Factual Information 2015:Appendix 1.6E

APPENDIX 1.6E - AIRCRAFT SYSTEMS DESCRIPTION

1.6E.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.

Two electronic controllers provide 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.

1.6E.2 Autopilot Flight Director System (AFDS)

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 1.6EA). 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 rudder pedals or control column. The autopilot can also be disengaged by pushing on the A/P Disengage Bar on the MCP. The AFDS consists of three autopilot flight director computers and the MCP.

Figure 1.6EA Autopilot Mode Control Panel

1.6E.2.1 Roll Modes

The following AFDS roll modes are available during climb, cruise and descent (Figure 1.6EB below):

a) Lateral Navigation (LNAV)

Pushing the LNAV switch arms, selects 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. The Control Display Units (CDUs) can send LNAV steering commands when there is no active FMCF. The AFDS use this priority to select a CDU command:

Left, if valid
Centre, if left not valid
Right, if left and centre not valid.

The CDU is the primary control and display interface for the FMCF. The CDU is used to enter flight plan data and performance data. The CDU can also be used to manually tune the navigation radios and access maintenance pages.

b) Heading Hold (HDG HOLD/Track hold (TRK HOLD)

Pushing the Heading/Track 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 1.6EB Lateral Mode Switches and Indicators

b) Heading select (HDG SEL)/Track select (TRK SEL)

Pushing the Heading/Track Select 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 degrees, depending on True Airspeed. When the other detented positions are selected, the value is the maximum, regardless of airspeed.

1.6E.2.2 Pitch Modes

The following AFDS pitch modes are available during climb, cruise and descent (Figure 1.6EC below).

a) Vertical navigation (VNAV)

Pushing the VNAV switch arms, selects or disarms the VNAV mode. The VNAV mode is a mix of throttle and elevator commands that control the vertical flight path. The FMCF vertical steering commands come from the active FMCF based on the navigation data and the active flight plan.

b) Vertical speed (V/S)/Flight Path Angle (FPA)

Pushing the 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 1.6EC Vertical Mode Switches and Indicators

c) Flight Level Change (FLCH)

Pushing the FLCH switch selects the Flight level change speed mode. The FLCH command is a mix of thrust and elevator commands to change flight levels. 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.

d) Altitude Hold (ALT)

Pushing the Altitude Hold switch selects the Altitude hold mode. In this mode, the aircraft holds the barometric altitude present when the pilot pushes the altitude HOLD switch.

1.6E.2.3 Landing Modes

The following AFDS functions are available for landing:

a) Localizer (LOC)

The LOC mode captures and holds the aircraft to a localizer flight path.

b) Glideslope (G/S)

The G/S mode captures and holds the aircraft to a vertical descent flight path.

c) Flare (FLARE)

The 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.

d) 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.

e) Rollout (ROLLOUT)

After touchdown, the rollout mode controls the aircraft to the runway centre line. Aircraft deviation from the localizer centre line supplies rudder and nose wheel steering signals.

f) Go-Around (TO/GA)

The go-around mode controls roll and pitch after an aborted approach. Also, the TMCF controls thrust during go-around.

Pushing the Localizer (LOC) switch arms, disarms or captures the localizer as roll mode. Pushing the Approach (APP) switch arms, disarms or captures the localizer as roll mode and glidepath (G/S) as pitch mode (Figure 1.6DD below).

Figure 1.6ED Approach Mode Switches

1.6E.3 Autothrottle (Thrust Management Computing Function – TMCF)

The autothrottle (A/T) commands the thrust levers to achieve an engine thrust setting, or a selected airspeed. The A/T is armed by the operation of two toggle switches and engaged by the operation of a pushbutton switch on the MCP (Figure 1.6DE below).

Figure 1.6EE Autothrottle Switches

During normal flight operations, the flight crew uses the TMCF to perform several routine or normal operations and tasks. These operations or tasks relate to autothrottle modes. The autothrottle (A/T) modes operate in these flight phases:

Take-off (TO)
Climb (CLB)
Cruise (CRZ)
Descent (DES)
Approach (APP)
Go-around (GA)

Autothrottle functions that relate to flight phases are flare retard during autoland and autothrottle disconnect. Autothrottle thrust mode annunciations relate to pitch mode annunciations on the Primary Flight Display (PFD).

