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A-320 STUDY NOTES

A-320 Systems Study Notes .

TRAINING PURPOSES ONLY -1-

A-320 STUDY NOTES

IMPORTANT NOTICE This manual is for training purpose only. It has been developed to assist new pilots on the A-320 fleet and to act as a refresher for more experienced type rated pilots. The information comes from several different sources, and therefore may differ from standard Airbus documentation. In all cases, the company’s SOPs, and standard Airbus documentation will take precedence. The manual is not designed to be totally comprehensive, nor does it cover the level of detail required for type rated pilots, it is exactly what it is described to be, Study Notes.

NOTICE OF LIABILITY The information in these study notes is distributed on an “as is” basis, without warranty. The author shall have no liability to any person or entity with respect to any liability, loss or damage caused, or alleged to be caused, directly or indirectly by information contained in these study notes or it’s application in practice.

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A-320 STUDY NOTES

TABLE OF CONTENTS ELECTRICS

4 - 12

HYRAULICS

13 - 18

LANDING GEAR

19 - 21

BRAKES

22-25

PNEUMATICS

26 - 30

AIR CONDITIONING

31 -34

VENTILATION

35 - 37

PRESSURIZATION

38 - 41

FIRE PROTECTION

42 - 45

FUEL

46 - 49

FLIGHT CONTROLS

50 - 69

NAVIGATION

70 -118

ICE AND RAIN

119 -122

ENGINES

123 - 126

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A-320 STUDY NOTES

ELECTRICS The electrical system is divided into two separate and isolated channels; number 1 and number 2. Electrical power is provided by two engine driven generators, an APU generator, an emergency generator and two nickel cadmium batteries. Each engine and APU generator, incorporating its own constant speed drive and drive oil system as a single unit, is called an integrated drive generator (IDG). Each IDG ensures a constant output of 90 KVA, 115/200 V 400 HZ, AC power. Should it be necessary to disconnect an IDG in flight, the disconnect push-button (pb) should be held for a maximum of 3 secs, and only when the engine is running otherwise the mechanism will be damaged. In normal operations, number 1 IDG provides power to AC BUS 1, which then powers the AC ESS BUS, which powers the AC ESS SHED BUS. TR 1 is also fed by AC BUS 1 which in turn powers DC BUS 1 and the DC BATTERY BUS. The DC BAT BUS powers the DC ESS BUS via the ESS DC TIE contactor, which then supplies the DC SHED BUS. The two aircraft batteries, numbers 1 and 2 are connected directly to their respective HOT BAT busses. Each battery has its own battery charge limiter (BCL) that monitors the battery charge level and if necessary, connects to the DC BATTERY BUS via its respective BATTERY contactor. The number 2 IDG normally provides power to AC BUS 2, which powers DC BUS 2 through TR 2. The entire electrical network can be powered by only one generator; either of the engine generators or the APU generator via the AC BUS TIE contactors. On the ground, external power can also be used to power the complete network. The priority of the generators is: • • •

Engine driven-generators External Power APU generators

The minimum battery voltage of 25.5v must be checked prior to departure. The batteries have built in protection to ensure they do not go below 22.5v, even if the switches are left on with the aircraft unpowered on the ground overnight. To check the battery voltage and the capability of the battery chargers on the ground, the batteries must first be switched off. If the voltage is below 25.5v, then connect TRAINING PURPOSES ONLY -4-

A-320 STUDY NOTES external power; turn the batteries on (AUTO) again; check on the ECAM ELEC page to ensure that the initial charge rate after 10 secs is less than 60 amps and is decreasing. After 20 minutes the batteries should be sufficiently charged.

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A-320 STUDY NOTES

IDG FAILURE •

If an IDG fails to power its own channel, the operative IDG will power both channels automatically through the AC TIE BUS and the Primary Galley Power will be shed. If the APU generator is available, it will automatically replace the failed generator via the AC TIE BUS and subsequently, the galley power can be restored as the electrical system will be back to normal output.

