b 737 Theory
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B73 37T 7Th heo ory ry
Bo oein ng 737N 7 7NG S stem Sys ms s
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Foreword: This booklet describes systems published in our Facebook pages: About This FB page is to interact throughout the B737 community and has NO direct link to any user company. THE CONTENT SHALL NOT BE USED FOR ACTUAL OPERATION OF THE AIRCRAFT. The administrator has NO RESPONSIBILITY to the content written on these pages. Description Creator: Ferdi Colijn Administrators: Ferdi Colijn (B737NG Type Rated) Maarten van Klaveren (B737‐900ER First Officer) Bert de Jong (Flight Engineer) B737Theory March 24 The goal of this FB page is to expand B737 theoretical knowledge among users and we try to achieve that by expanding the amount of visitors aiming for interaction. There rest no copyright on our stories but we rather see you recommending us on your private FB pages iso sharing the posts. Also feel free to "donate" your experiences and stories on B737Theory and drop us a line by sending a message. We will evaluate and post them in time. Be aware that it must not be a copy from any manual or else we interfere with a copyright that is also the reason why we do not publish schematics on the page. Thank you
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Contents: Foreword: ................................................................................................................................................ 2 APU .......................................................................................................................................................... 6 Auto Slat System...................................................................................................................................... 7 Engine Electronic Control (EEC) ............................................................................................................... 8 When things go wrong and beyond basic systems knowledge ............................................................... 9 Engine fire detection ............................................................................................................................. 11 Feel Differential ..................................................................................................................................... 12 Fuel Scavenge Jet Pump ........................................................................................................................ 13 Fuel valves ............................................................................................................................................. 14 AC Generator ......................................................................................................................................... 15 Isolation valve ........................................................................................................................................ 17 Manual gear extension. ......................................................................................................................... 18 Mechanical pressure relief valves. ........................................................................................................ 19 Nitrogen Generating System ................................................................................................................. 20 Outflow valve. ....................................................................................................................................... 21 Flight Control “Breakaway” Devices ...................................................................................................... 22 Pack & pack control ............................................................................................................................... 23 Recirculation fans .................................................................................................................................. 24 Hydraulic Reservoirs .............................................................................................................................. 25 The APU Starter/Generator. .................................................................................................................. 26 Landing Gear Transfer Valve ................................................................................................................. 27 PTU ........................................................................................................................................................ 28 Wing Thermal Anti Ice (WTAI) ............................................................................................................... 29 B737 Yaw damping ................................................................................................................................ 30 Zone temperature control ..................................................................................................................... 31 Lavatory “fire protection”. .................................................................................................................... 32 Center tank boost pumps ...................................................................................................................... 33 Antiskid .................................................................................................................................................. 34 Leading Edge Flaps ................................................................................................................................ 35 Thrust Reverser ..................................................................................................................................... 37 Tail Skid .................................................................................................................................................. 39 Vortex generators.................................................................................................................................. 40 Window heating .................................................................................................................................... 41 Wing& Body Overheat ........................................................................................................................... 42 Horizontal Stabilizer Trim. ..................................................................................................................... 43
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Display Electronic Units. ........................................................................................................................ 44 Proximity Switch Electronic Unit ........................................................................................................... 45 Nose wheel steering lockout ................................................................................................................. 46 Weather radar ....................................................................................................................................... 47 Dissolved air .......................................................................................................................................... 49 Frangible fittings .................................................................................................................................... 50 Rudder(vertical stabilizer) load reduction ............................................................................................. 51