1.6E.3.1 Autothrottle Modes

a) Take-off (TO)

In TO, the autothrottle controls thrust to the take-off 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.

b) Climb (CLB)

i. These are the three autothrottle mode selections in climb:

Vertical navigation (VNAV)
Flight level change (FLCH)
Autothrottle (MCP) speed mode or thrust mode.

ii. 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/GO are not active and the aircraft is in the air.

c) Cruise (CRZ)

i. These are the two autothrottle modes in cruise:

VNAV
Autothrottle speed mode.

ii. These are the autothrottle mode annunciations in cruise:

SPD when VNAV engages
SPD, VNAV is not active

d) Descent (DES)

i. These are the three autothrottle modes in descent:

VNAV
FLCH
Autothrottle speed mode.

ii. These are the autothrottle mode annunciations in descent:

IDLE, THR, or HOLD shows for VNAV
THR, or HOLD shows for FLCH
SPD.

e. Approach (APP)

SPD is normal mode in approach with glideslope active or in a manual approach.

i. Go-Around (GA)

A 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 lever must be pushed to request GA.

ii. Flare Retard

Flare retard occurs when a specified altitude threshold has been achieved during approach with a command from the autopilot flight director system (AFDS). The autothrottle mode changes to IDLE during a flare retard.

1.6E.3.2 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.

1.6E.4 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 airplane electrical power. Figure 1.6EF shows the cockpit Electrical panel where electrical switching can be made. It also shows the associated lights.

1.6E.4.1 Electrical Load Management System (ELMS)

The 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 sheds 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.

1.6E.4.2 AC Electrical System Power Sources

The entire aircraft 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.

1.6E.4.3 Integrated Drive Generators (IDGs)

Each engine has an 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.

1.6E.4.4 APU Generator

The 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.

1.6E.4.5 AC Electrical Power Distribution

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.

When external power is connected:

primary external power normally powers the right main bus
secondary external power normally powers the left main bus

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 airplane 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: ·

respective IDG
APU generator
opposite IDG

1.6E.4.6 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
the backup system powers the right AC transfer bus, the right DC bus, and the first officer’s flight instrument bus.

1.6E.4.7 Backup AC Electrical System

The backup electrical system is designed to automatically provide power to selected aircraft systems. The backup 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
the system is automatically tested after engine starts

The system transfers power without interruption.

1.6E.4.8 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 airplane equipment. To reduce electrical loading on the backup generator, the following systems are inoperative:

TCAS
SATCOM
Right HF radio

1.6E.4.9 DC Electrical System

The 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.

The left TRU powers the left main DC bus, which provides a second DC power source for:

left flight control power supply assembly (PSA)
right main DC bus.

The right TRU powers the right main DC bus, which provides a second DC power source for:

right flight control PSA
left main DC bus.

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.

1.6E.4.10 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.

1.6E.4.11 Flight Control DC 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.

1.6E.4.12 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

a) 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.

b) 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.

c) Ram Air Turbine (RAT) Generator

The 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 pushing the RAM AIR TURBINE switch on the overhead panel.

1.6E.4.13 Cabin Systems and Utility Power

Electrical power to some cabin and utility systems are controlled from the flight deck. 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.

Figure 1.6EF Electrical Power Panel Switches/Lights

Electrical Power Panel Switches/Lights (Figure 1.6EF)
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

1.6E.5 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. The cockpit door is reinforced and electrically locked. There is a cabin crew rest area in the aft cabin lower lobe. Access is through a compartment door adjacent to Door 3R.

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.

1.6E.6 Fire Protection

Fire protection and overheat detection systems are provided for the Engines, APU, wheel well, cargo compartment and the pneumatic ductings.

The Rolls Royce engine has a dual loop fire detection system that monitors the external areas of the engine. The detectors monitor the engine for fire and overheat conditions. Detector signals are monitored by a detection card and sends signals to the cockpit indication system. The engines have a two shot fire extinguishing system. The pilot can select to discharge the fire bottle. Halon gas will be discharged through nozzles positioned around the engine.