IDG 1 failure with IDG 2 feeding the entire system Note: Galley shed indication

IDG 2 failure, with APU feeding AC BUS 2 etc. Note: Galley power has been restored

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A-320 STUDY NOTES AC BUS 1 FAILURE •

Failure of AC BUS 1 leads to the momentary loss of the AC ESS BUS; however it can be restored by selecting the AC ESS FEED pb which will enable AC BUS 2 to feed the AC ESS. The ESSENTIAL TR will then be powered by the AC ESS BUS which will in turn feed the DC- ESS BUS. In this configuration the DC BATT BUS will be fed automatically by DC BUS 2 via the DC TIE contactor and the DC BUS 1 will be fed by the DC BATT BUS, via the other DC TIE contactor (the automatic changeover takes approximately 5 seconds).

AC BUS 1 failure with the ESS TR feeding the DC ESS.

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A-320 STUDY NOTES TR1 OR 2 FAILURE •

If TR 1 or TR 2 fail, their respective busses will automatically be fed through the DC BAT BUS by the opposite DC channel via the DC BUS TIE contactors. In the example below; TR 1 has failed and DC BUS 2 is feeding the DC BAT BUS which in turn is feeding DC BUS 1. The DC ESS BUS is being supplied by the ESS TR powered by the AC ESS BUS.

TR 1 failure with the ESS TR feeding the DC ESS.

The TR contactors will open automatically in case of the following: • TR Overheat • Current below the minimum required

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A-320 STUDY NOTES BOTH TR1 AND TR2 FAILURES •

If both TR 1 and TR 2 fail; DC BUS 1, DC BUS 2 and the DC BAT BUS will be all lost. The DC ESS BUS will be fed by the ESS TR powered by the AC ESS BUS.

TR 1 and TR 2 failure with the loss of DC 1, DC 2, and DC BAT busses. Only DC buses available are the HOT BAT and DC ESS busses.

BOTH AC BUS FAILURES •

If you are unlucky enough to lose both main AC busses and, your Ram Air Turbine (RAT), the good news is that the Static Inverter will connect to HOT BAT BUS 1 and provide power to the AC ESS BUS. HOT BAT BUS 2 will supply the DC ESS BUS. Both AC SHED and DC SHED Busses are also lost in this configuration. The bad news here is that the batteries will only last for approx. 22 minutes however, if you go through the procedure in QRH 2.03 (Flight on Bat Only), then you can increase your chances by increasing the battery life to at least 30 minutes. Reference to QRH 1.01 will let you know what equipment you have left in this configuration. Below 100 knots the DC BATT BUS is connected and below 50 knots the Static Inverter will disconnect and you will lose your AC ESS BUS, leading to a loss of the remaining CRT’s.

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A-320 STUDY NOTES RAM AIR TURBINE (RAT) •

If both AC BUS 1 and AC BUS 2 are lost and the speed is above 100 knots, the RAT will extend automatically. The RAT powers the blue hydraulic system which in turn will drive a constant speed hydraulic motor/generator. RAT extension and coupling takes approximately 8 seconds, during which time the aircraft is powered by the batteries only. Like the previous example, AC ESS is fed by the static inverter which gets its power from HOT BAT BUS 1 and the DC ESS is powered by HOT BAT BUS 2. (Both AC and DC SHED busses are un-powered). The red light on the EMER ELEC PWR panel (the only red fault light on the overhead panel) remains illuminated during the RAT extension. When the RAT is coupled it powers the AC ESS BUS which then powers the ESS TR and the DC ESS BUS. When the landing gear is extended the system reverts back to being powered by the batteries only. Once again, in this configuration the time available on batteries is approximately 22 minutes. The DC BAT BUS connects below 100 knots and the AC ESS is lost below 50 knots. The recommended minimum speed for flight with the RAT powering the aircraft is 140 knots, which prevents the RAT stalling. If the AC busses have been lost due to a short circuit, it is unlikely that the APU generator will couple and therefore, it might be wise to conserve battery power. Each APU start attempt will drain approximately 3.5 minutes of battery power. If only the batteries are powering the aircraft electrical system, the APU will only start below 100 knots when the DC BAT BUS connects to the system.

See next page for ECAM diagnostic page:

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A-320 STUDY NOTES

EMER ELEC Configuration

CIRCUIT BREAKERS •

The circuit breakers (CBs) are either green or black. The green ones are monitored and will display C/B TRIPPED ON OVHD PNL or REAR PNL on the ECAM if one has tripped for more than 1 minute. The Yellow collared (black) CBs on the overhead panel are for use if you are unlucky enough to be flying on batteries only (see FLT ON BAT ONLY checklist, QRH 2.03). There are also red collared CBs on the rear panel that must never be pulled in flight; these CBs are for the wing tip brakes. See QRH 2.34 the policy of reengaging tripped CBs, either on the ground or in flight. Essentially, resetting CBs is not recommended except in special circumstances which are outlined in then policy.

