Safety Investigation Report 2018:1 Factual Information/1.6/1.6.8 Aircraft Systems
SAFETY INVESTIGATION REPORT MH370 (9M-MRO)
1.6.8 Aircraft Systems Description
Most of the electronic equipment on the aircraft are mounted on equipment racks in the various equipment centres.
The Main Equipment Centre (MEC) contains most of the electronics equipment on the aircraft. The MEC is below the passenger cabin, rear of the nose wheel well and forward of the forward cargo compartment. Access to the MEC is possible on ground or in flight. The equipment in the MEC includes electronics for these functions:
- Information Management
- Generator Control
- Transformer Rectifier
- Flight control and autopilot
- Environmental control
- Recording
- Navigation
- Communication
- Cabin Management
- Weight and balance
- Air data
- Inertial data
- Warning
- Proximity sensing
- Engine control
- Electrical Load Management.
The Forward Equipment Centre is forward of the nose wheel well and contains the two weather radar receiver/transmitters. Access to the Forward Equipment Centre is through the access door forward of the nose landing gear or through the MEC.
The passenger compartment above the Door 3 cross-aisle at station 1530 on the left of the aircraft centre line contains the satellite communication equipment.
A rack in the passenger compartment above the rear galley at station 2100 on the right side of the aircraft contains the flight recorders.
There are also equipment racks adjacent to the forward, aft and bulk cargo doors. The forward cargo racks contain the primary flight control, actuator control, radio altitude, fuel quantity and cargo handling electronics. The aft cargo racks contain the HF communication, brake and tire and main gear steering electronics. The bulk cargo racks contain the APU battery and charger.
1) Air Conditioning and Pressurisation
The aircraft has two air conditioning systems divided into left pack and right pack. Engine bleed air provides the pneumatic source for air conditioning and pressurisation.
There are two electronic Controllers, each of which can provide both pack and zone control. Each Controller has two channels that alternate command cycle. Cockpit and cabin temperature selection is monitored, and the Air cycle machine and temperature control valves will be commanded to deliver temperature conditioned air to the various cabin zones.
Conditioned air is also used for electronic equipment cooling. This is supplied through a series of pneumatic valves with supply and exhaust fans. Exhaust air from the equipment cooling flow is routed to the forward cargo and used for forward cargo compartment heating.
Two cabin pressure Controllers regulate the aircraft pressurisation and command the pneumatic system. System operation is automatic and works in conjunction with the forward and aft outflow valves that are used for pressurisation. The outflow valves can also be manually operated from the cockpit by switches on the overhead panel.
Loss of cabin pressure will be indicated to the flight crew by a Cabin Altitude warning message on the Engine Indicating and Crew Alerting System (EICAS) display together with the associated aural warning.
2) Autopilot Flight Director System
The autopilot is engaged by operation of either of two A/P pushbutton switches on the Mode Control Panel (MCP) located on the glareshield panel (Figure 1.6C [below]). Once engaged the autopilot can control the aircraft in various modes selected on the MCP. Normal autopilot disengagement is through either control wheel autopilot disengage switch. The autopilot can disengage if the flight crew override an autopilot command through the use of the control column, control wheel or rudder pedals (when the yaw axis is engaged for approach).
Figure 1.6C - Autopilot Mode Control Panel
The autopilot can also be disengaged by pulling down on the A/P Disengage Bar on the MCP. The autopilot will also disengage automatically for failures of systems on which it relies upon for specific operations. The Autopilot Flight Director System (AFDS) consists of three Autopilot Flight Director Computers (AFDCs), one MCP, and six backdrive actuators (one each for the Captain’s and First Officer’s control column, control wheel, and rudder pedals). The left and right 28V DC buses power the left and right AFDCs, respectively and the MCP while the 28V DC battery bus powers the centre AFDC.
Emergency power from the Ram Air Turbine (RAT) generator does not power these busses and as a result the autopilot will not function with RAT electrical power.
- a) Take-off Mode
The Take-off (TO/GA) mode controls roll and pitch during take-off. Also, the Thrust Management Computing Function (TMCF) controls thrust during take-off. Turning a flight director on while the aircraft is on the ground, or activating either TO/GA switch while on the ground, will engage Take-off mode.
- b) Roll Modes
The following AFDS roll modes are available during climb, cruise and descent (Figure 1.6D [below]):
- i) Lateral Navigation
Pushing the Lateral Navigation (LNAV) switch arms or disarms the LNAV mode. The commands come from the active Flight Management Computing Function (FMCF) when there is a valid navigation data base and an active flight plan.
- ii) Heading Hold/Track Hold
Pushing the Heading Hold (HDG HOLD)/Track hold (TRK HOLD) switch selects Heading or Track hold. In this mode, the aircraft holds either heading (HDG) or track (TRK). If the HDG/TRK display on the MCP shows TRK, the aircraft holds track. If the HDG/TRK display on the MCP shows HDG, the aircraft holds heading.
Figure 1.6D - Lateral Mode Switches and Indicators
- iii) Heading Select/Track Select
Pushing the Heading Select (HDG SEL)/Track Select (TRK SEL) switch (inner) selects Heading Select or Track Select modes. In this mode, the aircraft turns to the heading or track that shows in the heading/track window. Pushing the Heading/Track (HDG/TRK) Reference switch alternately changes the heading/track reference between heading and track. Rotating the Heading/Track selector (middle) sets the heading or track in the heading/track window. If the HDG/TRK display shows HDG, the aircraft goes to and holds the heading that shows in the heading/track window. If the HDG/TRK display shows TRK, the aircraft goes to and holds the track that shows in the heading/track window. Rotating the Bank Limit selector (outer) sets the bank limit when in the Heading Select or Track Select modes. In the AUTO position, the limit varies between 15 - 25°, depending on True Airspeed. When the other detented positions are selected, the value is the maximum, regardless of airspeed.
- iv) Roll Attitude Hold
The Roll Attitude Hold mode is used to hold the roll attitude that exists at the time the flight director is first turned on, or the autopilot is first engaged. The Roll Attitude Hold mode is activated, and ATT annunciated, if the bank angle is greater than 5 degrees when either:
- A flight director is turned on with the autopilot not engaged; or
- The autopilot is initially engaged with no flight director on.