The APU has a dual loop fire detection system around the APU compartment. If a fire condition is detected, a signal is sent to the detection card. This will automatically activate a single shot fire extinguishing system that will discharge Halon gas into the APU compartment.

The wheel well fire detection system monitors the wheel well for brake and tyre fires.

The forward cargo compartment and aft cargo compartment have smoke detectors to monitor the air. The smoke detectors analyse the air for smoke particles. The forward cargo smoke detector also process signals from the main equipment centre cooling smoke detector.

The cargo compartment has an extinguishing system that comprises of five extinguishing bottles fitted in the forward cargo compartment. There are series of pipes that connect the bottles to the forward and aft cargo compartments. The detection and extinguishing system is monitored through a smoke detection card file and annunciated in the cockpit.

The high pressure pneumatic ducting from the engines and the APU have overheat protection. Should a duct leak be detected the associated pneumatic system will be isolated.

1.6E.7 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 uses 3 dedicated ARINC 629¹ 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 directed through four Actuator Control Electronic (ACE) units that change analogue signals to digital format to send to three Primary Flight Computers (PFC).

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 (FSCU).

The FSCU 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.

_________________________

¹ 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.

1.6E.8 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 fuels 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 fuelling 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.

1.6E.9 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 loss of all three hydraulic power.

1.6E.10 Ice and Rain Protection

Ice and rain protection is carried out on the wings, engine intake, air data probes, cockpit windows, water and waste lines.

The aircraft is fitted with two ice detector probes on each side of the forward fuselage. When ice collects on either detector, a signal is sent to the engine Airfoil Cowl Ice Protection System (ACIPS) card. The engine ACIPS card shares information with the wing ACIPS card. The cards then operate the wing and engine anti ice systems automatically when engine and wing anti ice switches are in auto and the aircraft is in the air. Engine bleed air is used for engine anti ice and wing anti ice functions.

Air Data Probes comprise of three Pitot probes, two Angle of Attack (AOA) probes and one Total Air Temperature (TAT) probe fitted on the forward fuselage. These probes have a built-in heater to prevent icing. The Electric Load Management System (ELMS) controls the level of heating in flight and ground mode.

The cockpit windshield and side windows are heated to prevent ice and fog. There is an in-built electrically resistive material in the window lamination that is electrically heated. Two window heat control units (WHCU) control the level of heating when the window heat switches are turned ON.

The cockpit has a windshield wiper system to remove rain from the LH and RH windshield. Two switches operate the LH and RH wiper operation.

The water supply and waste water line are fitted with inline heaters at selected areas to prevent freezing. The waste water drain mast is also heated to prevent icing.

1.6E.11 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)
the 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 (PFD)

The PFDs present 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 (ND)

The NDs provide a mode–selectable color flight progress display. The modes are:

MAP
VOR
APP (approach)
PLN (plan)

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 (EICAS)

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.

System Alert Level Definitions

i) 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.

ii) Warnings

Warnings alert the crew to a non-normal operational or system condition requiring immediate crew awareness and corrective action.

iii) Cautions

Cautions alert the crew to a non-normal operational or system condition requiring immediate crew awareness. Corrective action may be required.

iv) Advisories

Advisories alert the crew to a non-normal operational or system condition requiring routine crew awareness. Corrective action may be required.

v) EICAS 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 dispatch capability

An EICAS alert, communications, or memo message is no longer displayed when the respective condition no longer exists.

d) Multifunction Display (MFD)

The electronic checklist (ECL) system shows normal and non-normal checklists on a multifunction display (MFD). The electronic checklist system is not required for dispatch, and a paper checklist or other approved backup checklist must be available on the flight deck.

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).

e) Standby Flight Instruments

The standby flight instruments include:

standby attitude indicator
standby airspeed indicator
standby altimeter
standby magnetic compass

i) Standby Attitude Indicator

The Standby Attitude Indicator displays Secondary Attitude Air Data Reference Unit (SAARU) attitude. A bank indicator and pitch scale are provided.

ii) 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 and with an airspeed pointer.

iii) 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.

iv) 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.