APU •

When shutting down the APU on the ground, it is important to wait for 2 minutes after the green AVAIL light is extinguished to ensure the APU flap is fully closed and fire protection is still available, before turning off the batteries.

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A-320 STUDY NOTES SMOKE •

•

If smoke is detected in the Avionics compartment, amber fault lights will illuminate on the EMER ELEC PWR panel and also the VENTILATION panel associated with an ECAM warning. The procedure calls for depressing the GEN 1 LINE pb.; this action opens the generator line contactor and sheds AC BUS 1 momentarily; AC BUS 2 will then automatically power AC BUS 1 through the BUS TIE contactors. The purpose of using the GEN 1 LINE pb is to isolate one inner tank fuel pump in each main tank and power them by the number one generator, thereby assuring positive fuel pressure while the rest of the procedure is carried out. The procedure then calls for the RAT to be extended and when the emergency generator is available, APU GEN and GEN 2 are selected off; this action will shed approximately 75% of the electrical network, including the remaining fuel pumps. The aircraft will now in the emergency electrical configuration.

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A-320 STUDY NOTES

HYDRAULICS Hydraulic power is supplied by three independent hydraulic circuits called, the Green, Blue and Yellow systems. The normal hydraulic system pressure is 3000 psi. Each system is supplied by its own pneumatically pressurized reservoir and it is not possible to transfer fluid from one system to another.

GREEN SYSTEM The green system is pressurized by the number 1 engine-driven hydraulic pump. There is a fire shutoff valve located upstream of the pump which is operated by the ENG 1 FIRE pb. There are two pressure sensors located downstream of the pump, one sends the system pressure to the ECAM and the other to the Power Transfer Unit (PTU). Downstream of the pressure sensors there is a non-return valve, a system accumulator, a leak measurement valve, the various user components, a priority valve, two load alleviation function (LAF) accumulators and the PTU hydraulic line.

BLUE SYSTEM The blue system is pressurized by an electric motor-driven hydraulic pump. There is also a ram air turbine (RAT) that can pressurize the blue system to approximately 2500 psi in an emergency. There are two pressure sensors located downstream of the pumps which send system pressure to the ECAM. Downstream of the pressure sensors there is a system accumulator, a leak measurement valve, the various user components and a priority valve.

YELLOW SYSTEM The yellow system is pressurized by the number 2 engine-driven hydraulic pump. There is a fire shutoff valve located upstream of the pump which is operated by the ENG 2 FIRE pb. This system also includes an electric motor-driven pump and a hand pump for operation of the cargo doors when electric power is not available. There are two pressure sensors located downstream of the pumps, one sends the system pressure to the ECAM and the other to the (PTU). Downstream of the pressure sensors there is a non-return valve, a system accumulator, a leak measurement valve, the various user components, a priority valve, two (LAF) accumulators and the PTU hydraulic line.

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A-320 STUDY NOTES

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A-320 STUDY NOTES POWER TRANSFER UNIT (PTU) The PTU is a bi-directional motor pump located between the green and yellow hydraulic systems. It enables either system to power the other in case of pump failure. The PTU is automatically activated if there is a pressure differential of 500 psi or more between systems. On the ground with the engines not running, the PTU can be used to power the green system by activating the yellow electric pump. The PTU is automatically inhibited: • • • •

During the first engine start When the cargo doors are operating When the parking brake is on and only one engine is running When the PTU switch is off

•

The PTU is automatically tested during second engine start. If the second engine is started within 40 seconds of the cargo doors closing, a HYD PTU FAULT will be triggered. To reset the warning, switch the yellow electric pump ON and then OFF. (3.02.29 P16) The PTU does not transfer fluid.