- i) Lateral Navigation
- c) Pitch Modes
The following AFDS pitch modes are available during climb, cruise and descent (Figure 1.6E [below]):
- i) Vertical Navigation
Pushing the vertical navigation (VNAV) switch arms or disarms the VNAV mode. In this mode, the AFDS uses vertical steering commands provided by the Flight Management Computer Function (FMCF). The FMCF vertical steering commands come from the active FMCF based on the navigation data and the active flight plan.
- ii) Vertical Speed/Flight Path Angle
Pushing the Vertical Speed/Flight Path Angle (V/S-FPA) switch selects the V/S or FPA mode. Rotating the V/S-FPA selector Up or Down sets the vertical speed or flight path angle in the vertical speed/flight path angle window. Pushing the V/S-FPA Reference switch alternately changes vertical speed/flight path angle window references between vertical speed and flight path angle. The vertical speed or flight path angle command is an elevator command. The pilot uses this mode to change flight levels. The pilot must set the engine thrust necessary to hold the vertical speed or flight path angle command. When the V/S/FPA display shows V/S, the aircraft goes to and holds the vertical speed that shows on the vertical speed/flight path angle window.
Figure 1.6E - Vertical Mode Switches and Indicators
- iii) Flight Level Change
Pushing the Flight Level Change (FLCH) switch selects the FLCH mode. In this mode, the AFDS will control to the speed target in the IAS/MACH window, providing climb and descent guidance and control. FLCH mode may be used with autothrottles, or with manual throttle control. When the IAS/MACH display shows IAS, the elevator command holds the speed that shows on the IAS/MACH window. When the IAS/MACH display shows MACH, the elevator command holds the MACH that shows on the IAS/MACH window. Rotating the IAS/MACH selector sets the speed in the IAS/MACH window. Pushing the IAS/MACH Reference switch alternately changes the IAS/MACH window between IAS and MACH. The Thrust Management Computing Function (TMCF) supplies the engine thrust commands.
- iv) Altitude Hold
Pushing the Altitude Hold (ALT) switch selects the Altitude hold mode. In this mode, the aircraft holds the barometric altitude present when the pilot pushes the altitude HOLD switch. Altitude Capture and Hold can also be engaged from a climb or descent as the aircraft approaches the altitude that is selected and displayed in the altitude window.
- i) Vertical Navigation
- d) Landing Modes
The following AFDS functions are available for landing:
- i) Localizer
The Localizer (LOC) mode captures and holds the aircraft to a localizer flight path. - ii) Glideslope
The Glideslope (G/S) mode captures and holds the aircraft to a vertical descent flight path. - iii) Flare
The flare (FLARE) mode controls the aircraft to a smooth touchdown at a point past the glideslope antenna. This is a computed command and is not part of the glideslope mode. - iv) Runway Alignment
In crosswind conditions, the runway alignment mode supplies roll and yaw control to decrease the aircraft crab angle for touchdown. The runway alignment mode also includes roll and yaw control for an engine failure in approach during autoland. - v) Rollout
After touchdown, the rollout (ROLLOUT) mode controls the aircraft to the runway centre line. Aircraft deviation from the localizer centre line supplies rudder and nose wheel steering signals. - vi) Go-Around
The go-around (TO/GA) mode controls roll and pitch after an aborted approach. Also, the TMCF controls thrust during go- around.
Pushing the LOC switch arms or disarms the localizer as roll mode. Pushing the Approach (APP) switch arms or disarms the localizer as roll mode and G/S as pitch mode (Figure 1.6F [below]).
Figure 1.6F - Approach Mode Switches
- i) Localizer
- e) Autothrottle (Thrust Management Computing Function)
The autothrottle (A/T) commands the thrust levers to achieve an engine thrust setting, or a selected airspeed. The A/T is armed by raising one or both A/T Arm switches, and is engaged by a pushbutton switch on the MCP (Figure 1.6G [below]).
During normal flight operations, the flight crew uses the Thrust Management Computing Function (TMCF) to perform several routine or normal operations and tasks. These operations or tasks relate to autothrottle modes. The A/T modes operate in these flight phases:
Figure 1.6G - Autothrottle Switches
- Take-off (TO)
- Climb (CLB)
- Cruise (CRZ)
- Descent (DES)
- Approach (APP)
- Go-around (GA)
Autothrottle thrust mode annunciations relate to pitch mode annunciations on the Primary Flight Display (PFD).
- f) Autothrottle Modes
- i) Take-off
In take-off (TO), the autothrottle controls thrust to the TO thrust limit. The autothrottle mode annunciation on the PFD is thrust reference (THR REF). At a threshold air speed, the autothrottle mode annunciation on the PFD changes to HOLD.
- ii) Climb
These are the three autothrottle mode selections in climb (CLB):
These are the autothrottle mode annunciations for these modes:
- THR REF when VNAV engages
- THR when FLCH engages
- SPD or THR REF when autothrottle mode engages.
The autothrottle speed mode only engages when VNAV, FLCH, and TO/GA are not active, and the aircraft is in the air.
- iii) Cruise
In cruise, the pitch mode could be VNAV PTH, VNAV ALT or MCP ALT; the corresponding A/T mode is SPD.
- iv) Descent
These are the three autothrottle modes in descent (DES):
- VNAV
- FLCH
- Autothrottle speed mode
These are the autothrottle mode annunciations in descent:
- v) Approach
SPD is normal mode in approach with glideslope active or in a manual approach (APP).
- Go-Around
A go-around (GA) mode request causes the autothrottle mode to change to THR. A second GA request causes the autothrottle mode to change to THR REF. The TO/GA switch must be pushed to request GA. - Flare Retard
Flare retard occurs when a specified altitude threshold has been achieved when in SPD mode, or during an Autoland approach with a command from the autopilot flight director system (AFDS). The autothrottle mode changes to IDLE during a flare retard.
- Go-Around
- vi) Autothrottle Disconnect
The autothrottle disconnects when there is a manual autothrottle disconnect or when there is thrust reverser application. This occurs after initial touchdown during rollout. The autothrottle will disconnect automatically for certain system faults.
