem mperatures wo ould be near tthe same whe ere case drain fluid goes thrrough a oil‐to‐‐fuel heat exchangeer on the bottom of both taanks. MP INDICATOR R comes from the 28 VAC transfer bus so o if it loses power the By the waay, power for the FUEL TEM indication n “freezes” (A AC lies, DC diess)
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Integrated Drive Generator (IDG) I’ve already discussed the operation of an AC generator but focus now on the drive that creates a constant RPM of the Accessory Gear Box (N2) mounted generator. To create the nominal 400 Hz of the AC generator, it needs to turn at a steady 24.000 RPM which is done by converting N2 RPM. N2 at IDLE is ± 8.400 RPM, where it is ± 15.183 RPM at its MAX allowed operational RPM of 105%. This conversion is achieved by a Constant Speed Drive (CSD) which hydraulically/mechanically transfers the, by the HPC driven Accessory Gear Box (AGB) RPM into the needed Generator operating RPM. The components needed for this speed adjustment are the “FIXED DISPLACEMENT HYDRAULIC UNIT “ and the “VARIABLE DISPLACEMENT HYDRAULIC UNIT” in the IDG, using oil as the name already explains. The CSD has its own oil system which is cooled by tapped off Fan air, through an oil–to–air heat exchanger/cooler and by fuel through an oil–to–fuel heat exchanger/cooler thereby increasing the fuel temperature on the latter. The DRIVE light is activated by an IDG oil pressure switch, and illuminates when pressure is below its limit of 165 PSI where normal operating pressure is 240 – 290 PSI. This is anytime when: • Engine is shut down • IDG is disconnected • IDG overtemp occurs (automatic at 182°C) • IDG oil pump failure • IDG oil loss • IDG under frequency occurs with the engine running!! • IDG drive shaft sheared. If the light illuminates by one of the previous causes, the QRH is directing you to disconnect the IDG from the AGB using the DISCONNECT switch on the “GENERATOR DRIVE AND STANDBY POWER PANEL”. This switch activates the DISCONNECT SOLENOID when the respective START LEVER is in the IDLE position. The switch is safety wired to the panel to prevent inadvertent operation and activating the switch with a shut down engine (preflight) doesn’t activate the solenoid. (START LEVER at CUT OFF) DON’T GO THERE ;‐) When the IDG is disconnected, the action is irreversible and the IDG has to be mechanically reset (reconnected) by maintenance. If this disconnected is due to an overtemp, the IDG has to be replaced and is not allowed to be “just” reset by maintenance. The IDG drives shaft shear device shears the drive shaft in case of an IDG mechanical malfunction to protect the IDG gear train in the AGB against damage.
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The oil leevel sight gage has an ind dicating silver band and is reliable >5 5 minutes affter engine shutdow wn. When thee oil level is b below the sillver band, th he IDG oil levvel is low and d servicing is required d. When (cold d) oil is abovve the silver b band, the IDG oil level is high and oil has to be drrained off. With h hot oil, a leevel above th he silver band d is acceptab ble provided it is below tthe DRAIN mark. Be aware that the left, and right engine side gage on each IDG reads different as a result of being mounted d on the sam me left side o of the enginee and wing diihedral.
