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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aware that it must not be a copy from any manual or else we interfere with a copyright that is also the reason why we do not publish schematics on the page.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 <100 ft. RA and AP selected, and when the EFS is not operational.
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Fuel Scavenge Jet Pump De fuel scavenge jet pump scavenges residual trapped fuel from the center tank to tank #1. Even at 0 Kgs indication there is still some residual fuel in the center tank. This fuel is trapped and cannot be sucked up by the scavenge line of the center tank boost pumps because of its elevated position. To be able to use this last bit of fuel, a center tank scavenge system is provided. To activate the system, next conditions need to be met; the LEFT FWD pump operating and tank #1 quantity lower than half full. (< 1990 Kgs) When the float type shutoff valve opens, it allows LEFT FWD fuel pump flow to create a negative pressure in the (non‐rotating parts) eductor type scavenge pump which in turn draws fuel from the center tank relieving it in tank #1. Of course this will create over time (the pump capacity is 100–200 Kgs/hr. (AMM)) a relatively small imbalance between the main tanks. The book says that the system continues to run for the remainder of the flight (can’t be shut off) but when you’ll remove the controlled condition (LEFT FWD fuel pump) also the jet pump stops operating. When the center tank is depleted, the scavenge pump draws air from the center tank to tank #1 which obviously does no harm to engine #1 operation. Note: the “dissolved air” story of fuel. When on suction feed with a high fuel temperature and a rapid decreasing pressure over the fuel, air bubbles (aeration) appear in the fuel possibly causing engines to run erratic or even flame out when sucked up though the bypass valve. Note: when both center tank fuel pumps are inoperative, fuel will be trapped in the center tank. There is no bypass valve provision for suction feed, also the left main tank quantity has to be below half full to even start the scavenge jet pump. Even so, the scavenge rate is insufficient to be used for emptying the center tank. Under these conditions you’ll use main tank fuel before the center tank is at required safe levels and a possible overstress of the wing roots arises. (>453, the main tanks have to be full and >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 < 50%. The A reservoir has a 20% standpipe to preserve fluid to the EMDP when a leak occurs at the EDP. The EDP is more likely to malfunction because of the engine gearbox mounted heavy design and higher capacity it puts out. (±4x) The B reservoir has a common standpipe for both system B pumps so when a leak occurs, fluid will drain the entire B reservoir until a 0% indication. In this case the B system cannot be pressurized anymore but the remaining 1.3 USG can be used for the PTU to operate the LE lift devices. A second standpipe at 72% preserves fluid to this level for both B system pump operation, in case a leak occurs while using the standby hydraulic system. Minimum quantity for the A & B reservoirs is 76% which triggers a white RF (refill) indication on the lower DU when on the ground and TE flaps are up, or no engines are operating. Besides that, when equipped with an update pin function to the lower DU on systems, there can also be a red dial indication when A or B quantities decrease to 0%, or increases to 106%. The pumps heated (case drain) cooling fluid return to the reservoirs, is routed through oil‐to‐fuel heat exchangers mounted on the bottom of the main tanks. To achieve enough cooling for on the ground operation, there should be at least 760 Kg of fuel in the tanks each.
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The APU Starter/Generator. The APU is started through a starter/generator and when on speed transfers to an AC generator. The start sequence of the APU starter/generator is determined by the Generator Control Unit (GCU) which receives power from the Switched Hot Battery Bus. That is the reason why the Battery Switch must be in the ON position (switched hot battery bus energized) to operate the APU. When switched OFF, the Switched Hot Battery Bus and ECU become de‐energized which in turn shuts down the APU immediately without the regular 1 minute cooling cycle. (trips the generator off line and closes the APU bleed valve to unload/cool the APU prior shutdown) Strangely enough power to the starter is provided by either the Battery (28 VDC), or Transfer Bus 1 (115 VAC). Both voltages are first changed/boosted to a whopping 270 VDC by the Start Power Unit (SPU), where after a Start Converter Unit (SCU) creates the 270 VAC which is needed to drive the starter/generator in the start mode. This signal lasts until 70% RPM where the SPU becomes de‐ energized and the APU becomes self‐sustaining and accelerates further to its operating RPM. When the APU RPM reaches ±95% the ECU commands the blue APU GEN OFF BUS light to illuminate as a signal that the APU generator can assume the electrical load. The AC generator consists of the same parts as the “regular” AC generator as described in an earlier post and can supply 90 KVA below 32,000 feet and 66 KVA at 41,000 because of APU load capabilities with low air densities.