3) Electrical Power
The electrical system generates and distributes AC and DC power to other aircraft systems, and is comprised of: main AC power, backup power, DC power, standby power, and flight controls power. System operation is automatic. Electrical faults are automatically detected and isolated. The AC electrical system is the main source for aircraft electrical power. Figure 1.6H (below) shows the cockpit electrical panel where electrical switching can be made. It also shows the associated lights.
As the various aircraft systems rely on electrical power, failure of the electrical buses will affect the systems operation which will in turn trigger the corresponding fault messages. These messages are collected by the CMCS which will transmit the messages, via the ACARS, to the Maintenance Control Centre (MCC).
| Note: Diagram not included | |||
| Electrical Power Panel Switches/Lights | |||
| 1. | Battery Switch | 11. | Backup Generator OFF Lights |
| 2. | Battery OFF Light | 12. | Backup Generator (BACKUP GEN) Switches |
| 3. | APU Generator (APU GEN) Switch |
13. | External Power AVAIL Lights |
| 4. | APU Generator OFF Light | 14. | External Power ON Lights |
| 5. | BUS TIE Switches | 15. | External Power (EXT PWR) Switches |
| 6. | BUS Isolation (ISLN) Lights | 16. | CABIN/UTILITY Power OFF Light |
| 7. | Generator Control (GEN CTRL) Switches |
17. | Cabin/Utility (CABIN/UTILITY) Power Switch |
| 8. | Generator OFF Lights | 18. | IFE/PASS SEATS OFF Light |
| 9. | Drive Disconnect Switches | 19. | In Flight Entertainment System/ Passenger Seats (IFE/PASS SEATS) Power Switch |
| 10. | Generator DRIVE Lights | ||
Figure 1.6H - Electrical Power Panel Switches/Lights
- a) Electrical Load Management System
The Electrical Load Management System (ELMS) provides load management and protection to ensure power is available to critical and essential equipment. If the electrical loads exceed the power available (aircraft or external), ELMS automatically shed AC loads by priority until the loads are within the capacity of the aircraft or ground power generators. The load shedding is non-essential equipment first, then utility busses. Utility busses are followed by individual equipment items powered by the main AC busses. When an additional power source becomes available or the loads decrease, ELMS restores power to shed systems (in the reverse order). The message LOAD SHED displays on the electrical synoptic when load shed conditions exist.
- b) Alternating Current Electrical System Power Sources
The entire aircraft alternating current (AC) electrical load can be supplied by any two main AC power sources. The main AC electrical power sources are:
- left and right engine integrated drive generators (IDGs)
- APU generator
- primary and secondary external power
The power sources normally operate isolated from one another. During power source transfers on the ground (such as switching from the APU generator to an engine generator) operating sources are momentarily paralleled to prevent power interruption.
- c) Integrated Drive Generators
Each engine has an Integrated Drive Generator (IDG). Each IDG has automatic control and system protection functions. When an engine starts, with the GENERATOR CONTROL switch selected ON, the IDG automatically powers the respective main bus. The previous power source is disconnected from that bus.
The IDG can be electrically disconnected from the busses by pushing the GENERATOR CONTROL switch to OFF. The IDG can also be electrically disconnected from its respective bus by selecting an available external power source prior to engine shutdown. The DRIVE light illuminates and the EICAS message ELEC GEN DRIVE L or R displays when low oil pressure is detected in an IDG. The IDG drive can be disconnected from the engine by pushing the respective DRIVE DISCONNECT switch. The IDG cannot be reconnected by the flight crew. High drive temperature causes the IDG to disconnect automatically.
- d) Auxiliary Power Unit Generator
The Auxiliary Power Unit (APU) generator is electrically identical to the IDG generators. The APU generator can power either or both main busses and may be used in flight as a replacement to an IDG source. If no other power source is available when the APU generator becomes available, the APU generator automatically connects to both main AC busses. If the primary external source is powering both main busses, the APU powers the left main bus, and the primary external source continues to power the right main bus. If the primary external source is powering the right main bus, and the secondary external source is powering the left main bus, the APU then powers the left main bus and the primary external source continues to power the right main bus. If the secondary external source is powering both main busses, the APU then powers both main busses.
The APU generator OFF light illuminates when the APU is operating and the APU generator breaker is open because of a fault or the APU GENERATOR switch is selected OFF. When the APU GENERATOR switch is ON and a fault is detected, the APU generator cannot connect to the busses.
In flight, when both transfer busses are unpowered, the APU starts automatically, regardless of APU selector position.
- e) Alternating Current Electrical Power Distribution
The AC power is distributed through the left and right main busses and the ground service bus. The right IDG normally powers the right main bus and the left IDG normally powers the left main bus. The APU normally powers both main busses when they are not powered by any other source.
Bus tie relays, controlled by BUS TIE switches, isolate or parallel the right and left main busses. When both BUS TIE switches are set to AUTO, the bus tie system operates automatically to maintain power to both main busses.
Power transfers are made without interruption when the aircraft is on the ground, except when switching between primary and secondary external power sources. The source order for powering left and right main busses in flight is the:
- f) Autoland
During autoland, the busses isolate to allow three independent sources to power the three autopilots:
- the left IDG powers the left AC transfer bus, the left main DC bus, and the captain’s flight instrument bus;
- the right IDG powers the battery bus and AC standby bus through the main battery charger; and
- the back-up system powers the right AC transfer bus, the right DC bus, and the first officer’s flight instrument bus.
- g) Backup Alternating Current Electrical System
The electrical system is highly reconfigurable to accommodate multiple failures. The electrical system is designed to automatically provide power to selected aircraft systems. The electrical system automatically powers one or both transfer busses when:
- only one main AC generator (includes APU) is available;
- power to one or both of the main AC busses is lost;
- approach (APP) mode is selected for autoland; and
- the system is automatically tested after engine starts
The system automatically transfers power without interruption.
- h) Backup Generators
Backup power is provided by one variable speed, variable frequency generator mounted on each engine. A frequency converter converts the generator frequency to a constant 400 Hz. Only one backup generator can power the converter at a time.