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APU The APU is a constant speed (± 49.000 RPM) gas turbine engine that can supply AC power and pressurized air. The starter/generator is powered from either directly the main battery (28VDC) or transfer bus 1 (115VAC) where either source is converted into 270VDC for starter operation. At 95% starter operation reverses to a 90 KVA generator, indicated by the blue APU OFF BUS light. (90 KVA until 32.000 ft. and 66 KVA until 41.000 ft.) Starter sequence is automatically determined by the Electronic Control Unit (ECU) that needs the battery switch to be in the ON position to energize. The APU can be used for air and AC power until 10.000 ft., just air to 17.000 ft. and just AC power until 41.000 ft. That is also the maximum starting altitude although recommended at 25.000 ft. Air takes the biggest performance from the APU as it takes air from the load compressor which is mounted on a common shaft with the combustion compressor. The more air taken in, the lower the performance of the APU. That is why there is a restriction in altitude use, especially with air and when the demand is large (high EGT), air use is squeezed by IGV’s toward the load compressor. When on suction feed the APU draws fuel from tank #1 and when operating for an extended time select a fuel pump to pressure feed which extends the lifetime of the APU. The ECU protects the APU and shuts down with a low oil pressure, overspeed or when a FAULT light illuminates. The latter represents more than just the foregoing, including ECU failure, loss of DC power, APU fire, overtemp (during start), high oil temp and many more. The start limit is 2 minutes and a FAULT light illuminates when the start is aborted through a protection or when the generator malfunctions. A blue MAINT light illuminates when oil quantity is low or a generator malfunction occurred, the APU is still allowed to operate. APU compartment and oil cooling is accomplished by exhaust air used as an educator to draw outside air into the compartment from an inlet just above the exhaust. When the APU is stopped by placing the switch to OFF, the ECU determines a cooling cycle of 1 minute before the APU actually stops. The cooling cycle closes the APU BAV and trips the generator OFF line. By doing so it reliefs the APU from load and decreases the EGT preventing so called cooking of the nozzles. (residual fuel forms carbon on the hot nozzles which can affect the flame pattern) Delay switching the Battery to OFF to 2 minutes after selecting the APU to OFF, this allows the inlet door to close. The door closes when the APU decelerates to ± 30% to prevent the inlet duct to collapse. The 1 minute is by‐passed when the APU shuts down through a malfunction, the Fire Switch is activated or when the Battery Switch is selected 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 BATTERY DISCHARGE is probably caused by the a (excessive) main battery discharge by powering the Battery Bus as also the DC 2 system (TR 2 & TR 3) were not powered anymore and illuminates when a battery output conditions exists of: • Current draw is more than 5 amps for 95 seconds • Current draw is more than 15 amps for 25 seconds • Current draw is more than 100 amps for 1.2 seconds. Mind you, normally when Transfer Bus 2 is de‐energized the Transfer 3 Relay would switch TR 3 to Transfer Bus 1 which obviously 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 Differential The FEEL DIFF PRESS indication on the flight control panel can illuminate in the following cases. (The feel system simulates “actual feel forces” at the control column from the hydraulically supported elevator panels) 1. The first one is related to a differential of A & B hydraulic pressure to the elevator feel system. When either hydraulic system pressure drops > 25% related to the higher pressure, the FEEL DIFF PRESS light illuminates on the flight control panel with a 30 second delay. The 30 second delay prevents the light from “flickering” when pressure drops in either system by a high demand such as gear selection. 2. The second is related to the dynamic air pressure supply to the Elevator Feel Computer. It receives dynamic pressure from the two pitot tubes mounted on either side of the vertical stabilizer. When the computer receives an erratic signal it’d be the same as the pressure drop and the light illuminates. (failed probe heater and icing conditions) 3. The third is related to the Stall Management and Yaw Damper (SMYD), and a so called Elevator Feel Shift module (EFS), which creates a ±4 times higher forward control column force when approaching the stall region. This force uses a reduced system A pressure and when this reducer fails, opening prematurely providing a higher than normal A system pressure to the feel actuator, the FEEL DIFF PRESS also illuminates after 30 seconds. Note on the last system, it’s inhibited 726, 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 the image shown protections in the CDU will de‐energize the GCR thereby de‐energizing the exciter field, which de‐energizes the generator. This de‐energizing GCR also occurs when the generator switch is selected OFF.
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Isolation valve The isolation valve separates the left, from the right side of the bleed manifold. It is powered from AC Transfer Bus 1 but also can be manually opened/closed by a control lever, accessible in the left air condition bay. Because it’s AC power* it will fail in the selected position when power is removed. When the Isolation switch is in the AUTO position the valve opening relies on the so‐called “corner switch” positions. They are the Pack and Bleed switches, when all these switches are NOT in the OFF position the isolation valve is closed. On the other hand if any corner switch is selected to OFF the Isolation valve opens in the AUTO selection. When a Pack switch is OFF, the Isolation valve opens to create equal performance of the engines. When a Bleed is selected OFF the Isolation valve opens to allow air from either side of the manifold to be used for the off side WTAI. Note the isolation valve logic is related to switch position so a tripped pack or bleed will not open the Isolation Valve when in AUTO. After flight the Isolation valve should be selected OPEN just in case you need to battery start engines when there is no APU or external electrical power available. The ground air connection is located on the right side of the manifold close to engine #2. When N2 >20% there is no personnel allowed in the vicinity of the turning engine so we have to start engine #1 first. When this would be a battery start you’ll need the isolation valve to be open, so when you removed AC power with the isolation valve switch OPEN, the valve is still in the open position. * A general rule for electrical power is; “AC lies, DC dies”. This is a nice thing to know also for analog instruments, an AC powered instrument stays where it lost power and a DC powered instrument will drop off to zero.