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Landing Gear Transfer Valve The Landing Gear Transfer Valve has two ways of operation. The simplest is to transfer the nose wheel steering operation from its normal hydraulic system A, to the alternate hydraulic system B on the ground (only), by a switch on the left front (Capt) panel. The second way of operation (in flight) is a bit more complex as it has 3 conditions that needs to be met before the LG transfer valve moves from its normal hydraulic system A operation for gear retraction to the alternate hydraulic system B. 1. Engine #1, N2 below 50% 2. Landing Gear Handle in UP 3. Any gear NOT in the UP and locked position The PSEU is triggered by those conditions and moves the LG transfer valve to system B. Note that the PSEU light is inhibited from T/O thrust until 30 seconds after landing but DOES guard and operate the 737’s systems. Losing engine #1 stops the EDP (hydraulic system A) output so the only way to pressurize the A system is by means of the Electric Hydraulic Pump which puts out 4 times less volume than the EDP. This would result in 4 times slower movement of its components including a gear retraction which is an unwanted situation just after takeoff or on a go‐around with N‐1 conditions when you need to clean that configuration as fast as possible to decrease the massive drag by any extended gear. In that case the retraction is transferred from the A, to the B system so a normal fast retraction of the gear is achieved. The Power Transfer Unit (PTU) is a backup to the LE lift devices if the hydraulic system B EDP fails or has low output. It supports the B system electric hydraulic pump to operate the lift devices in a higher speed as it would be 4 times slower with just the EMDP. The PTU can also operate the lift devices when system B fluid is lost to a 0% indication, still holding ± 1.3 USG residual fluid in the reservoir to be used by the PTU.
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PTU The PTU operates when the next conditions are met: 1. Airborne and, 2. System B EDP pressure low (< 2350 PSI) and, 3. TE flaps less than 15° but not UP. If this occurs the PTU control valve opens, allowing system A pressure to operate the PTU hydraulic motor. The motor drives a hydraulic pump through a common shaft and uses the 1.3 USG from below the standpipe on the bottom of the B reservoir to operate the selected lift devices. Of course there are return lines back to the B reservoir from the PTU hydro motor and used devices which are not visible on common simplified (FCOM) schematics. Note that the PTU does NOT transfer fluid from A to B, and that the selected devices can be extended AND retracted by use of the PTU but will operate according the used pumps. (EMDP + PTU or PTU only)
Teaser . . . .how CAN you transfer hydraulic fluid from A→B or B→A?? A →B 1. EMDP's OFF. 2. Release parking brakes, deplete accumulator (<1800 PSI) 3. EMDP A, ON and apply parking brakes. 4. EMDP A, OFF and depressurize by control column movement. 5. EMDP B, ON and release parking brakes. (Sends the fluid back to system B) A →B 1. EMDP's ON. 2. EMDP B, OFF and depressurize by control column movement. 3. EMDP A, ON and apply parking brakes. (Uses fluid from system A) 4. EMDP B, ON and release parking brakes. (Sends the fluid back to system B) B →A 1. EMDP's OFF 2. Either FLT CONTROL to STBY RUD. 3. No1 thrust reverser OUT (uses standby hyd sys) 4. FLT CONTROL to ON. 5. EMDP A, ON. 6. Stow No 1 thrust reverser (using sys A)
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Wing Thermal Anti Ice (WTAI) Wing anti‐ice is provided for the inner three LE slats only and is preferably used as a DE‐icer. ANTI‐ ice would constantly heat up the LE thereby melting the ice crystals immediately, creating water “runback” over the wing and possibly freezing up on flight controls. Besides that it would have a negative effect on engine performance and fuel consumption. Note that use above FL 350 may cause a dual bleed trip off by the request of the amount of air also note that (ENG) anti‐ice is not required when < ‐40°C SAT. The outer slats are not de‐iced because the narrow outer slat cannot hold the hardware needed such as, a bleed manifold, telescopic tube and spray tubes. The wing is actually not producing much lift in that area anyway and they realized that some ice accretion on that part of the wing would not hurt too much. Eventually some drag and increase in stall speed occurs, not to forget that in case you use WTAI the stall warning computer remains set with increased speed logic. Where there is little cooling airflow over the LE on the ground, they are protected against overheating. First the engine bleed air is extra cooled through the pre‐cooler which allows tapped off fan air to extra cool the engine bleed air for maximum LE cooling on the ground. Second there is an overheat sensor (± 125 °C) which closes both WTAI valves when exceeded and opening up again at a predetermined value. During the design/test phase it turned out that ice does not accumulate on the empennage, mainly due to its position in relation to the engines causing hot air from the engines striking the empennage. Although some ice can build up in that area, it doesn’t have any adverse consequences (the stabilizer regularly changes the AOA and eventually shedding ice under the new conditions). The military version of the Boeing 737, the P8 Poseidon, does have a so‐called electro‐mechanical expulsion de‐icing systems (EMEDS) installed on the leading edges of the raked wingtips, horizontal and vertical stabilizers. The system is specially designed for the aggressive slow and low level cold weather mission assignments of this aircraft and does basically the same as a de‐icing boot but deforms the LE self by using low electrical current (28VDC and 25 Amps).