Each backup generator contains two permanent magnet generators (PMGs) that supply power to the flight control DC electrical system (refer to DC Electrical System). If both IDGs and the APU generator are inoperative, a backup generator powers essential aircraft equipment. To reduce electrical loading on the backup generator, the following systems are inoperative:
- i) Direct Current Electrical System
The direct current (DC) electrical system includes the main DC electrical system and the flight control DC electrical system. The main DC electrical system uses four transformer-rectifier units (TRUs) to produce DC power. The TRUs are powered by the AC transfer busses.
TRU DC electrical power is distributed to various DC busses as follows:
- (1) The left TRU powers the left main DC bus, which provides a second DC power source for:
- (2) The right TRU powers the right main DC bus, which provides a second DC power source for:
- (3) The C1 TRU powers the captain’s flight instrument bus and the battery bus. The captain’s flight instrument bus provides a second DC power source for:
- centre flight control PSA
- first officer’s flight instrument bus
The C2 TRU powers the first officer’s flight instrument bus, which provides a second DC power source for the captain’s instrument bus.
- (1) The left TRU powers the left main DC bus, which provides a second DC power source for:
- j) Batteries
The main battery is connected directly to the hot battery bus and provides standby power to other busses. The main battery charger normally powers the hot battery bus and maintains the main battery fully charged.
The APU battery is connected directly to the APU battery bus and provides dedicated power to the APU electric starter, which is used when sufficient bleed air duct pressure is unavailable for the APU air turbine starter. The APU battery charger normally powers the APU battery bus and maintains the APU battery fully charged.
- k) Flight Control Direct Current Electrical System
The flight control DC electrical system is a dedicated power source for the primary flight control system. Primary power for the flight control DC electrical system comes from permanent magnet generators (PMGs) housed within each backup generator. Variable frequency PMG AC power is used by individual power supply assemblies (PSAs) to provide DC power to the three flight control DC busses. To ensure a high level of system reliability, each PSA also has multiple DC power sources. If primary PMG AC power is not available, secondary power for the left and right PSAs is provided by the related main DC bus. Secondary power for the centre PSA is provided by the captain’s flight instrument bus. The hot battery bus provides additional backup power for the left and centre PSAs only. Each PSA also uses a dedicated battery to prevent power interruptions to the related flight control DC bus. The batteries have limited capacity and are incorporated to supply power for brief periods during PSA power source transfers.
- l) Standby Electrical System
The standby electrical system can supply AC and DC power to selected flight instruments, communications and navigation systems, and the flight control system, if there are primary electrical power system failures. The standby electrical system consists of:
- the main battery
- the standby inverter
- the RAT generator and its associated generator control unit
- the C1 and C2 TRUs
- (1) Main Battery
The main battery provides standby power to the:
- hot battery bus
- battery bus
- captain’s flight instrument bus
- left and centre flight control PSAs
- standby inverter.
Note:
The main battery can power the standby system for a minimum of 10 minutes.
- (2) Standby Inverter
The standby inverter converts DC power to AC power. The inverter powers the AC standby bus if the left transfer bus is not powered.
- (3) Ram Air Turbine Generator
The ram air turbine (RAT) generator provides standby power to the C1 and C2 TRUs. The RAT can supply electrical and hydraulic power simultaneously. If the RAT is unable to maintain RPM, the RAT generator electrical load is shed until RPM is satisfactory. Power for standby electrical loads is provided by the main battery during deployment of the RAT and when RAT generator loads are shed. The RAT is deployed automatically if both AC transfer busses lose power in flight. The RAT can be manually deployed by using the RAM AIR TURBINE switch on the overhead panel.
- (4) Cabin Systems and Utility Power
Electrical power to some cabin and utility systems are controlled from the cockpit. The IFE/PASS SEATS Power switch controls power to the IFE and passenger seats. The CABIN/UTILITY Power switch controls power to cabin and utility systems.
4) Cabin and Cargo Compartments
The aircraft, 9M-MRO was configured to 35 business class and 247 economy class seats. The business class and economy class seats were procured from BE Aerospace. An approved Lay Out of Passenger Accommodation (LOPA) determines the cabin interior configuration. Safety and emergency equipment are fitted and positioned throughout the cabin.
There is one crew rest area in the forward cabin behind the cockpit. There is a cabin crew rest area in the aft cabin lower lobe. Access is through a compartment door adjacent to Door 3R.
The cockpit door provides selective entry to the cockpit and is resistant to ballistic penetration. When closed, the door locks when electrical power is available and unlocks when electrical power is removed. A viewing lens in the door allows observation of the cabin. The door can be manually opened from the cockpit by turning the door handle.
An emergency access code is used to gain access to the cockpit in case of pilot incapacitation. Access is provided by the use of a Keypad Access System which consists of a numeric keypad outside the cockpit area and a chime module and electric strike that is not accessible from outside the cockpit. The chime module provides an audible alert to the pilots that the correct code has been entered into the keypad. There is also an indicator light in the cockpit and a Light Emitting Diode (LED) on the keypad that indicates that the correct code has been entered.
The pilots have a 3-position switch by which they can open the door lock, close the door lock, or permanently lock the door for a specified amount of time to prevent access by anyone regardless if the correct code is entered into the keypad.
The door has blowout panels that will open in the event of a rapid decompression of the passenger compartment. A pressure sensor controls an electric strike and allows the door to open inward in the event of a rapid decompression in the cockpit. These features serve to equalise the pressure between the passenger compartment and the cockpit in case of decompression either side of the door.
The aircraft is also fitted with a Flight Deck Entry Video Surveillance System (FDEVSS) which provides the pilots surveillance capability of the cockpit doorway and the forward galley areas. This allows the pilots to see the person who wants to access the cockpit before they allow entry.
There are four Type A passenger and service doors on each side of the aircraft. Each door has a window. The passenger compartment has windows along both sides of the passenger compartment. Each exit is fitted with a slide raft system for emergency use.
The overhead passenger cabin is fitted with Passenger Service Units (PSU) above each seat row. They are hinged and secured by a magnetic latch that is electrically controlled. In the event of cabin depressurisation, the PSU magnetic latch will be electrically released and allow the oxygen masks to drop for passenger use.
The aircraft cabin lighting system comprises of ceiling lights, sidewall lights, entry lights and emergency lights. The cabin management system (CMS) controls the passenger cabin lighting.