•

PTU in operation GREEN system being powered by YELLOW system

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A-320 STUDY NOTES ELECTRIC PUMPS The blue electric pump will operate and pressurize the blue system on the ground if the switch is in AUTO, at least one engine is running and AC power is available. The pump can also be operated on the ground by pressing the BLUE PUMP OVRD switch on the overhead maintenance panel. The pump runs continuously in flight, unless the ELEC PUMP switch is OFF. The yellow electric pump will start automatically, regardless of switch position, when the cargo doors are operated. If there is a power interruption during cargo door operations, the pump is de-energized until its switch is cycled. During cargo door operations, all other yellow system users are inhibited except, number 2 reverser and the brake accumulator. A yellow system hand pump can be used to operate the cargo doors if there is no electrical power available.

YELLOW ELEC pump powering YELLOW system, on the ground

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A-320 STUDY NOTES RAM AIR TURBINE (RAT) The RAT can be deployed in an emergency to pressurize the blue hydraulic system, by pressing the RAT MAN ON switch on the Hydraulic overhead panel. In the event of a loss of both AC BUS 1 and AC BUS 2, and the speed above 100 knots, the RAT will automatically extend to power both the hydraulic motor that pressurizes the blue system and the emergency generator via a hydraulic motor. The RAT can also be manually extended for these purposes by pressing the EMER ELEC PWR MAN ON switch on the EMER ELEC PWR PANEL located on the overhead panel. • •

•

The RAT can only be stowed on the ground. If the aircraft is powered, so are the RAT circuits, so don’t press the switch unless you intend to deploy the RAT; on the ground, someone might have a nasty surprise! The minimum speed for RAT operation is 140 knots

- RAT pressurizing the BLUE system. - ENG 1 pump switch is off - PTU switch is off

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A-320 STUDY NOTES LOAD ALLEVIATION ACCUMULATORS (LAF) During turbulence, the ailerons and spoilers 4 and 5 deflect rapidly to alleviate some of the wing structural loads; the LAF accumulators assist the hydraulic system with this function. PRIORITY VALVES The priority valves shut off hydraulic power to the flaps, slats, landing gear, nose wheel steering and the emergency generator if system pressure falls below a certain predetermined value. The reason for this is to ensure sufficient hydraulic pressure is available for the flight controls, brakes, spoilers, and thrust reversers. SYSTEM ACCUMULATORS •

The accumulators help maintain constant hydraulic pressure to the users during normal operations or heavy demands.

LEAK MEASUREMENT VALVES •

These are used by maintenance.

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A-320 STUDY NOTES

LANDING GEAR The green hydraulic system provides power for landing gear extension and retraction. In case of hydraulic, electrical or mechanical failure the gear can be extended by gravity using a hand crank that physically disengages the up-locks, opens the gear doors, depressurizes the system and allows the gear to free-fall. The hand crank, located aft on cockpit pedestal, must be rotated three full turns to enable the gravity extension. The nose wheel steering, normally powered by the green hydraulic system, will be lost after gravity extension and the doors will remain open.

LANDING GEAR CONTROL AND INTERFACE UNITS (LGCIU) Electrical signals are provided for gear and door actuation by two LGCIUs. The LGCIUs receive information from the landing gear, cargo door and flap systems. They then process gear and door position, sequencing, control and gear lever selection. The LGCIUs also send information and signals to the ECAM and other aircraft systems regarding the ground or flight mode; for example the active LGCIU will signal that the aircraft is on the ground and landing gear retraction will be inhibited; another example would be the inhibition of flap extension during cargo door operations on the ground. In the air, examples would be the inhibition of reverser operation or the closure of the hydraulic safety valve above 260 knots, preventing the gear being extended above this speed. There are at least 37 LGCIU outputs which are obviously beyond the scope of these notes; for more information see FCOM 1.32.10. p8. • LGCIU 1 provides landing gear position information to the landing gear indicator panel (next to the gear lever) and the ECAM. LGCIU 2 only provides this information to the ECAM, so if you lose LGCIU 1, you will loose the landing gear position information on the indicator panel. On the SD (WHEEL page) each landing gear position is indicated by two triangles. Each triangle is controlled by one LGCIU. • A green triangle indicates that its respective LGCIU detects a landing gear downlocked. • A red triangle indicates that a landing gear is in transit • No triangle indicates that a landing gear is uplocked • Amber crosses indicate an LGCIU failure