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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 temperature control Temperature control is achieved by mixing cool Pack air with hot Pack by‐pass air. The normal temperature range selection is from 18°C – 30°C through mixing cold air from the Packs with trim air for each individual compartment. The left Pack provides 20% to the Control Cabin and 80% to the mix manifold where the right Pack provides 100% to the mix manifold. The Zone/Pack controllers hold the various control electronics for the Cont Cabin and Passenger zones. The Cont Cabin has two controllers, a primary and a backup where the Passenger Cabin has only one controller for each area from either Zone/Pack controller. (see previous image) If both Cont Cabin controllers fail you’d get a Cont Cabin ZONE temp light with a Master Caution, if one fails they illuminate on recall. If a Passenger Cabin Controller fails the ZONE temperature light and Master Caution illuminates on recall and the two cabin requirement will be averaged. A ZONE temperature light also illuminate when there is an exceedance of duct temperature, the respective trim valve will close which can be reset by the reset button when cooled down. (select colder on that area) In the normal mode the Packs produce a temperature according the selection of the lowest temperature, the remaining zones use trim (hot) air required for their selected temperature. Unbalanced mode (Control Cabin trim air malfunction) The left Pack produces the selected Control Cabin temperature and the right Pack puts out the lowest Passenger Cabin selected temperature, the Passenger zone trim valves still operate. Unbalanced average mode (any Passenger Cabin trim air malfunction) The left Pack produces the selected temperature but the trim air valve still operates and the right Pack puts out an average of both Passenger Cabin selected temperatures. Single Pack operation and Trim ON results in normal temp control, with Trim switch OFF all trim valves close and the Pack averages the three compartment requirements. Trim switch OFF, all trim modulating valves are close and the left pack produces the selected Control Cabin temperature where the right pack produces an average of the Passenger Cabin selected temperatures. Temp selectors OFF will create a fixed 24°C output from the left and 18°C from the right Pack.
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Lavatory “fire protection”. I noted also B737 cabin crew “Likes” to our FB page, so I’ll try to aim a couple of subjects in that direction, of course also “need to knows” for flight crews. Let’s start with Boeing’s approach of “fire protection”, of course we’re discussing fire detection & extinguishing NOT protection ;‐) The lavatory smoke detection system needs 28 VDC from DC Bus # 1 to operate. The lavatory is equipped with a smoke detection system and a fire extinguishing system. In some 73’s you still find a “SMOKE” annunciator light at the P5 forward overhead panel but mostly there is no indication on the flight deck. In the cabin we find smoke detection indications through the next components: 1. Smoke Detector Unit As the name says, it’s a smoke detection and the unit is mounted against the ceiling of the lavatory. It has a green (power) light and a red (smoke detected) light, also an alarm horn will sound when smoke is evident for > 8 seconds. 2. Lavatory Call Light Located above the lavatory and is a Call/Reset Light that flashes amber when smoke is detected. 3. Master Lavatory Call Light At each EXIT locator light there are three indicator lights where a flashing amber Master Call Light indicates there is smoke detected in the lavatory in that respective area (fwd. or aft). 4. Attendant Control Panels (fwd.& aft) On these panels there are more options than just smoke detection as you can test the system here and detect FAULTS. When smoke is detected a red light flashes together with a flashing locater light that identifies the area where the smoke is detected and an intermittent horn is sound through the panel. The switches and lights on the panel are self‐explanatory, when a FAULT is detected during a test the failing detector is indicated through the location indicator. 5. Passenger Address (PA) system The PA sounds a repetitive high chime when smoke is detected.