The lower section of the fuselage houses forward, aft and bulk cargo compartments. A cargo handling system is fitted for the forward and aft cargo to command power drive units (PDU) to move cargo containers laterally and longitudinally.
Cargo compartment sidewalls, ceilings and walkways are constructed of fire resistant materials. There is a smoke detection warning system and fire extinguishing system installed to contain any smoke or fire eventualities.
5) Flight Controls
The flight control system is an electronic fly by wire system. It is divided into two separate systems to control the aircraft in flight.
Primary Flight control system (PFCS) is a modern three axis, fly by wire system. It controls the roll, yaw and pitch commands using the ailerons, flaperons, spoilers, elevators, rudder and horizontal stabilizer. The high lift control system (HLCS) comprises of inboard and outboard trailing edge flaps, leading edge flaps and Kruger flaps. It supplies increased lift at lower speeds for take-off and landing.
The PFCS and HLCS use 3 dedicated ARINC 6299 Flight Control digital busses to transmit data signals to command the flight controls. Mechanical control is available to two spoilers and horizontal stabilizers.
The PFCS has three operational modes of command - Normal mode, Secondary mode and Direct mode. The PFCS command signals are computed by three redundant Primary Flight Computers (PFCs) in Normal and Secondary modes and directed through four Actuator Control Electronic (ACE) units. In Direct mode, the control surface command signals are computed by the ACEs without reliance on the PFCs.
The PFC also receives airspeed, altitude and inertial reference data from Airplane Information Management System (AIMS), Air Data Inertial Reference Unit (ADIRU) and Secondary Attitude and Air Data Reference unit (SAARU). The PFCs calculate the flight control commands based on control laws, augmentation and envelop protections. The digital command signals from the PFCs go to the ACEs that will change the digital signal to analogue format and send to the power control units (PCU) that will command the control surface movement.
The HLCS operates in three modes, primary, secondary and alternate. Command signals are transmitted from the flap lever to two Flap Slat Electronic Units (FSEU).
The FSEU process the flap command and control the sequence of flaps and slats operation. It also commands auto slat, load relief and asymmetry protection.
Two spoilers and the horizontal stabilizer receive mechanical control signals from pilots input.
_________________
9 Aeronautical Radio, Incorporated (ARINC) 629 is an aeronautical standard which specifies multi- transmitter data bus protocol where up to 128 units can share the same bus.
6) Fuel System
The fuel system has three fuel tanks, two integral wing tanks and one centre tank. The tanks are part of the wing structure and have many fuel system components located inside the tanks and on the rear spar. The fuel tanks are vented through channels in the wing to allow near ambient pressure during all phases of flight.
An integrated refuel panel (IRP) on the lower left wing and two refuel receptacles on each wing allows rapid pressure refueling of the aircraft. The refueling operation is automatic with fuel load selection on the IRP. Fuel quantity indicating system (FQIS) processor unit controls all fueling operations and measuring of fuel quantity.
Several enhanced features were incorporated in the design to include the following:
- Ultrasonic Fuel Quantity Indicating system
- Automatic centre tank scavenge system
- Ultrasonic water detection system
- Densitometers
- Jettison system
Fuel quantity is displayed on the fuel synoptic page and the upper EICAS fuel block.
- a) Engine Fuel Feed System
There are two boost pumps for each main tank and two override/ jettison pumps in the center tank to supply fuel to the engines. The fuel flows through the crossfeed manifold to the engines. Redundant crossfeed valves isolate the left and right sides of the manifold.
At the start of a flight, when all the tanks are full, the normal procedure is to turn on all the fuel pumps. The override/jettison pumps supply center tank fuel to both engines. This occurs because the override/jettison pumps have a higher output pressure than the main tank boost pumps. When the override/jettison pump output pressure decreases because of low fuel quantity in the center tank, the boost pumps automatically supply fuel to both engines from the main tanks.
- b) Auxiliary Power Unit Fuel Feed System
The Auxiliary Power Unit (APU) can receive fuel from any tank. A DC pump supplies fuel from the left main tank if no AC power is available. APU fuel is supplied from the left fuel manifold. APU fuel can be provided by any AC fuel pump supplying fuel to the left fuel manifold or by the left main tank DC fuel pump. On the ground, with the APU switch ON and no AC power available, the DC pump runs automatically. With AC power available, the left forward AC fuel pump operates automatically, regardless of fuel pump switch position, and the DC fuel pump turns off. In flight, the DC fuel pump operates automatically for quick left engine relight with the loss of both engines and all AC power. Figure 1.6I (below) shows the Engine and APU Fuel Feed System.
Figure 1.6I - Engine and APU Fuel Feed
(Copyright © Boeing. Reprinted with permission of The Boeing Company )
- c) Fuel Inlets
The fuel intake inlet for the APU (in the left main tank) is located lower than that for the engine. As the fuel level drops below the engine fuel intake level the engine will be starved of fuel, however fuel will still be available for the APU as its fuel intake is lower. This difference in level between the engine and APU fuel intakes, accounts for approximately 30 pounds of fuel in a standard flight attitude (1° pitch). The APU is estimated to consume (when electrically loaded) approximately 2 pounds of fuel in 55 seconds which will amount to a maximum APU run time of 13 minutes and 45 seconds. The pitch attitude and in-flight accelerations can affect the actual amount available for the APU.
7) Hydraulics
There are three independent hydraulic systems using electrical, pneumatics or engine driven power source. They are identified as Left, Centre and Right. Each hydraulic system can independently operate the flight controls for safe flight and landing.
Each hydraulic system uses a Hydraulic Interface Module Electronics Card (HYDIM) for automatic control and indications. The three systems operate independently at 3000 psi nominal pressure.
The left system is powered by an engine driven pump (EDP) and an electric motor pump (ACMP). The right system is also powered by an EDP and ACMP. The centre system has two ACMP and two air driven pumps (ADP) and a ram air turbine (RAT) pump.
Hydraulic pumps control and indication are on the P5 overhead panel. During normal operation the flight crew will select the switches to the auto position before flight. The pressure and quantity indication is provided on the hydraulic synoptic page and the status page.
The primary pumps are the EDPs in the left and right system and the ACMPs for the centre system. These pumps operate continuously. The demand pumps are the ACMPs for the left and right systems and the ADPs for the centre system. These pumps normally operate only during heavy system demands. The operation logic is controlled and monitored by the HYDIM cards.