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Manual gear extension. Let’s have a look at this Non Normal procedure and its components. When the gear is UP and the LG lever in the OFF position, hydraulic system A pressure is removed from the uplines to the actuators which causes the three struts to “hang” in their respective uplock. This is also the preferred position of the LG lever during a manual extension attempt because of the depressurized hydraulic lines. When the gear (all or any) does not extend after a down selection, follow the QRH procedure in an attempt to lower the gear. Manual extension of the gear is accomplished by pulling the three “T” handles, accessible through the Manual Gear Extension Access Door just behind the FO seat on the cockpit floor. The need for this Non Normal procedure could be caused by: • Disrupted electrical signal to the LG selector valve • No system A hydraulic pressure available • LG lever stuck in the UP or OFF position When opening the Manual Gear Extension Access Door, a “door open” micro switch commands the LG selector valve electrically down regardless of the LG handle position. This action activates the LG selector bypass valve which connects the hydraulic lines to return so the manual down selection does not hydraulically restricts (locks) the actuators down capability. This also prevents the LG to retract when the door is not flush closed after take‐off and selected UP. This procedure is covered in the QRH by the LG disagree procedure with the LG handle UP and all red and green indicator lights illuminated, telling you the gear is down and locked but not in the selected position. When you’d pull any (or all) “T” handle it simply releases the uplock by cable action where after the respective gear free‐falls down, supported by gravity (weight) and airflow to the extend position. When the gear is fully down, the downlock “bungee” springs will hold the downlock struts in an over centered locked position. Normally this is accomplished by a downlock actuator but with the absence of system A pressure, the springs enforce a mechanical downlock which is indicated by (6) down and locked green lights. By the way, there are 6 green lights as a redundant indication. Neither gear is visible on the NG and the double green lights for each strut will give a backup for the down indication.
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Mechanical pressure relief valves. There are three mechanical adjusted pressure relief valves on the 737. Positive safety pressure relief is accomplished by 2 mechanical adjusted pressure relief valves, located on each side of the outflow valve. They are totally independent of the pressurization system and prevent the inside/outside pressure to exceed +9.1 PSID in the event of a pressurization system/outflow valve malfunction. (stuck closed outflow valve) The fuselage airframe structure cannot withstand large negative pressures and is protected for that at a very low value. The negative pressure relief valve is located at the right lower side of the fuselage just fwd. of the outflow valve. This spring‐loaded door is also not depending on the pressurization system and adjusted at just a –1.0 PSID value. This will prevent the aircraft to collapse when the inside/outside pressure becomes negative for example during a (very) fast descent.
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Nitrogen Generating System Following two Boeing 737CL explosion investigations in Asia (and others including the B747 TWA 800 midair explosion), a protection was developed by Boeing to minimize explosive vapors in the center tank. The 737 explosions were caused by trapped fuel high temperatures due to radiant heat from the Packs under the tank which formed highly explosive vapors. The fuel was ignited by the center tank fuel pumps which were still running with an empty center tank. Early days center tank fuel pumps did not had an automatic shut off with LOW PRESSURE as the newer modified ones that shut down after ±15 seconds of LOW PRESSURE. This is also the reason that someone has to be on the flight deck when a center tank pump is running as by the FCOM, the book does not cover explicit modifications to each aircraft. This protective device (NGS) divides Nitrogen from Oxygen by a separation module and leaves Nitrogen enriched air (NEA) in the center tank to a level which will not support combustion. The oxygen level is decreased by the NGS to ±12% which is sufficient to prevent ignition. The NGS has only an indication available in the right main wheel well next to the APU fire control panel, so it has no visible clew for crews of its operation during flight. Indications are: • OPERATIONAL (green) • DEGRADED (blue) • INOPERATIVE (amber) The nitrogen generation system gets bleed air from the left side of the pneumatic manifold where after its cooled, driven through the separation module and directed to a flow valve into the center tank. The NGS operates automatically only in flight and shuts down in the next conditions: • Either engine is shut down in flight • Fire or smoke detection in any compartment • Left Pack overheat