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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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Antiskid The 73 is equipped with a system that prevent wheels from skidding (decelerating), thereby optimizing braking capabilities on any runway surface condition. An antiskid condition releases brake pressure to the affected wheel(s) which stops the skid condition when: • Uncommand deceleration. (Antiskid protection) • Wheel(s) stops instantaneously. (Locked wheel protection) • Landing with (parking) brakes ON. (Touchdown protection) • Hydroplaning To detect a wheel uncommanded deceleration, an electrical so‐called transducer is mounted underneath the hubcap of each wheel and is monitored by the Antiskid/Autobrake Control Unit (AACU). This signal is compared to information from both Air Data Inertial Reference Unit’s (ADIRU’s) and is also used for auto brake system wheel speed functions. The AACU controls the anti‐skid system and monitors for malfunctions which are indicated on the flight deck by an Antiskid Inoperative Light. An additional signal to the AACU comes from the parking brake system because the normal antiskid system returns (releases) hydraulic fluid through the parking brake valve. When the parking brake valve has a disagree with the lever (switch) the antiskid inoperative light also illuminates. Antiskid is provided during operating normal (system B), alternate (system A), and operation of the brakes with residual accumulator pressure. When in normal operation, antiskid is provided through 4 antiskid valves for each wheel separately and during alternate or emergency (accumulator) operation through 2 antiskid valves whereby the wheels are protected in pairs. To allow retract brakes to operate (Alternate brake pressure, system A) the antiskid system is de‐ energized when the gear retracts. Be aware that the antiskid system releases brake pressure, also during emergency (accumulator) operation which would reduce emergency brake applications when stepping on the brakes too 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. 35
In the past of “my field of experience” I saw a vertical stabilizer of a P3 Orion totally being sheared off like it was removed with a chain saw when it struck a wash rack when the aircraft has been swapped around by a twister at NAS Jacksonville and when a P3 hits a power cable at Pago Pago Hawaii. Be aware of the structural design of your aircraft!!
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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 translating sleeve has two deactivation points, installing two pins at these points prevents T/R deployment. Follow the “thrust reverser deactivation for flight dispatch procedure” from the current (AMM) manual to operate the aircraft with deactivated T/R’s.
An auto–restow circuit compares actual reverser sleeve position to the commanded position. When it determines an incomplete stowage or uncommand movement of the sleeves to the deployed position, the circuit commands to stow the T/R. When activated, the isolation valve remains open and the control valve is held in the stowed position until the thrust reverser is commanded to deploy.
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Tail Skid To protect the aft lower fuselage from over rotation damage the 737NG is equipped with a Tail Skid. It consists of a sort‐of shock absorber cartridge, a skid fairing and a skid “shoe”, where the last two parts are outside the fuselage where all other parts are inside. A light touch to the runway causes the shoe to wear off, indicated by 4 dimples on the shoe indicating the amount of wear and is an indication when the shoe needs to be replaced. When the shoe hits an object or an uneven part of the runway during the skid, the lower part of the shoe shears off as on the left image to indicate a tail drag but does not damage the skid fairing. A firm touch crushes the cartridge pushing the skid fairing inside the aft fuselage. The higher the force the further the skid disappears indicated by colored decals. If the green decal is still visible the skid is still “serviceable” but if the green decal disappears inside the fuselage, the red decal indicates that the skid must be replaced. When the “kiss to the runway” is more than firm, the skid disappears totally inside the aft fuselage and a safety pin (fuse pin) allows the cartridge to pivot inside (other than crushing) thereby protecting the aircraft structure against massive loads. There is also an option for a retractable tail skid that extends on take offs and landings which is under control of the Supplemental Proximity Sensing Electronic Unit (SPSEU) and operates with hydraulic system A pressure. The SPSEU commands the tail skid to extend if: • In the air for 2 minutes and, • Landing gear lever is DN and, • Either engine is running The SPSEU commands the tail skid to retract if: • On the ground 5 seconds or, • Landing gear lever is not DN or, • No engines operating