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Electtrical Lo oad Shedding Load shed dding is comm manded by thee Bus Power C Control Unit (BPCU) to prevvent a single A AC source from m overloadiing by de‐eneergizing certain n electrical bu usses in seque ence of prioritty. u, each generaator has its ow wn Generator Control Unit (GCU) that ho old individual protections su uch as; Mind you * Over‐ , under voltagee, * Over‐ , under frequency, urrent, * Unbalanced phase cu ure, * Generator diode failu * Phase ssequence, * Over cu urrent, * Differen ntial fault, * Under sspeed protecttion, * GCU faiil‐safe, taking thee respective ggenerator OFFF LINE when any of the previouss occur. d shedding. ENGINE ggenerator load or load sheddiing sequence;; Single engine generato R bus 2 1st) Gallley and main busses on XFR 2nd) Galley and main busses on XFR R bus 1 3d) IFE buses hanges to a seecond operating generator, automatic load restoration n of the main buses, When configuration ch buses occurs. If this doesn’tt happen, man nual restoratio on can be atteempted by mo oving the galley busses and IFE b n back ON. CAB/UTILL Power Switch to OFF, then hedding APU geneerator load sh Ground APU GEN attempts to carry all electrical lloads. A single A When an overload occcurs, Galley an nd main bussees are de‐enerrgized until the load is within limits. Flight PU GEN autom matic load shed dding sequence; Single AP 1st) Gallley busses. 2nd) Main busses. (see explanation n below) 3d) IFE busses. U detects an EEGT rise abovee limits which can be cause ed by electricaal and air (enggine starts) demand, If the ECU the Main busses de‐en nergizes. pted by selectiing the CAB/U UTIL Power Sw witch to OFF, tthen back ON.. Restoringg bus power can be attemp used by high EEGT on the grround, the bussses automatiically restore w when EGT is w within limits. When cau
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Comm mon Dissplay Sy ystem (C CDS) ma alfunctiions. I’ve earlieer posted an article about the DEU’s bu ut want to exxplain the CDSS (malfunctions) a little further. The CDS displlays performaance, navigatio on and enginee information on the six Dissplay Units (DU’s). omponents; The CDS cconsists of thee following co Display seelect panels Engine display control panel EFIS conttrol panels Display so ource selectors Display electronics units (DEUs) Coax couplers Display units (DUs) nels Brightnesss control pan Remote light sensors (RLSs). CDS on the gro ound only If a fault occurs in any card in the C he 2nd enginee start, it is presented as an a amber before th CDS FAULLT indication o on both PFD’ss below the sp peed tape or a white CDS M MAINT message. When botth engines are operaating or in thee air, the undisspatchable CD DS FAULT chan nges to an am mber DISPLAY SSOURCE. AULT/DISPLAY SOURCE indiccates a total D DEU failure tellling you that a “critical card” in the CDS has A CDS FA failed. ngle card malfunction includ de: These sin • Input/output controller • Power supplyy • Processor. or two DEU’s. It also could indicate a combination of (less important) failing ccards in one o Multiple cards also incclude: G rator • Graphic gene • Discrete inputt/output A /output • Analog input/ OR is on one D DEU supplyingg data to all 6 DU’s. The DISPLAY SOURCE aalso indicates when the SOURCE SELECTO 1 / ALL ON 2) (ALL ON 1 Note: witching betweeen sources, (ALL ON 1 – AU UTO – ALL ON N 2) leave the switch 1 – 2 sseconds at the e When sw intermed diate position or else the 2 D DEU’s can sho ow incorrect d data. Note: OM 2 at “Fligh ht Instrumentts, Displays” h how a DISPLAYY SOURCE affeects your auto omated Check your current FCO flight relaated to AP usee. DS MAINT show ws when one of these SING GLE (less impo ortant) circuit cards The whitee dispatchablee message CD fails in eitther DEU: • Graphic gene G rator • Discrete inputt/output A /output • Analog input/
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Cargo Compartments air. There is some confusion about the cargo compartments related to air inside the holds. The current Boeing CBT’s explain that air from the Equipment Cooling System and from the Recirculation Fans enter the (Fwd) Cargo Compartment what appears to refresh air inside and warm the compartment. This is NOT true as the Cargo Compartments are "Class C" fire compartments, meaning they confine a fire. So when the fire agent is relieved in the compartment, it takes away the O² without fresh air circulation to feed the fire. The air from the Recirculation Fans are diffused in a “double skin” around the Cargo Compartments, thereby warming the compartments up without any regulation. The liner that creates that double skin is designed as a fire barrier to isolate the compartments from the rest of the aircraft. The Forward Cargo Compartment also receives Equipment Cooling air when the aircraft is above ± 3.000 feet (> 1 PSID), when the Overboard Exhaust Valve is closed and reliefs into the liner. The Aft Cargo Compartment only receives air from the Recirculation Fans and from the Cabin where the Outflow Valve creates an increased airflow through the liner depending on differential pressure at that moment warming up the Aft Compartment. All together this normally results in a higher temperature in the Forward Compartment compared to the Aft Compartment. The compartments are pressurized through an Pressure Equalization Valve, so as the aircraft climbs or descents there is a flow out and into the compartments for pressurization. At level altitude the compartments are “closed” and there is no airflow in or out creating those fire confining areas. Last, the compartments are equipped with Blowout Panels to backup the Equalization Valves in case of a fast rate of change in pressure around the compartments as in a rapid decompression.