1.6E.12 Airplane Information Management System (AIMS)

The 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)
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 which do the calculations for other avionic systems. These cabinets (Left AIMS Cabinet and Right AIMS Cabinet) are located in the Main Equipment Centre (MEC). To do these calculations, each AIMS cabinet has the following:

A cabinet chassis
Four Input/output modules (IOM)
Four Core processor modules (CPM)

The IOMs and the CPMs are in the cabinet chassis. The chassis also has a backplane data bus and a backplane power bus to distribute data and power to the IOMs and CPMs. The input/output module (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.

1.6E.13 Landing Gear

The landing gear is of tricycle design with two main landing gears and one nose landing gear. The nose landing gear is of conventional two wheel design while the main landing gear truck has six wheels.

The nose landing gear strut includes an air oil shock absorber. The nose landing gear uses centre hydraulics pressure to extend and retract. Sequence valves control the door and landing gear movement.

The main landing gear has three axles’ trucks and fitted with six wheel and brake assemblies. The main gear uses centre system hydraulic pressure to extend and retract. Sequence valves control the door and gear movements. Drag braces and downlock actuators lock the gear in the extend position. Uplock hooks lock the gear in the retract position.

Alternate extension system permits landing gear extension if the centre hydraulic system has no pressure. An alternate extend power pack will unlock the landing gear doors and the uplock. The door opens and the gear extend under its own weight. The gear doors will remain open after an alternate extension.

Nose gear steering is available when aircraft is on the ground. Two steering tillers in the cockpit allows nose wheel steering up to a specified maximum in each direction. The rudder pedal input allows smaller nose wheel steering in each direction.

Main gear steering operates when nose wheel steering commands are more than a specified limit. The main gear aft axle is able to steer left or right when commanded by main gear steering control unit (MGSCU).

Two sets of brake pedal controls the brake. Normal braking uses right system hydraulic pressure and alternate brakes use centre system hydraulic pressure. The braking system pressure is applied through the brake system control unit (BSCU) to control brake pressure to prevent tyre skid. Each wheel has a wheel speed transducer which supplies signal to the BSCU.

1.6E.14 Navigation Systems

Navigation systems include Global Positioning System (GPS), Air Data Inertial Reference System (ADIRS), Very High Frequency Omni Range (VOR), Distance Measuring Equipment (DME), Instrument Landing System (ILS), Automatic Direction Finder (ADF), Weather Radar, and the Flight Management System (FMS).

1.6E.14.1 Global Positioning System (GPS)

Left and right GPS receivers are independent and supply very accurate position data to the FMC. 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.

1.6E.14.2 Inertial System

The Air Data Inertial Reference System (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 Air Data Inertial Reference Unit (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.

1.6E.14.3 Radio Navigation Systems

a) Automatic Direction Finding (ADF): Two ADF systems are installed. Either ADF can be manually tuned from the left or right CDU on the NAV RADIO page.
b) Very High Frequency Omni Range (VOR): Two VOR receivers are usually tuned by the FMC but, can be tuned manually by the crew. The tuned VORs display on the ND and may be used for position updates.
c) Distance Measuring Equipment (DME): Two DME systems are installed. The DMEs are usually tuned by the FMC, but may be tuned manually.
d) Instrument Landing System (ILS): Three ILS receivers are installed. They are usually tuned by the FMC, but can be tuned manually on the NAV RADIO page.

1.6E.14.4 Weather Radar

The weather radar system consists of receiver–transmitter unit, antenna, and control panel. Radar returns display on the Navigation Display (ND). The EFIS control panel weather radar (WXR) map switch controls power to the transmitter/receiver and controls the weather radar display on the ND.

1.6E.14.5 Flight Management System (FMS)

The FMS aids the flight crew with navigation, in–flight performance optimisation, automatic fuel monitoring, and flight deck 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 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.