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Vortex generators. The 737 is equipped with several boundary layer control devices or, vortex generators (VG’s). They are mounted on the next locations: • the wings • the tail cone • the inner engine nacelle • the nose • the APU inlet door The vortex created by the VG affects the boundary layer on the respective surface behind the device by “pulling” air from outside, into the boundary layer. It creates an air swirl that draws air from above the boundary layer into this layer intensifying it and making it more compact. VG’s are mounted to slow, control or even prevent boundary layer separation. VG’s are used on the 737 wings to improve high Mach pitch characteristics beyond initial buffet and to lower stall speeds in the landing configuration. The (back swept) wing design creates a relative weak boundary layer where the outboard wings are more sensitive to initial flow separation. The purpose of the wing VG’s is to strengthen the boundary layer (especially with high AOA’s) and direct the airflow on the surface controls. On the tail cone, VG’s are mounted to separate the flow field from the horizontal tail thereby reducing drag, improving performance and reducing elevator vibrations. A Vortex Control Device (or nacelle chine) is installed on the inboard side of the nacelles. The engines are mounted relatively close to the wing which results in air disturbance at high angles of attack. To control the air flow at high AOA’s and slow speed, a Boeing invented VCD is mounted on the inner side of the engine nacelle. The created nacelle vortex is delayed with high AOA’s to support the airflow over the wing, increasing lift in those conditions. There are a number of VG’s mounted on the nose of the aircraft just before the windows. The general purpose is to reduce airflow noise by ± 3 – 4 Dbs. on the flight deck, directing the airflow away from sharp edges and corners of the windows. On the APU inlet door, there is a VG installed to improve high altitude starting of the APU. When the inlet door is opened during flight, the VG improves inlet ram recovery and thereby the pressure difference across the APU even to assist (electrical) starting.
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Window heating Window heat is provided to improve impact resistance (bird hit), ice buildup and defogging and should be selected ON at least 10 minutes before take‐off. Each frontal window (L & R SIDE & FWD) have an own Window Heat Control Unit (WHCU) which receive power from their respective 115 VAC Transfer Bus. This heating is accomplished by laminated glass and vinyl window layers with in‐between a conductive layer that allows electrical current to flow through it when heating is selected. This “gradual increasing” current flow creates heat by resistance in the layer towards a target temperature of ±43°C. The WHCU adjusts heating current in its operating range to prevent a thermal shock and reduces the current flow at higher ranges to prevent an over temperature. Window heat becomes active when selected to ON and the window temperature is < 37°C, indicated by a green ON light (or extinguished OFF light) meaning there is current flow through the conductive layer. When the window reaches the target temperature the WHCU interrupts electrical power and extinguishes the ON light (or illuminates the OFF light) which could already be the case when parked into the sun on a hot day. To make sure the system operates when needed, there is a POWER TEST switch that by‐passes the thermistor and sends electrical current through the windows. Be aware this action will bypass the control unit temperature regulation so when activated too long it could cause an overtemp condition in that window. Another test function is the OVHT TEST switch that simulates an overheat condition with the window heat switches in ON, indicated by amber OVHT lights on the control panel and extinguished ON lights (or illuminated OFF lights). The simulated overheat must be reset by selecting the window heat switches to OFF than back to ON. When a window “overtemps” at values higher than 62°C, the WHCU interrupts power to the affected windshield and illuminates an OVERHEAT indication together with an extinguished ON light (or Illuminated OFF light). This condition can be reset by selecting the affected window to OFF and back to ON when allowed sufficient cooling first (2 – 5 minutes according the QRH). When window heat is inoperative prevent speeds above 250 Kts at altitudes below 10.000 ft. to minimize the effect when a bird strike occurs on the window.