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NiCad Battery operation. The B737NG is equipped with either 1, or commonly 2, a Main and Auxiliary NiCad Battery located in the E&E bay. The purpose of the Batteries is to provide Emergency Power and to be able to start the APU from the Main Battery when no AC Power (XFR bus #1) is available. The Battery has a capacity of 48 Ampere‐Hour which can, when fully loaded provide “normal” Emergency Electrical Power operation to the electrical system for 30 minutes(each). The standard Voltage range is from 22 – 30 VDC as charged to its maximum Battery capacity by their respective Charger from AC Ground Service bus #2 (Main), and AC GS bus #1 (Aux). When supplying Emergency Power, the Batteries are paralleled by the Remote Current Circuit Breaker (RCCB) to equalize their discharge and basically can be explained as, any time when the Emergency Inverter (that converts Battery DC into AC) powers the AC Standby bus, the RCCB is closed. When starting the APU from the Battery, the power comes from the Main Battery ONLY as the RCCB is commanded open, most likely preventing draining both Batteries in an attempt to start the APU with a loss of both Generators but also applies when starting on the ground without AC power on the aircraft. Emergency power is provided for: (Battery Switch ON) • Hot Battery Bus (always connected to the Main Battery) • Switched Hot Battery Bus • Battery Bus • DC Standby Bus • AC Standby Bus (through the Emergency Inverter)
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When selecting the Battery Switch to ON, the batteries are discharged which can be seen on the Electrical Power Module Panel display on the fwd overhead panel as selected. An amber BAT DISCHARGE light illuminates when there is an excessive discharge load (Amps) of a Battery for: • 95 seconds more than 5 amps • 25 seconds more than 15 amps • 1.2 seconds. more than 100 amps So initially there is no DISCHARGE light until one of the previous values has been exceeded, it takes some time to determine that the Inverter drains the Batteries at a higher current rate. Of course the light is accompanied by a MASTER CAUTION and the amber ELEC annunciator. The BAT DISCHARGE does NOT illuminate when this load is the result of an APU start using the Main Battery. When a Battery discharge is detected on the ground, a horn will sound ± 2 minutes after detection to alert the ground crew, meaning the Battery discharges without a charger providing power to the Battery. The APU start attempts are restricted by using the APU’s Starter Power Unit (SPU) and Starter Control Unit (SCU) that convert 28 VDC or 115 VAC to the required 270 VAC which heats up these components. The restriction is three attempts, where after 15 minutes cooling is required of these SPU and SCU. Here is my approach of how the Battery is affected by an APU start using the max starting time of 120 seconds where the starter is cutoff at 70% meaning 84 seconds at ± 400 Amps. One attempt would take 9.3 Amp/hr from the Battery leaving (48 – 9.3) 38.7 Amp/hr and 20.1 Amp/hr after three attempts. This discharge would be affected by the quality of the Battery, the rate of discharge, the time the starter is engaged and temperature.