The RAT deploys automatically during flight when both engines are shutdown or for loss of all three hydraulic power. The RAT hydraulic pump supplies hydraulic power to some of the center hydraulic system flight controls. When the aircraft is operating on RAT power only, the flap drive hydraulic motor is isolated from the center hydraulic system and as a result the flaps will not respond to the cockpit flap handle inputs.
8) Instrumentation
The flight instruments and displays supply information to the flight crew on six flat panel liquid crystal display units:
- Captain and First Officer Primary Flight Display (PFD)
- Captain and First Officer Navigation Display (ND)
- Engine Indication and Crew Alerting System (EICAS)
- the Multifunction Display (MFD)
Standby Flight Instruments provide information on separate indicators. Clocks display Airplane Information Management System (AIMS) generated UTC time and date, or manually set time and date.
- a) Primary Flight Display
The Primary Flight Display (PFD) presents a dynamic color display of all the parameters necessary for flight path control. The PFDs provide the following information:
- flight mode annunciation
- airspeed
- altitude
- vertical speed
- attitude
- steering information
- radio altitude
- instrument landing system display
- approach minimums
- heading/track indications, engine fail, Ground Proximity Warning System (GPWS), and Predictive Windshear (PWS) alerts.
Failure flags are displayed for aircraft system failures. Displayed information is removed or replaced by dashes if no valid information is available to the display system (because of out-of- range or malfunctioning navigation aids). Displays are removed when a source fails or when no system source information is available.
- b) Navigation Display
The navigation displays (ND) provide a mode-selectable color flight progress display. The modes are:
The MAP, VOR, and APP modes can be switched between an expanded mode with a partial compass rose and a centered mode with a full compass rose.
- c) Engine Indication and Crew Alerting System
The Engine Indication and Crew Alerting System (EICAS) consolidates engine and aircraft system indications and is the primary means of displaying system indications and alerts to the flight crew. The most important indications are displayed on EICAS which is normally displayed on the upper centre display.
- i) System Alert Level Definitions
- (1) Time Critical Warnings
Time critical warnings alert the crew of a non-normal operational condition requiring immediate crew awareness and corrective action to maintain safe flight. Master warning lights, voice alerts, and ADI indications or stick shakers announce time critical conditions.
- (2) Warnings
Warnings alert the crew to a non-normal operational or system condition requiring immediate crew awareness and corrective action.
- (3) Cautions
Cautions alert the crew to a non-normal operational or system condition requiring immediate crew awareness. Corrective action may be required.
- (4) Advisories
Advisories alert the crew to a non-normal operational or system condition requiring routine crew awareness. Corrective action may be required.
- (5) Engine Indication and Crew Alerting System Messages
Systems conditions and configuration information are provided to the crew by four types of EICAS messages:
- EICAS alert messages are the primary method to alert the crew to non-normal conditions.
- EICAS communication messages direct the crew to normal communication conditions and messages.
- EICAS memo messages are crew reminders of certain flight crew selected normal conditions.
- EICAS status messages indicate equipment faults which may affect aircraft capability.
An EICAS alert, communications, or memo message is no longer displayed when the respective condition no longer exists.
- (1) Time Critical Warnings
- i) System Alert Level Definitions
- d) Multifunction Display
The electronic checklist (ECL) system shows normal and non- normal checklists on a multifunction display (MFD). The electronic checklist system is not required for, and a paper checklist or other approved backup checklist must be available in the cockpit.
The checklist display switch on the display select panel opens the electronic checklist. The flight crew operates the checklist with the cursor control devices (CCDs).
The MFD has also communications functions which are used to control data link features. Data link messages not processed by the Flight Management Computer (FMC) are received, accepted, rejected, reviewed, composed, sent, and printed using communications functions on the MFD. ACARS and data link radio management functions are provided through communications management menus. The COMM display switch, located on the display select panel, displays the communications main menu on the selected MFD.
Communications functions are selected using the cursor control device. Message text entry is accomplished by entering data into the Control Display Unit (CDU) scratchpad and transferring it to the appropriate area. Messages can be printed on the cockpit printer. Incoming message traffic is annunciated by EICAS communications messages.
- e) Standby Flight Instruments
The standby flight instruments include:
- standby attitude indicator
- standby airspeed indicator
- standby altimeter
- standby magnetic compass
An external Power Supply Assembly supplies power to the standby attitude and airspeed indicators and the standby altimeter. The standby magnetic compass does not require any electrical power except for its lighting.
- (1) Standby Attitude Indicator
The Standby Attitude Indicator displays Secondary Attitude Air Data Reference Unit (SAARU) attitude. A bank indicator and pitch scale are provided.
- (2) Standby Airspeed Indicator
The Standby Airspeed Indicator displays airspeed calculated from two standby air data modules (one pitot and one static). It provides current airspeed in knots as a digital readout box with an airspeed pointer.
- (3) Standby Altimeter
The standby altimeter displays altitude from the standby (static) air data module. Current altitude is displayed digitally. A pointer indicates altitude in hundreds of feet. The pointer makes one complete revolution at appropriate intervals.
- (4) Standby Magnetic Compass
A standard liquid–damped magnetic standby compass is provided. A card located near the compass provides heading correction factors.
- f) Clock
A clock is located on each forward panel. Each clock displays Airplane Information Management System (AIMS) generated UTC time and date, or manually set time and date. The AIMS UTC time comes from the global positioning system (GPS). In addition to time, the clocks also provide alternating day-month and year, elapsed time, and chronograph functions.
9) Airplane Information Management System
The Airplane Information Management System (AIMS) collects and calculates large quantities of data. The AIMS manages this data for several integrated avionics systems. These systems are the:
- Primary display system (PDS)
- Central maintenance computing system (CMCS)
- Airplane condition monitoring system (ACMS)
- Flight data recorder system (FDRS)
- Data communication management system (DCMS) - including ACARS datalink
- Flight management computing system (FMCS)
- Thrust management computing system (TMCS)
The AIMS has software functions that do the calculation for each of these avionics systems. The AIMS supplies one other software function that many aircraft systems use. It is the data conversion gateway function (DCGF).