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Outflow valve. To stay in line with the previous post, let us look at this pressurization component of the 73. The outflow valve restricts/regulates the flow of conditioned air overboard, thereby creating a pressurized environment in the aircraft. The valve is located at the aft lower side of the fuselage and has raked edges for noise reduction purposes. The valve is moved by a common actuator which can be operated by either of the three outflow valve electro motors. Two motors are operated by the pressure system controllers and one is directly operated by a switch when in Manual operation. Automatic control is accomplished by means of 2 Cabin Pressure Controllers (CPC’s) which alter control each flight or when a malfunction occurs on the operating controller. A third way of controlling the outflow valve is by a manual toggle switch on the pressurization panel. The switch is spring loaded to neutral and has three positions, CLOSE – Neutral – OPEN. The outflow valve indicator shows the actual position of the outflow valve in all modes of operation provided the Battery Bus is powered through the PRESS CONT IND C/B. Electrical power to the three electro motors is provided by: • AUTO mode 1 electrical power to the auto electro motor 1 is supplied by the 28 VDC Bus 1 through CPC 1. (PRESS CONT AUTO 1 C/B) • AUTO mode 2 electrical power to the auto electro motor 2 is supplied by the 28 VDC Bus 2 through CPC 2. (PRESS CONT AUTO 2 C/B) • MANUAL mode electrical power to the manual electro motor is supplied directly by the 28 VDC Battery Bus. (PRESS CONT MAN C/B) A mode selector is used to determine the operation of the outflow valve, either AUTO, ALT(ernate) or MAN(ual). The outflow valve receives a closed signal when the cabin altitude reaches 14.500 feet in the AUTO mode of operation so it is not affected through the MANUAL mode. Just for the “mind set” when at a high altitude and a pressure loss, you’d have to close the outflow valve to increase pressure in the aircraft which results in lowering cabin altitude. Aircraft control override 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 control There are two Packs activated by an AUTO/HIGH selection that individually has two airflow directions, one that goes through a three stage cooling cycle (2 air to air heat exchangers and an expansion turbine) and one that bypasses the cooling machine and its components. The two flow directions are mixed at the output of the expansion turbine of the cooling machine. Air that enters the Packs through the Pack Flow Control and Shutoff valve is at ± 212°C and is conditioned and cooled to a mixed minimum Pack output of ± 18°C as set the lowest on the zone temperature control selectors. (auto zone temperature range is 18°C – 30°C)When these selectors are all in the OFF position, the left Pack puts out a fixed 24°C and the right Pack 18°C. There are two combined Zone/Pack controllers that control the required output temperature of each Pack. These two Pack Controllers have an auto “on side”, and a standby “off side” control, the latter takes over if an auto controller fails. In this case a PACK OFF light illuminates on recall together with a Master Caution light. When both Pack Controllers fail, a Pack OFF light illuminates with a Master Caution light, the packs will still operate until a temperature exceedance occur. When a Pack becomes overloaded by the demand of cool air, a PACK trip off light illuminates with a Master Caution light and the Pack Flow Control and Shutoff valve closes shutting down that Pack. When the Pack cools down and the light extinguishes, the Pack can be reset by the reset button on the Bleed panel. To prevent this condition from re‐occurring select a higher temperature to “unload” that Pack by demanding less cold air from the cooling machine bypassing it. A Pack automatically provides a high airflow when the other Pack is selected to OFF provided the aircraft is in the air with flaps up. The other conditions require engine performance and inhibits the automatic high flow. Note: the image is just a simplified flow and pack component, and controller image to illustrate the flow through the pack and the components in both controllers.
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Recirculation fans The recirculation fans are located under the cabin floor on the forward cargo compartment’s aft bulkhead. The purpose of these fans is to re‐use air drawn from the cabin and distribution compartment back into the mix manifold. Doing so there is no need for air from the Packs, thereby relieving the Packs from producing conditioned (cool) air improving engine performance. The left recirculation fan circulates air back into the mix manifold from the distribution compartment underneath the cabin floor (mix manifold/fan area), the right recirculation fan from the passenger compartment. When a higher amount of fresh air is needed from the packs, the recirculation fans are automatically shut down under several conditions with the recirculation fans selected to AUTO, and the isolation valve selected to AUTO or OPEN: On the ground using engine bleed air: Left RECIRC FAN shuts down when both Packs are selected to high flow On the ground using APU bleed air: Left RECIRC FAN shuts down regardless of Pack selection In flight using engine bleed air: Left RECIRC FAN shuts down when either Pack is selected to high flow Both RECIRC FANS shut down when both Packs are selected to high flow In flight using APU bleed air: Both RECIRC FANS shut down regardless of Pack selection Reading the first part it makes sense that the left fan (distribution compartment) shuts down first as this area heats up by the several operating components. (my personal point of view)
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Hydraulic Reservoirs The 3 hydraulic fluid reservoirs are located in the front of the main wheel well. They are pressurized from the bleed manifold to supply positive fluid to the pumps, preventing cavitation and foaming.The standby system reservoir is pressurized through the B reservoir.These pressures (45 – 50 PSI) can only be checked on 2 gages mounted on the forward main wheel well bulkhead. Quantity of the A & B reservoirs is displayed directly through gages on the reservoir by a float type transmitter which also sends a signal to the DEU’s for display on the lower DU. The standby system reservoir only has a low quantity switch, which displays the STANDBY HYD LOW QUANTITY light on the flight control panel when
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