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Climb Thrust Reduction When you’ve selected a fixed derate and/or ATRT thrust reduction for takeoff on the N1 LIMIT page, the FMS computes on top of this selection an additional thrust reduction during climb. This recomputed value is automatic, and is required to avoid a climb N1 value greater than the reduced thrust takeoff N1value. There are two fixed climb thrust reductions available on the N1 LIMIT page: CLB–1, which gives a climb thrust limit reduction of 3% N1, and is an equivalent of ± 10% thrust reduction. CLB–2, which gives a climb thrust limit reduction of 6% N1, and is an equivalent of ± 20% thrust reduction. Normally, selecting TO–1 automatically arms CLB–1 and selecting TO–2 automatically arms CLB–2 but also could be the outcome of a combination with ATRT selection. Automatic arming of CLB‐1 or CLB‐2 by the FMS depends on various additional conditions such as environmental and aircraft and engine configuration. The FMC automatically selects the highest climb thrust available (CLB, CLB‐1, CLB‐2) which would not result in a thrust lever push, when the aircraft transitions from takeoff to climb. is displayed inboard of the selected climb N1 limit and If a CLB–1 or CLB–2 is selected, the N1% for CLB and the N1 cursors still display values for full rated climb. Climb thrust reduction initiates at 1500 ft AGL indicated by the N1 rolling back the required percentage (3% or 6%) where after it immediately starts to slowly increase to the full (fixed) rated thrust selection. This full rated thrust will be reached when not interfered at 15.000 ft.
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The ““white b bug”. On takeo off the whitee bug is automatically sett to V2 +15 aand on appro oach to Vref +20. At takeo off and speed ds above whiite bug, the aaircraft has FFULL maneuvverability and is provided d when the airplane has acceelerated to the recommeended maneuver speed ffor the selectted flap posiition. This meaans up to 40°° AOB, i.e. 25 5° AOB + 15 overshoot upset. When below whitee bug, it has LLIMITED maneuveerability to 3 30° AOB i.e. 1 15° AOB + 15 5 overshoot. The bug disappears at the initial flap ps retractio on after takeoff or when VREF is seleccted in the C CDU. On approach, when at normal landing flaps aare set and aabove Vref, the aircraft has full G/A, and the e Flaps are reetracted to 15, the full maneuveerability. However in thee event of a G maneuveerability & limited maneuverability ccriteria applyy to the whitee bug again. This is becau use Vref 30 (&Vreef40 + 5) = V2 F15, and th hus the whitte bug is V2 FF15 + 15, and d equals the takeoff case e as if a F15 takeeoff. (This is aalso why thee magenta CM MD speed bu ug automaticcally moves aabove the white bug as the Flaps are retraacted to F15 for a 2 engin ne G/A). In the caase of a single engine app proach Vref 15 = V2 F1. TThus in the G G/A the samee limits on Baank Angle ap pply as the Fllaps are retraacted from FF15 to F1. In this case thee magenta sp peed bugs sttays at Vfly to give best Ratee of Climb up p to1000 ft. ((Note the QR RH saying baank angle lim mited to 15° ffor a single en ngine G/A un ntil at safe sp peed)
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Standby Hydraulic System operation. The standby hydraulic system is a backup system for the normal hydraulic A and/or B system in the event of a failure, for the next components: • Thrust reversers • Leading edge flaps • Leading edge slats • Rudder o Standby Yaw damper The Standby Hydraulic System operates manually by selection of: • Leading edge devices by: o FLT CONTROL A or B switch to: Standby Rudder o ALTERNATE FLAPS arm switch o ALTERNATE FLAPS control switch • Thrust reversers o Operating the thrust reverser handles • Standby Yaw Damper o FLT CONTROL A and B switch to: Standby Rudder The Standby Hydraulic System operates automatically for the: • Rudder PCU o By command of the Force Fight Monitor Also the standby pump operates automatically if ALL of the next conditions exist: • FLT CONTROL A or B switch ON and, • ALTERNATE FLAPS arm switch OFF and, • Trailing edge flaps not up and, • Aircraft in the air, or wheel speed more than 60 kts and, • Low flight control hydraulic pressure. The purpose of the automatic standby hydraulic system operation is to have enough rudder control during takeoff, approach, and landing if either or both of the main hydraulic systems fail. When the standby hydraulic system activates, the amber STBY RUD ON light illuminates.
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The force fight monitor (FFM) is installed on modified (NG) 73’s to detect opposite pressures on the A, and B hydraulic system at the main rudder PCU actuator. This could be an indication of either hydraulic system, input rod or control valve experiences a malfunction. When such a conditions is detected > 5 seconds, the FFM will automatically activate the standby hydraulic pump thereby pressurizing the standby rudder PCU.