1.6E.15 Oxygen Systems

1.6E.15.1 Crew Oxygen System

The 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 (3150 litres) of oxygen at 1850 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 flight deck. It can also be selected to ‘100%’, in which case 100% oxygen will be supplied. Table 1.6A below shows the expected duration of oxygen supply from the two cylinders with the dilution control in ‘Normal’ position. 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.

Aircraft Altitude: 36,000 ft. Cabin Altitude: 8,000 ft.		Aircraft Altitude: 36,000 ft. Cabin Altitude: 36,000 ft.
No. of Crew Members	Expected Duration (hour)	No. of Crew Members	Expected Duration (hour)
1	42	1	27
2	21	2	13
3	14	3	9
4	10.5	4	6.5

Table 1.6EA Expected Duration of Crew Oxygen

1.6E.15.2 Passenger Oxygen System

The passenger oxygen system is supplied by 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). Oxygen flows from a PSU generator when any mask hanging from that PSU is pulled. Oxygen is available for approximately 22 minutes. The masks automatically drop from the PSUs if cabin altitude exceeds approximately 13,500 feet. The passenger masks can be manually deployed from the flight deck by pushing the overhead panel PASSENGER OXYGEN switch to the ON position.

1.6E.15.3 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 feet (301 litres) capacity. The flow of oxygen can be controlled by an ‘Off-On’knob from 0 to 20 litres per minute.

1.6E.16 Structures

The B777-200 is a transport category aircraft certified to Federal Aviation Regulations (FAR) Part 25. The structure construction is a conventional skin, frame, stringers and longeron to handle the flight load on the aircraft during its operation. The structure components are fuselage, wing and empennage which also consist of the horizontal and vertical stabilizers. Control surfaces to control the aircraft manoeuvrability are attached to the wing, and empennage. Engine assemblies are attached to the wings and the landing gear system to the part of the fuselage. The fuselage has numerous cut-outs for doors, windows and inspection doors.

1.6E.16.1 Fuselage

The fuselage is a semi-monocoque thin wall structure, which consists of skin panel, frames and stringers. The fuselage is designed with frames and stringers. Keel beam, located at the underside of the fuselage, provides structural reinforcement and provides protection to the underside of the centre fuel tank in the event of an emergency landing. Boeing has introduced modern materials on the B777-200 structure, which are improved Aluminum Alloy 7055 and carbon fiber with toughened resin (composite).

1.6E.16.2 Wings

The B777-200 wings comprise of two outer wings and a centre wing box. The wings are conventional design with front and rear spars together with the upper and lower wing skins reinforced with stringers. The two outer wings are attached to the wing box using wing joint fittings.

1.6E.16.3 Engine Nacelle Attachments

The B777-200 engine nacelle and pylon assemblies are attached to the wing by four fuse pins. The fuse pins are designed to fail in the event of abnormal loads being applied to the nacelle, such as during an emergency landing, in order to preserve the wing structure and allow the engine assembly to separate cleanly.

1.6E.16.4 Landing Gear System

The main landing gear drag brace is joined to the rear face of the rear spar web, the terminal fitting and the back-up fitting. Attached to the innermost rib is a further back-up fitting which again attaches through the rear spar web to the main landing gear drag brace fitting. The wing box also forms part of the centre fuel tank and the wing fuel tanks.

1.6E.17 Central Maintenance Computing System (CMCS)

The 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.

1.6E.18 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 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
Flight deck 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 Ration (EPR) trim – thrust balancing
Condition monitoring – performance tracking
Maintenance data – trouble shooting
Primary display system data – indication.

The RB211 TRENT 892B-17 engine has the capability to generate snapshot reports of engine data for the purpose of Engine Health Monitoring.

1.6E.19 Auxiliary Power Unit (APU)

The aircraft is fitted with an 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. Pneumatic pressure is available up to an altitude of 22,000 feet (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.

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Disclaimer

Source: Malaysian ICAO Annex 13 Safety Investigation Team for MH370, 8 March 2015, Factual Information MH370/01/15

The Factual Information was updated in 2018 by the Safety Investigation Report MH370/01/2018 which added new content but did not include all of the previous data.

Note: Each extract can be viewed separately using the links below:-