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Wing& Body Overheat On the P5 overhead bleed panel, there are 3 caution lights mounted that warns you for an overheat condition in the bleed system. Two of them are INSIDE the manifold and considered “safe” as they have a reset function after the overheat condition is corrected, the BLEED TRIP OFF and PACK light. The third light is the WING‐BODY OVERHEAT light that indicates an over temperature OUTSIDE the manifold and is considered “not safe”. This indication determines an overheat in the area where the duct is located indicating a duct leak or worse, a duct rupture. The areas covered by the left indication are: • Left engine strut (154°C) • Left inner leading edge (154°C) • Left air condition compartment (124°C) • Keel beam area (124°C) • APU bleed duct area (124°C) The right indication covers the engine strut, leading edge and air condition on the right side. When a wing & body overheat condition (leak) exists, use the current QRH to determine the location and isolate the leak by selecting a combination of pneumatic system related switches to OFF. When located and isolated, the temperature drops and extinguishes the indication knowing that the overheat has disappeared but not the cause of the indication. When the correct QRH procedure was followed the overhead condition should clear as the source has been removed somewhere during the procedure. If not . . . the QRH doesn’t suggests steps beyond the procedure so use common sense to fly out of this condition. The system has a test switch located on the P5 bleed panel to test the continuity of the sensing loops. The test starts when the button is pushed > 5 seconds and indicates the same as in an overheat condition by the next amber indications: • WING‐BODY OVERHEAT • AIR COND annunciator • MASTER CAUTION
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Horizontal Stabilizer Trim. One of the important features related to pitch and load balancing is the movable horizontal stabilizer trim control (stab trim). A control jackscrew moves the leading edge of the horizontal stabilizer as a trim in order to achieve this goal and can be operated: • Manually by two trim wheels which operate the stabilizers gearbox and jackscrew through cables and cable drums. • Electrically either through yoke trim switches or Auto Pilot command to the stabilizer electrical trim actuator. o AC power – AC Transfer Bus 2 o DC control – DC Bus 2 Electrical movement of the trim actuator by either the yoke switch‐ or the Auto Pilot will backdrive the trim wheels on the control stand. When the handle on the wheels are extended during electrical operation, this can injure the operators leg/knee. Extreme UP of the leading edge is restricted at 4.2°, and DOWN at 12.9°. Indication in “Units” is mechanically provided on the control stand through a flexible cable that is driven off the trim control mechanism underneath the flight deck floor. As reference, the 0° neutral position equals 4 units on the trim position scale. Stabilizer Trim Cutout switches are located on the control stand in order to interrupt either control column switch–, or AP electrical power toward the trim motor when an uncommanded movement or “runaway” trim occurs. A Stabilizer Trim Override switch is located on the aft electronic panel in case a counter movement of the trim is required opposite of the control column movement. When not in OVERRIDE, a mechanical control column actuated stabilizer trim cutout switch will interrupt electric power to the trim motor when attempting to trim opposite of control column or AP commanded force. (column nose DOWN vs. trim UP or vv) The override switch can also be used to by‐pass the control column actuated stabilizer trim cutout switches in the event both (yoke switch or AP) fail in the open position, to be able to operate the stab trim. Electrical movement by the yoke switches can vary between high speed (0.4 unit/sec) when the flaps are NOT UP and low speed (0.2 unit/sec) when the flaps are fully UP. When the trim is under AP control high speed is 0.27 unit/sec while low speed with the flaps UP is 0.09 unit/sec.
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Display Electronic Units. There are two DEU’s in the Common Display System (CDS) which are located in the E & E bay that receive data from the B737’s aircraft systems and avionics. This data is converted into a video signal that is send to the six Display Units (DU’s) on the flight deck. Both DEU’s provide data toward all six DU’s but are differently powered. DEU 1 is powered through the 28 VDC Standby Bus and DEU 2 through the 28 VDC Bus 2. They both have a “hold‐up power” from the Hot Battery Bus which is used to supply power to the DEU’s during power surges of maximum 2 seconds or else the DEU powers down. Both DEU’s “crosstalk” to compare critical data and when there is a difference, this could create an amber CDS FAULT indication as described below. The same “split” is made for powering the components that distinguish DU operation when powering up the aircraft. I mean when the Battery Switch is selected ON, DEU 1 is powered through the DC Standby Bus but also both Captains–, and Upper DU’s, as well as the Captains EFIS control panel. Note that it takes ± 90 seconds to get display because the DEU has to become operational. When the DC Bus 2 becomes powered the same applies for the First Officers side. An (undispatchable) amber CDS FAULT displays on both PFD’s when there is a DEU operational failure on the ground and one engine operating. When both engines are operating or when in the