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Transformer Rectifier Units. (TRU) There are three TRU’s located in the E & E bay racks that convert 115 VAC into 28 VDC with a maximum load of 75 Amps with sufficient (enforced) cooling or 50 Amps with regular cooling. The TR’s output can be individually checked on the (overhead) metering panel. TRU 1 receives power from AC Transfer Bus 1 and feeds DC Bus 1 in normal operation. TRU 2 receives power from AC Transfer Bus 2 and feeds DC Bus 2 in normal operation. TRU 3 receives power from AC Transfer Bus 2 and feeds the Battery Bus in normal operation. TRU 3 receives power from AC Transfer Bus 1 in the event of a failure of AC Transfer Bus 2 through the energized TR3 Transfer Relay (TR 3 XFR RLY). DC Bus 1 and 2 have a cross redundancy by the Cross Bus Tie Relay (or DC Bus Tie Relay) if a TRU fails operation but automatically opens: • At glide slope capture during a flight director or autopilot ILS approach. o This isolates DC Bus 1 from DC Bus 2 during an approach as a redundancy, to prevent a DC Bus malfunction (such as a short) from affecting both navigation receivers and flight control computers. Note: When a DC Bus experiences a short, the electricity tries to follow the path of least resistance thereby possibly affecting the operation of the “healthy” DC Bus. • When the Bus Transfer Switch is positioned to OFF. o This enables the crew to isolate the left (1) DC system from the right (2) DC system together with the left (1) and right (2) AC system when needed. When on the ground, any malfunctioning TRU will illuminate the amber TR UNIT Off light on the Metering Panel together with the ELEC master caution annunciator light. In flight this light illuminates when either TR 1 fails or a combination of TR 2 & 3. The reason of these combinations is to warn the crew that at glide slope capture during a flight director or autopilot ILS approach you’ll lose either DC Bus 1 (TR 1) or DC Bus 2 (TR 2 & 3). Any combination of two TRU’s is capable of powering the complete electrical system. The diode after TR 3 allows TR 3 to provide a backup for TR 2 & 3 but prevents for some (to me yet unexplained) reason TR 2 from powering the Battery Bus.
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RAM AIR DUCT doors. The RAM AIR DUCT system supplies and controls cooling air to the air‐to‐air heat exchangers of the air condition packs. As the word describes it uses ram air created by forward motion of the aircraft. On the ground airflow is created by a fan which is turned by the AIR CYCLE MACHINE (ACM). On the ground, a DEFLECTOR DOOR mounted at the inlet of the duct extends to prevent debris from entering the duct when taking off or landing on a contaminated runway. The nose wheels spurt contamination sideways and up in the direction of the ducts where it becomes deflected away from them, to minimize clogging the air‐to‐air heat exchangers inside the ducts. When the heat exchangers clog up, it might affect temperature control of the PACK possibly creating an overheat and a PACK tripping of. Inside the duct are the RAM AIR DOORS (RAM AIR MODULATION PANELS) mounted that modulate to accomplish a constant temperature of 110°C measured between the ACM compressor and the secondary air‐to‐air heat exchanger. As most of the air condition components this is measured and controlled by the AIR CONDITION ACCESSORY UNIT (ACAU) in conjunction with the PACK/ZONE CONTROLLER. On ground; The ACAU commands the RAM AIR DOORS to be completely open indicated by the RAM DOOR FULL OPEN light(s) to establish the best cooling flow with slower and no forward motion of the aircraft. In flight: The ACAU commands the RAM AIR DOORS to move from the open position, modulating to such a position to maintain that 110°C duct temperature, mainly at a faired position to reduce drag. The FCOM states that the doors will be fully open indicating the RAM DOOR FULL OPEN in slow flight with the flaps not fully UP. This is a result of low airflow causing the air temperature to reach the 110°C not being able to maintain a lower temperature which drives the door to full open. In the ACAU and PSCU (AIR/GND) systems that control the doors are no speed reference signal inputs. Other possibilities of an illuminated RAM DOOR FULL OPEN indication but with flaps UP even at cruise altitudes could be: ‐ The ram air duct could have an obstruction ‐ On or both air to air heat exchangers are dirty not allowing enough airflow through them ‐ An electrical failure causing a high temperature sensing or door fails in open