See next page

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A-320 STUDY NOTES

- Two red triangles indicate that both LGCIUs detect that the gear is in transit. - The amber STEERING indicates that either the nose wheel steering has failed or the antiskid feature has failed

- The amber L/G CTL indicates that the landing gear lever and landing gear position disagree

The DOWN arrow illuminates red with an ECAM warning if the landing gear is not downlocked with any of the following: •

•

Radio altimeter reading less than 750’ and both engines N1 below 75% or N! below 97% with one engine Radio altimeter less than 750’ and flap 3 or FULL

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A-320 STUDY NOTES NOSE WHEEL STEERING The nose wheel steering is electrically controlled by the Brake and Steering Control Unit (BSCU) and hydraulically operated by the green hydraulic system. The BSCU receives inputs from the rudder pedals and the Captain’s or First Officer’s steering hand wheels. • The rudder pedals deflect the nosewheel a maximum of 6 ° until 40 knots, and then progressively reducing to 0° at 130 knots. • The hand-wheels deflect the nosewheel a maximum of 75° until 20 knots, and then progressively reducing to 0° at 70 knots. The nose wheel steering can be deactivated by turning off the A/SKID NW STRG switch or by operating the towing lever on the nose wheel (used for pushing back or pulling the aircraft on the ground). Indeed, until the system is modified by Airbus, it is SOP to turn the A/SKID NW STRG switch OFF prior to push back, and turn it back on after push back is complete, thereby ensuring that the nose wheel steering is deactivated during the process. A green NW STRG DISC message appears on the ECAM after de-activation and this message turns amber after the second engine start. Green Hydraulic pressure will be applied to the nose wheel when all of the following conditions are met: • A/SKID NW STRG SWITCH is ON • At least one engine running • Towing control lever in normal position • Main landing gear are compressed • ADIRU 1 or 3 operative

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A-320 STUDY NOTES

BRAKES The brakes are carbon multi-discs. The normal braking system includes antiskid and autobrakes and is powered by the green hydraulic system. The alternate braking system is powered by the yellow system and has a brake accumulator. The alternate brakes may or may not have antiskid depending on the level of redundancy. A dual channel Brake and Steering Control Unit (BSCU) controls all brake functions including normal and alternate brakes, antiskid, autobrakes and temperature indications. Normal brake pressure is between 2000 psi and 2700 psi with full pedal deflection. Normal brakes are available when: • The A/SKID NW STRG switch is ON • The green hydraulic pressure is available • The parking brake is OFF ANTISKID The antiskid system ensures maximum braking capability by keeping the wheels at the limit of an impending skid. If a skid is detected by the BSCU, it will send a release signal to the normal and alternate servo valves and the ECAM. When a skid is no longer detected, the system returns to normal protection and detection. The antiskid is deactivated below 20 knots or when the A/SKID NW STRG is selected OFF. ALTERNATE BRAKES If the green system loses pressure, the alternate system with antiskid, takes over automatically, with the same capabilities as the normal system except, the autobrakes are not available (nose wheel steering will also be lost with a green system failure). Antiskid will be lost with alternate braking if: • • • •

There is an electrical failure affecting the system The BSCU fails The A/SKID NW STRG switch is OFF The yellow system fails, forcing the accumulator to take over

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A-320 STUDY NOTES The maximum brake pressure to be applied without antiskid is 1000 psi, read on the brake triple indicator (top of the green band for left and right brakes).

With normal brakes there will be no pressure indication. Alternate braking will show the actual braking pressure and the parking brake will also show the actual brake pressure. The accumulator pressure will always be shown and it should maintain its pressure for least 12 hours on the ground with no power on aircraft. Operation of the yellow pump for cargo door operations will charge the accumulator. If the yellow system fails and the accumulator takes over, the system is designed to allow 7 full braking applications.

BRAKE TEMPERATURE LIMITATIONS The maximum brake temperature allowed for take-off is 300° C. Brake temperatures are shown on the ECAM wheel page; they will indicate HOT when any brake temperature exceed 300° C. The hottest brake will show a green arc over its temperature indication, turning amber above 300° C. After landing delay using the bake fans for approximately five minutes or until you get to the gate, however they must be turned on whenever brake temperatures exceed 500° C. The reason for this is that carbon brakes actually perform better when they are warm and using the brake fans increase wear due to possible oxidation.