The AIMS has two cabinets, for redundancy, which do the calculations for other avionic systems. The Left cabinet is located in the forward rack of the Main Equipment Centre (MEC) while the Right cabinet is located in rear rack of the MEC. To do these calculations, each AIMS cabinet has the following:
The IOMs and CPMs are considered Line Replaceable Modules (LRM). The IOM transfers data between the software functions in the AIMS CPMs and external signal sources. The CPMs supply the software and hardware to do the calculations for several avionic systems. The software is called functions. To keep a necessary separation between the functions, each function is partitioned. The partitions permit multiple functions to use the same hardware and be in the same CPM.
The Left AIMS cabinet gets electrical power from the 28V DC Capt Flight Instrument bus and the 28V DC F/O Flight Instrument bus. The Right AIMS cabinet gets electrical power from the 28V DC Left bus and the 28V DC Right bus. Each cabinet receives the power from four 28V DC circuit breakers in the overhead circuit breaker panel. The four 28V DC bus inputs are known as power 1 through power 4. Power 1 and power 2 enter the cabinet through a connector on the left side of the cabinet and therefore they are considered as left power. Power 3 and power 4 enter the cabinet through a connector on the right side of the cabinet and are considered as right power.
Each LRM receives power from four sources, two for main power and two for monitor power. The main circuitry uses the main power. Special circuits that monitor the condition of the power supply in the LRM use the monitor power. The two main and two monitor sources of power for each LRM come from different power sources.
Each AIMS cabinet also receives power through one hot battery bus circuit breaker in the standby power management panel. The connection to the hot battery bus keeps the LRMs internal memories active. The hot battery bus also makes the AIMS cabinet less likely to have faults due to power transients.
The Navigation systems of interest include Global Positioning System (GPS), Air Data Inertial Reference System (ADIRS) and the Flight Management System (FMS).
- a) Global Positioning System
The Left and right GPS receivers are independent and use navigation satellites to supply very accurate position data to the FMC. One is powered by the 115V AC Standby bus and the other by the 115V AC Transfer bus. They pass data to aircraft systems including the ADIRS via the AIMS. GPS tuning is automatic. If the Air Data Inertial Reference Unit (ADIRU) becomes inoperative during flight, the EICAS displays the message NAV ADIRU INERTIAL and the FMC uses only GPS data to navigate.
- b) Inertial System
The ADIRS calculates aircraft altitude, airspeed, attitude, heading, and position data for the displays, flight management system, flight controls, engine controls, and other systems. The major components of ADIRS are the ADIRU, Secondary Attitude and Air Data Reference Unit (SAARU), and air data modules. The ADIRU supplies primary flight data, inertial reference, and air data. The ADIRU is fault-tolerant and fully redundant. The SAARU is a secondary source of critical flight data for displays, flight control systems, and other systems. If the ADIRU fails, the SAARU automatically supplies attitude, heading, and air data. SAARU heading must be manually set to the standby compass magnetic heading periodically. The ADIRU and SAARU receive air data from the same three sources. The ADIRU and SAARU validate the air data before it may be used for navigation. The three air data sources are the left, centre, and right pitot and static systems.
- c) Flight Management System
The FMS aids the flight crew with navigation, in-flight performance optimisation, automatic fuel monitoring, and cockpit displays. Automatic flight functions manage the aircraft lateral flight path (LNAV) and vertical flight path (VNAV). The displays include a map for aircraft orientation and command markers on the airspeed, altitude, and thrust indicators to help in flying efficient profiles. The flight crew enters the applicable route and flight data into the CDUs. The FMS then uses the navigation database, aircraft position, and supporting system data to calculate commands for manual and automatic flight path control. The FMS tunes the navigation radios and sets courses. The FMS navigation database supplies the necessary data to fly routes, SIDs, STARs, holding patterns, and procedure turns. Cruise altitudes and crossing altitude restrictions are used to calculate VNAV commands. Lateral offsets from the programmed route can be calculated and commanded.
The basis of the flight management system is the flight management computer function. Under normal conditions, one Flight Management Computer (FMC) accomplishes the flight management tasks while the other FMC monitors. The second FMC is ready to replace the first FMC if system faults occur. The FMC uses flight crew-entered flight plan data, aircraft systems data, and data from the FMC navigation database to calculate aircraft present position and pitch, roll, and thrust commands necessary to fly an optimum flight profile. The FMC sends these commands to the autothrottle, autopilot, and flight director. Map and route data are sent to the NDs. The EFIS control panels select the necessary data for the ND. The mode control panel selects the autothrottle, autopilot, and flight director operating modes.
Crew Procedure on the operations and programming of the Flight Management System safeguards and protects against incorrect execution of erroneous Information for the Navigation and Performance Data Input. Different levels of verification and cross checking between the Captain and Co-Pilot ensure that any error would be captured and corrected during the crew preparation.
In addition, system logics will also prevent the crew against selection of the wrong co-ordinates from the stored Navigation Database if a particular waypoint code happens to be used by many different places worldwide.
11) Oxygen Systems
- a) Flight Crew Oxygen System
The flight crew oxygen system provides oxygen to the flight crew for emergencies and other procedures which make its use necessary. The oxygen is supplied by two cylinders located in the left side of the main equipment centre. Each cylinder is made of composite material and holds 115 cubic feet (3,256 litres) of oxygen at 1,850 psi. The oxygen is supplied, through regulators, to four oxygen masks in the cockpit, one each for the Captain, the First Officer, the First Observer and the Second Observer. The mask has a dilution control which is normally set at ‘Normal’ position. In this position the oxygen is diluted with ambient air according to the pressure altitude in the cockpit. It can also be selected to ‘100%’, in which case 100% oxygen will be supplied. Table 1.6E (below) shows the expected duration of oxygen supply from the two cylinders with the dilution control in ‘Normal’ position.
AIRCRAFT ALTITUDE: 36,000 ft Cabin Altitude: 8,000 ft. Cabin Altitude: 36,000 ft. No. of
Crew
MembersExpected
Duration
(hour)No. of
Crew
MembersExpected
Duration
(hour)1 42 1 27 2 21 2 13 3 14 3 9 4 10.5 4 6.5 Table 1.6E - Expected Duration of Crew Oxygen
Aircraft altitude is assumed to be 36,000 ft. A cabin altitude of 8,000 ft. would indicate a normally pressurised cabin and a cabin altitude of 36,000 ft. would indicate an unpressurised cabin. At this cabin altitude of 36,000 ft, 100% oxygen will be supplied even with the dilution control in the ‘Normal’ position.