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Standby Power. The Standby Power System is required for safe flight operation to power the most important electrical components that receive power through the Standby–, and Battery busses in the event of a partial or total electrical failure. It also powers those busses during ground operation without AC electricity available. The Standby Power Switch controls power to the Standby busses and can be positioned to BAT, OFF and its normal (guard down) position AUTO. The next selections of the Standby Power Switch will energize their respective busses; AUTO position, AC power de‐energized and BAT switch ON: ‐ AC standby bus from the batteries through the static inverter ‐ DC standby bus from the batteries ‐ Battery bus from the batteries AUTO position with AC xfr bus 1 energized and BAT switch ON: ‐ AC standby bus from AC xfr bus 1 ‐ DC standby bus from the TR that provides the highest load ‐ Battery bus from TR 3 OFF position and the BAT switch ON: (STANDBY PWR OFF light illuminated) ‐ Battery bus from the batteries BAT position and the BAT switch ON or OFF: ‐ AC standby bus from the batteries through the static inverter ‐ DC standby bus from the batteries ‐ Battery bus from the batteries (Switched Hot Battery bus de‐energized when BAT switch is OFF) The amber STANDBY PWR OFF light illuminates together with the master caution and ELEC annunciator light when low voltage is detected on one of the next busses. ‐ AC standby bus < 100 VAC > 2 seconds ‐ DC standby bus < 17,5 VDC > 2 seconds ‐ Battery bus < 17,5 VDC > 2 seconds The STANDBY PWR OFF light only illuminates when the Battery bus has low power output with the BAT switch in the ON position.
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Fueling panel Single point pressure fueling or de‐fueling (or ground fuel transfer) is accomplished through this panel which is normally not a crew duty but certainly doesn’t harm if known. By the way SP 12 in FCOM 1 will guide you when refueling or tank fuel transfer is needed. For this explanation I use Kgs where Lbs is also possible and I don’t cover the Aux Tank version utilized on the BBJ. When refueling, the crews present knowledge on limitations should be used to stay within limits. ‐ Refuel the Main Tanks equally. (<453 Kg limit) ‐ Refuel the Main Tanks to full if there is >453 Kg fuel in the Center Tank. ‐ Fuel truck nozzle pressure should not exceed 50 PSI. (Placard states 55 PSI MAX) I remember a ‐7 PSI minimum when de‐fueling or else the hose collapses but isn’t a value with the B737 obviously and never used by flight crews. (except Flight Engineers) The fueling panel controls pressure, and manual refueling of the tanks and uses Hot Battery Bus DC power for operation when the door is opened. Refueling is possible with one of the next electrical power sources: ‐ External power connected with the system buses energized ‐ External power connected but no buses energized ‐ APU generator ‐ Battery power (Battery switch ON) Opening the Fueling Panel Door energizes the Refueling Power Control Relay by a magnet which allows Hot Battery Bus Power to the Fueling Panel. If this does not happen, the Fuel Indication Test switch should be used to the Fuel Door Switch Bypass position to accomplish the same. The Tank Fueling Valves are controlled by their respective switches, and a light will illuminate when the valve receives power. When there is also Fuel Pressure present!!, the valve will open so they open when the next conditions are met: ‐ Power on the Fueling Panel ‐ Fueling valve control switch OPEN ‐ Fueling valve solenoid energized ‐ Refuel pressure on the Fueling Valve ‐ De‐fuel suction on the Fueling Valve ‐ Fueling Valve Float switch not in the full position ‐ Preselected value (if applicable) is not reached
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When the tank reaches FULL, the indicators start flashing and power will be removed to the Fueling Valve which will close. On some newer Fueling Panels, the fuel load can be pre‐selected and will close the Fueling Valve when that level is reached. When the Fueling Valve solenoid fails, there is a possibility of manual operation of the valve by a Manual Override Plunger. Be aware that the FULL protection is inoperative during this action