See next page

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A-320 STUDY NOTES Maintenance action is due if: •

•

• • •

The temperature difference between two brakes on the same gear exceeds 150° C and the temperature of either one of the brakes exceeds 600° C, or The temperature difference between two brakes on the same gear is greater than 150° C and the temperature of the lower brake is below 60° C, or The difference between the LH and RH brake’s average temperature is higher than or equal to 200° C, or A fuse plug has melted, or Any one brake temperature exceeds 900° C.

Avoid using the parking brake if the brake temperature exceeds 500° C. AUTOBRAKES Autobrakes are available with normal braking (green system) only. LO and Med can be used for landing and MAX is for take-off only. To arm the autobrakes, depress the switch for at least 1 second. During take-off the autobrakes will not activate below 72 knots (ground spoiler extension). During landing, with LO selected the autobrakes will activate 4 seconds after touchdown and in MED, 2 seconds after touch down. A green DECEL light in the autobrakes pb comes on to indicate that the actual deceleration is within 80% of the selected rate. (it does not indicate that the autobrake is activated) • The autobrakes can be armed with the parking brake on. The autobrake will be disarmed under the following conditions: • • • •

De-selection of the AUTO/BRK switch Ground spoiler retraction 10 seconds after landing gear retraction By applying sufficient brake pedal force during auto brake operation: • In Max mode, when both pedals are depressed • In LO/MED mode, when one pedal is depressed

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A-320 STUDY NOTES

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A-320 STUDY NOTES

PNEUMATICS The pneumatic system provides high pressure air for the following: • • • • • • •

Air conditioning Pressurization Engine starting Wing anti-icing Potable water tank pressurization Hydraulic reservoir pressurization Aft cargo heating

The high pressure air is supplied by the engine bleeds, APU bleed and external ground air units. Each source can be connected to the crossbleed duct where a cross-bleed valve enables interconnection. The pneumatic system is controlled and monitored by two Bleed Monitor Computers (BMCs). Basically, BMC 1 controls the left side engine and APU bleed systems and BMC 2, the right side engine bleed system. If BMC 1 fails, BMC 2 takes over all monitoring functions except engine 1 and APU leak detection. If BMC 2 fails, BMC 1 takes over all monitoring functions except for engine 2 leak detection. ENGINE BLEED SELECTION Bleed air is normally taken from the intermediate pressure (IP) stage of the high pressure compressor. If IP stage pressure is too low, then the high pressure (HP) stage automatically supplies bleed air through the HP valve. In flight, if the pressure is still too low, the BMC will signal the FADEC to increase engine RPM as required (usually might happen at idle thrust).

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A-320 STUDY NOTES TEMPERATURE AND PRESSURE REGULATION The bleed air temperature is controlled by a pre-cooler that uses fan air. The bleed air pressure is regulated by the bleed valves, which are downstream of the IP and HP valves. The bleed valves are electrically controlled and pneumatically operated valves act as pressure regulators as well as shutoff valves. There is an overpressure valve downstream of the bleed valves which will close to protect the system if the bleed valve fails to regulate high pressure. A bleed valve will close pneumatically if: • •

The upstream pressure falls below 8 psi There is reverse flow

A bleed valves will close electrically if: • •

The bleed pb is switched off An engine fire pb is pressed

The BMC will close the engine bleed valves for the following: • • • • •

APU bleed switch is ON Engine start valve not closed Over-temperature System leakage Over-pressure

APU When the APU bleed valve is opened, the BMC sends a signal to the crossbleed valve to open and the engine bleed valves to close. The cross-bleed valve has two electric motors; one is controlled automatically by the BMC and one manually. During normal operations with the APU bleed valve closed, the crossbleed will also close to isolate the two systems. The cross-bleed is opened manually during a cross-bleed engine start.