- b) Passenger Oxygen System
The passenger oxygen system is supplied by separate and individual chemical oxygen generators. The oxygen system provides oxygen to:
- passenger seats
- attendant stations
- lower crew rest compartment
- lavatory service units
The passenger oxygen masks and chemical oxygen generators are located in passenger service units (PSUs). A door with an electrically operated latch keeps the masks in a box until the oxygen deployment circuit operates. The deployment circuit operates, and the masks automatically drop from the PSUs if cabin altitude exceeds approximately 13,500 feet. The passenger masks can be manually deployed from the cockpit by pushing the overhead panel PASSENGER OXYGEN switch to the ON position. Oxygen flows from a PSU generator when any mask hanging from that PSU is pulled. Oxygen is available for approximately 22 minutes. The electrical power to the latch is supplied through a circuit breaker located in the Main Equipment Centre. It is not possible to deactivate automatic deployment of the masks from the cockpit.
- c) Portable Oxygen
Portable oxygen cylinder lets the flight attendants move in the aircraft when oxygen is in use. It is also a gaseous oxygen supply for medical emergencies. The bottle is fitted with disposable mask. 15 cylinders are located throughout the passenger cabin. Each cylinder is of 11 cubic ft (310 litres) capacity. The flow of oxygen can be controlled by an ‘Off-On’ knob which can be rotated to control the flow from 0 to 20 litres per minute. Therefore, the minimum time for the portable oxygen supply from full is 15.5 minutes.
12) Central Maintenance Computing System
The Central Maintenance Computing System (CMCS) collects and stores information from most of the aircraft systems. It can store fault histories as well as monitor and conduct tests on the various systems. The fault history contains details of warnings, cautions and maintenance messages.
At regular intervals, during flight, the CMCS transmits any recorded fault messages, via the Aircraft Communications Addressing and Reporting System (ACARS), to the Maintenance Control Centre (MCC) of Malaysia Airlines. This helps in the planning and preparation for the rectification of any potential aircraft defects at the main base or line stations. Refer also to Section 1.6.4 para. 9).
13) Engines
The aircraft is fitted with two engines (Model: RB211 TRENT 892B- 17) manufactured by Rolls-Royce. The RB211 TRENT 892B-17 engine is a high bypass turbofan (bypass ratio of 6.4:1 at a typical cruise thrust) axial flow, three-rotor with a single low pressure fan driven by a five-stage, low-pressure turbine.
The engine has an eight-stage intermediate pressure compressor driven by a single-stage turbine and a six-stage high pressure compressor driven by a single-stage turbine.
The engine take-off thrust is 92,800 lb and weighing approximately 15,700 lb (7,136 kg). The engines are certified in accordance with the US FAA Type Certificate E00050EN.
The FAA Type Certificate Data Sheet certifies that the engines meet the smoke and gaseous emission requirements of the US FAR 34. The engine is certified under FAR Part 36 Stage 3 Noise regulation.
The engine is fitted with a digital Electronic Engine Fuel Control System and it interfaces with many systems and components in the form of primary analogue or ARINC 629 buses.
The following analogue engine fuel and control system interfaces and correlates with the other systems for supply and feedback:
- Engine ignition - ignition unit power
- Engine air - actuator and valves
- Engine controls - resolver excitation and position
- Engine indicating - engine parameter data
- Engine exhaust - thrust reverser operations
- Engine oil - oil cooling and indications
- Engine starting - auto-start and manual start
- Electrical power - aircraft power from the Electrical Load Management System (ELMS)
The following ARINC 629 engine fuel and control system interfaces and correlates with other systems for supply, control and indication data:
- AIMS - indication, air data and flight management control
- Cockpit controls - switch position and indication
- Flap Slat Electronic Unit (FSEU) - Flap indication
- Proximity Switch Electronic Unit (PSEU) - Landing gear lever position
- Air Supply Cabin Pressure Controller (ASCPC) - Pneumatic system demand
The RB211 TRENT 892B-17 engine Electronic Engine Control (EEC) serves as the primary component of the engine fuel control system and uses data from the engine sensors and aircraft systems to control the engine operations. The EEC controls most of the engine components and receives feedback from them. These digital data go to the Engine Data Interface Unit (EDIU) and send the signal to the AIMS. The AIMS transmits and receives a large amount of data to and from the EEC. These include:
- Engine bleed status - EEC thrust limit calculations
- Air data - EEC thrust limit calculations
- Engine data – system requirements
- Autothrottle Engine Pressure Ratio (EPR) trim - thrust balancing
- Condition monitoring - performance tracking
- Maintenance data - trouble shooting
- Primary display system data - indication.
14) Auxiliary Power Unit
The aircraft is fitted with an Auxiliary Power Unit (APU) - Model: GTCP 331-500 - manufactured by Allied Signal. The Allied Signal GTCP 331-500 gas turbine APU is a two-stage centrifugal flow compressor, a reverse flow annular combustion chamber and a three-stage axial flow turbine. It supplies the auxiliary power system for the aircraft pneumatic and electrical power. This permits independent operations from the ground external power sources or the main engines.
The APU generator supplies 120 KVA electrical power at any altitude. The APU can start at all altitudes up to the service ceiling of the aircraft (43,100 ft/13,100 m). Electrical power is available up to the service ceiling and pneumatic power is available up to 22,000 ft (6,700 m).
The ELMS contains the APU autostart logic and sends signal to the APU Controller (APUC).
The APU Controller serves to control the APU functions for:
- Starting and ignition
- Fuel metering
- Surge control
- Inlet guide vane (IGV) control
- Data storage
- Protective shutdown
- BITE/Fault reporting
- APU indication
The APU is designed to automatically start when certain logic conditions are met when the aircraft is in the air or electrical power removed from left and right transfer buses from respective No. 1 and No. 2 engine generators.
15) Communications
For Communications Systems description, refer to Section 1.9.