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Brake accum mulator When bo oth the norm mal (B system m), and altern nate (A syste em) hydraulic brakes are inoperative, you still have (wh hen charged) residual brake accumulator pressurre available ffor braking. The brakke accumulattor is located d just aft of tthe right maiin wheel welll, accessible through an access panel un nderneath th he belly of the aircraft and has a direcct reading gaage at the aftt wheel well wall. It is a cylin nder free floaating piston ttype accumu ulator with a Nitrogen pre charge of 1 1000 PSI whiich is also indicated at the right forwarrd panel on tthe flight decck. When it indicates 100 00 PSI, there e is no or braking as this is just aa pre charge of the accum mulator. Minimum indicaated pre pressuree available fo charge p pressure should be 1000 PSI to enable the maxim mum possiblee amount of eemergency b brake applicatiions with ressidual pressu ure, when all (Hyd A & B)) brake presssure supply fails. The accu umulator is ccharged by B system presssure and routes through h the same h hydraulic tub bing as the norm mal brake sysstem applyin ng brake presssure througgh the individ dual wheel anti skid valve es toward tthe brake un nits. A fully ch harged brakee accumulato or is capablee of applying at least 6 full brake applicatiions but consider preven nting hard brraking as the anti skid vallves will relieef pressure to prevent a brake lockkup, thereby diminishing to less brake e application ns. Accumulator pressurre can also b be used to seet the parkingg brakes wheen there is no hydraulic ssystem operatin ng. When fullly charged, the accumulaator can hold d the parkingg brakes up tto approximaately 8 hours. A brake pressure relief valve is seet at 3500 PSSI to preventt damage to the accumulator which aalso restricts maximum p pressure whicch, when opened, closess again at 310 00 PSI. This is NOT the hyydraulic system p pressure relieef which are separate system relief vvalves set at 3500 PSI.
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Control column shaker Stall Buffet is identified when the critical Angle Of Attack is approaching a Stall condition, and Stall where that AOA is reached disturbing the airflow over the wing in such a way that it does not produce the required Lift. There are systems to warn the crew on approaching a Stall (buffet) condition by an indication on the PFD and a Stick Shaker system, and to prevent Stall by a hydraulic forward force on the Control Column (Elevator Feel Shift) and the Auto Slat System. Central of stall management are the two Stall Management and Yaw Damper computers (SMYD) which uses inputs on aircraft configuration such as Gear and Flap position together with Angle Of Attack and Mach number inputs. The PFD’s show the minimum safe operating speed (Vmin) related to the current configuration as red blocks on a black background at the lower inner part of the speed tape, indicating the speed where the stick shaker activates for normal stall warning. When the Stick Shaker has failed, the red blocks next to the speed tape are removed and a SPD LIM flag appears next to the speed tape on the PFD. (see right image) At first Flap retraction there is also an amber band visible above the red blocks that indicates minimum safe maneuvering speed Vmvr where on the approach the amber band is visible after Vref is entered. Note: Vref calculated by the FMC through your inputs are standard Vref values not affected by ice control systems so you have to add the 10 Kts to Vref when determining Vref “ice”. The Auto Slat System commands the LE Slats from the Extended position to the Full Extended position when the Flaps are selected at the 1, 2, and 5, (+ 10, 15, 25 depending on model) position, and the aircraft approaches the Stall region. When entering the Stall region, the Stick Shaker(s) are activated to warn the crew they are dangerously close to stalling the aircraft’s wings. The SMYD computer activates at Vmin and operates the Stick Shaker devices at the back of each Control Column where the Captains Stick Shaker uses 28 VDC Standby Bus power and the FO’s Stick Shaker 28 VDC Bus 2 power. The DC motors consist of unbalanced rings that shake their respective column when activated, and of course also the other column as they are interconnected underneath the flight deck floor boards. When close to a Stall, the SMYD computer commands the Elevator Feel Shift module (and actuator) to a ± four times higher nose down force to prevent further nose up motion (pulling on the Column) and transit into a Stall condition.