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A-320 STUDY NOTES PNEUMATIC SYSTEM Note: engine 2 bleed system omitted for clarity

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A-320 STUDY NOTES OVERHEAT AND LEAK DETECTION The system detects overheat conditions near the hot air ducts in the fuselage, pylons and wings. There are double loops in the wings and the fuselage back as far as the APU check valve and single loops for the pylons and APU. The following occur for an APU leak: • • •

The APU bleed valve closes The APU BLEED FAULT light illuminates The cross-bleed valve closes

For a wing leak: • • • •

The bleed valve closes on the related side The associated ENG BLEED FAULT light illuminates If open, the cross-bleed valve closes (except during engine start) If open, the APU bleed valve closes for a left wing overheat (except during engine start)

For a pylon leak: • •

The related bleed valve closes The associated ENG BLEED FAULT light illuminates

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A-320 STUDY NOTES OVERHEAT AND LEAK DETECTION

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A-320 STUDY NOTES

AIR CONDITIONING Pneumatic air from the cross-bleed duct passes through two electrically controlled, and pneumatically operated, pack flow control valves, which regulate the air flow in accordance with commands from the two pack controllers. The pack flow control valves or pack valves will close automatically for the following reasons: • Low air pressure • Pack overheat • Engine starting • Ditching (when ditching pb is pressed) • Engine fire switch pressed The hot air is then either ducted towards hot air pressure regulating valve and then to the trim air valves, or to the primary heat exchangers and the two packs. The cooled air from the packs then passes to the mixing unit where it mixes with recirculated cabin air and is ducted to one of the three zones: Cockpit, FWD cabin and AFT cabin. The hot trim air is mixed with the conditioned air after it leaves the mixing unit. The pack controllers regulate the temperature according to the demands of the zone controllers by modulating the pack valves, the pack turbine bypass valves as well as the ram air inlet and outlet flaps. During take-off the ram air inlet and outlet flaps close when TO power is set. During landing, they close as soon as the gear struts are compressed and re-open 20 seconds after reaching 70 knots. The zone controller modulates the airflow through the trim air valves to optimize the temperature regulation. The cabin zone commanding the coldest temperature will drive both packs to that temperature. Trim air is then added to the other zones to optimize their temperatures. EMERGENCY RAM AIR INLET If both packs fail, and for smoke removal, the ram air inlet can be opened to ventilate the cockpit and cabin. Do not confuse the emergency ram air inlet with the ram air inlet and outlet flaps operated by the pack controller. (Yes, I totally agree, they really could have come up with different names here!) Anyway, this one, the ‘emergency ram air inlet’ opens when the RAM AIR switch on the overhead panel is selected ON provided the differential pressure is less than 1 psi and the DITCHING switch is not selected ON. TRAINING PURPOSES ONLY - 31 -

A-320 STUDY NOTES When the ‘inlet’ opens, ram air is supplied to the mixing unit and the outflow valve opens 50%, if the CABIN PRESS MODE SEL is in AUTO. PACK CONTROLLER FAILURES •

•

•

Each pack controller has two channels known as the primary and secondary channels. If the primary channel fails, the secondary channel automatically takes over and the pack flow is fixed at the pre-failure setting. If the secondary channel fails the primary channel automatically takes over with no effect on the optimization however the ECAM signals are lost. If both controllers fail, the pack outlet temperature is then controlled by the pack anti-ice valve to a temperature between 5°C and 30°C in a maximum of 6 minutes. ECAM signals to corresponding pack are also lost.

ZONE CONTROLLER FAILURES Each zone controller has two channels, known as the primary and secondary channels. If the primary channel fails, the secondary channel takes over and: • The hot air and trim air valves close. • The zones are controlled to 24°C. • Pack one controls the cockpit temperature and pack two, the FWD and AFT cabin temperatures • ALTN MODE appears in the ECAM COND page If the secondary channel fails, there is no effect on zone temperature regulation, however the backup is lost. If both channels fail, optimized temperature regulation is lost and: • The packs deliver a fixed temperature: 20°C for pack 1 and 10°C for pack 2. • The failure removes all information from the ECAM COND page and PACK REG is displayed

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A-320 STUDY NOTES HOT AIR PRESSURE REGULATION FAILURE Fails open – no effect Fails closed – Optimization is lost • Trim air valves close • Pack 1 controls the cockpit to selected temperature • Pack 2 controls FWD and AFT zones to the mean of the selected temperatures. AIR CYCLE MACHINE FAILURE If an air cycle machine fails (compressor or turbine seizure), the affected pack may be operated in heat exchanger cooling mode with reduced flow.

NOTE: Do not connect an external LP air unit when the packs are running, it’s either one or the other, never both!