air the CDS FAULT message is replaced by a DISPLAY SOURCE message. The DISPLAY SOURCE also shows when one DEU is selected (ALL ON 1 or 2) to provide all six DU’s with data. Note: Switching too fast between SOURCE selections can create a possible incorrect data display, use a 1 – 2 second interval when switching between the SOURCES. A (dispatchable) white CDS MAINT indication tells us that there is a partial data input malfunction on a DEU when on the ground and 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 steering lockout Yesterday I answered a question to one of our followers and want to share this (obvious) info. When towing or pushing the aircraft with a tug and tow bar, system A pressure has to be removed from the nose wheel steering system. It prevents unwanted dangerous movements of the tow bar injuring people or damaging the nose wheel steering system. It is done either by switching both A pumps to OFF or by use of the towing lever and lockout pin to depressurize the nose wheel steering system. Moving the lever to the towing (fwd.) position moves the towing shutoff valve to such a position that it shuts off A hydraulic pressure to the steering valve and connects both sides of the steering valve to each other, and return. This allows the nose gear to rotate freely to a maximum of 78° indicated by red stripes on the lower side of the fuselage. Moving the nose wheel beyond the stripes requires disconnecting the torsion links or even the taxi light wiring. Most companies wants you to always switch the A pumps OFF even if the lockout pin is installed as a safety measure. This is done to prevent miscommunication with the multi‐cultural ground observers related to pin installation. Just thinking about the system as an engineer . . . . . remember the NOSE WHEEL STEERING switch. When it is selected in Alternate on the ground, it moves the Landing Gear Transfer valve to select B system pressure to nose wheel steering!!!
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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 weather radar actually enters the wind shear scanning mode at 2,300 feet AGL to provide time for the system to power up (if necessary) and update the displays before the aircraft reaches the 1,200 feet AGL level. TEST During the test: • The R/T transmits a few pulses to let the BITE monitor for correct operation • The R/T makes a test pattern and sends it to the DEU to show on the ND’s • The R/T sends test messages, mode, gain and tilt information to the DEUs to show on the ND’s • WXR test pattern shows on ND’s. • The test pattern shows until another mode on the WXR panel or EFIS panel is selected.
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Dissolved air A fuel phenomena within high altitude aviation is called “dissolved air” or “aeration”. It is a result from the highly aerated condition of fuel caused by rapidly decreasing tank pressure during climb, allowing entrapped air in the fuel to expand. Reduced air pressure above the fuel surface promotes the release of dissolved air from the fuel. Air released from the fuel can have degrading effects on the performance and safe operation of a fuel system. 2 Main tank fuel pump LOW PRESSURE’s during rapid climbout can cause thrust deterioration or even an engine flameout on the affected engine at higher altitudes (> ± 13,000 feet). (QRH, FUEL PUMP LOW PRESSURE) This altitude varies with the prevailing fuel temperature in the tank (the higher the fuel temperature, the lower the altitude at which the gradual power loss occurs). Once pressure has stabilized and excess air has escaped from the fuel, loss of both fuel boost pumps has no effect on engine operation with maximum power settings at altitudes up to above 30,000 feet. The time required to stabilize the fuel from this highly aerated condition cannot be determined exactly, since it is a function of both rate‐of‐climb and fuel temperature. Solution to the problem is level off and let the engines stabilize at altitude or pressure feed the engines as suction feed increases the aeration effect and ads in the possibility to ingest aerated fuel into the engine feed line. Fuel stabilization should occur after a few minutes of stabilized cruising operation or back on pressure feed.
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Frangible fittings Frangible fittings are mounted in the rim of the main wheel wells to prevent a rotating blown tire to enter the wheel well. If it shears, only that side will freefall back down by relieving Landing Gear Actuator up pressure overboard. (4 green and two red indications) Note that we do have retract brakes through the alternate brake system (hydraulic system A) but when a tire blows there is a good chance that the brake lines will be substantially damaged causing the retract brakes not to work. Retract brakes (and nose wheel snubbers) are mounted to stop the wheels from rotating, hanging in their uplocks. A high speed rotating wheel would cause tremendous precession forces to the structure during a turn that’s why they are stopped after retraction.
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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. In the past of “my field of experience” I saw a vertical stabilizer of a P3 Orion totally being sheared off like it was removed with a chain saw when it struck a wash rack when the aircraft has been swapped around by a twister at NAS Jacksonville and when a P3 hits a power cable at Pago Pago Hawaii. Be aware of the structural design of your aircraft!!
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