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The stall warning can be tested on the Aft Overhead Panel by separate test switches which activate the respective Stick Shaker motors. When the system does not operate (systems not malfunctioning) it could be that the test is performed within 4 minutes after AC power was selected as the SMYD computers uses this time for a self test. Another possibility is that one or both LE Flap panels have drooped off by the lack of B system pressure. They have no up lock as the Slats have and can move from the up position by weight and gravity when B system pressure has “leaked” away from the hydraulic lines. The test should not last >20 second as you might damage to the DC motor, and the stall warning test is inhibited when the aircraft is in flight.
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Wheel thermal fuse plugs. Thermal Sensitive Inflation Pressure Release Devices for tubeless aircraft wheels or easier thermal fuse plugs prevent violent explosion of the tires when maximum temperatures are exceeded. Note; when a wheel explodes (Main tire pressure can be as high as 205 PSI), it will happen sideways so NEVER approach a suspected overheated wheel from the side. Four thermal fuse plugs mounted in each inner wheel half (not visible because of the mounted brake units) prevent tire explosion caused by hot brakes. The plugs are designed to completely release the contained inflation pressure from a tubeless tire when brake generated heat causes the tire or wheel to exceed a safe temperature level and melt to release tire pressure at approximately 177°C or 192 °C depending on model. FCOM 1 SP 16 (adverse weather) talks about brake heat radiation and its negative effect on temperature levels that may melt the (thermal) fuse plug, deflating the respective tire. FCOM 1 PD section provides in Quick Turn Around Limit Weight tables that indicate the maximum Aircraft weight against OAT and PA which also need to be corrected for slope and wind. When this limit weight is exceeded you should honor the respective cool down period on the ground depending on the category brakes mounted on the aircraft, where after a check has to be performed to determine if the fuse(s) have not melted (deflated tire) before commencing takeoff. When a Brake Temperature Monitoring System (BTMS) is installed note the Brake Temp light, when illuminated honor the respective timeframe as above and check the fuse (tire) before takeoff. FCOM 1 PI section contains a Recommended Brake Cooling Schedule to determine the Adjusted Brake Energy (or indicated by the BTMS) which indicates what action is required covering a No Action, a Caution and Fuse Melt Zone area. The Caution and Melt zones indicate dangerously heated brakes and require safety actions such as: ‐ Caution on ground, delay takeoff ‐ Caution in flight, delay raising the gear ‐ Melt zone on ground, vacate runway, do not set parking brakes and wheel/brake replacement could be necessary. ‐ Melt zone in flight, delay raising the gear.
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Battery busses The 73 has three Battery Busses which in order of priority are: Hot Battery Bus Switched Hot Battery Bus Battery Bus The HOT BATTERY BUS is the most important DC Bus which normally receives power directly from the Main Battery Bus Bar or charger, in non‐normal conditions Main Battery power is supported by the Aux Battery in parallel. This Bus is the main power supply for all fire extinguishers and powers also other important DC components as shown in the image. The SWITCHED HOT BATTERY BUS becomes energized by selection of the Battery Switch to ON and receives power from the Battery Bus Bar/charger. This Bus powers some interesting components like the APU ECU (APU shuts down when ECU becomes de‐energized), Fwd airstair (STANDBY position) and the L & R ADIRU as an DC emergency power source. The BATTERY BUS receives power from the Main Battery/charger or from TR 3 and is energized by selecting the Battery Switch to ON, the Standby Power Switch to BAT or if the TR 3 has no output. The Battery Bus has the most and heaviest DC users of all Battery Busses.
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Electrical schematic
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Fuel schematic
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Hydraulic schematic
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Bleed schematic
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Air condition schematic
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Engine oil & fuel schematic
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Flight Mode Annunciations (FMA)
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INTENTIONALLY LEFT BLANK
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Power Sources (NG)
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