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A-320 STUDY NOTES AIR CONDITIONING SCHEMATIC

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A-320 STUDY NOTES

VENTILATION The fully automatic ventilation system cools the avionics compartment, the flight deck instruments and the circuit breaker panels. The system uses two continuously running electric fans to force the circulation of cooling air. There are two valves located in the aircraft skin on either side of the aircraft that facilitate the intake of cooling air, and the extraction of warm air overboard. The valve on the left side of the aircraft is called the skin air inlet valve and the one on the right side, the extract valve. Cooling air is sucked in through the inlet valve where it passes through a filter, through the blower fan, then through the avionics equipment. The warm air enters the extract duct where together with air from the cockpit panel ventilation system, is passed through the extract fan where it is extracted overboard, directed under the cargo compartment or recirculated through a skin air heat exchanger. VENTILATION SCHEMATIC (Open circuit configuration)

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A-320 STUDY NOTES The system has three normal configurations, determined by the ground/flight modes and aircraft skin temperature: • • •

Open configuration Closed configuration Intermediate configuration (in flight only)

OPEN CONFIGURATION The open configuration occurs on the ground only. Both inlet and extract valves are open, and air is extracted overboard. During heavy rain, it is recommended to select the EXTRACT pb to OVRD, which will close both the inlet and extract valves and prevent water entering the avionics bay. (both packs should be ON if this procedure is attempted). CLOSED CONFIGURATION This occurs on the ground and in flight. If the outside skin air temperature is below a certain threshold, both inlet and extract valves close automatically, and the air is circulated past the skin air heat exchanger, and exhausted below the cargo compartment. This configuration is the same as the abovementioned ‘rain’ procedure, except the BLOWER and EXTRACT pb are in the AUTO position. INTERMEDIATE CONFIGURATION This occurs only in flight if the skin air temperature is above a certain threshold. This configuration is the same as the closed circuit, except that the extract valve opens partially to allow the exhaust of hot air overboard. ABNORMAL CONFIGURATION This occurs when either the BLOWER or EXTRACT pb are in the OVRD position. The system reverts to the closed circuit configuration except the cooling air is supplied by the air conditioning system. When the BLOWER switch is in OVRD, the blower fan stops and the extract fan keeps running. When the EXTRACT switch is in the OVRD position, both fans continue to run. Either a BLOWER FAULT or EXTRACT FAULT, ECAM warning will be displayed.

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A-320 STUDY NOTES

SMOKE CONFIGURATION If an AVIONICS SMOKE warning is triggered by the sensor beside the extract fan, the BLOWER and EXTRACT lights will illuminate on the VENTILATION panel. Both are switched to OVRD during the ECAM procedure which leads to the blower fan stopping and the extract fan continuing to run. The cooling air is provided by the air conditioning system and all ventilation air is extracted overboard in an attempt to clear the smoke.

BATTERY VENTILATION A venture in the skin of the aircraft draws air from the space around the batteries and vents it overboard, thereby ventilating the batteries. LAVATORY AND GALLEY An extraction fan draws air from the galleys and lavatories and exhausts it overboard. CARGO VENTILATION The cargo compartments are ventilated with cabin air. An extraction fan draws air from the cargo compartments and exhausts it overboard. The system can add hot bleed air to the cabin air entering the cargo compartments, thus heating them.

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A-320 STUDY NOTES

PRESSURIZATION The pressurization system is made up of two identical, independent, automatic systems (system 1 or system 2) that control the cabin altitude. The overall system consists of two cabin pressure controllers (CPCs), one outflow valve, two safety valves and a control panel. The outflow valve actuator incorporates three electric motors; two for automatic control and one for manual control. In normal operation the system is operated in automatic control, however it can also be operated in semi-automatic or manual control. The CPCs receive the landing elevation, QNH and pressure schedule from the FMGS, and the pressure altitude from the ADIRS. The engine Interface Unit (EIU) supplies the throttle lever position and the LGCIU sends the flight/ground signal. In case of FMGC failure, the crew sets the landing elevation manually by pulling the LDG ELEV knob out of the auto detent, rotating it to the desired landing elevation and the system then uses its own internal pressure schedule (semi-automatic control). In this case, the controller uses the Captain’s Baro reference from the ADIRS. In manual mode the crew controls the cabin altitude by use of the switch control on the CABIN PRESS panel, which in turn operates the manually controlled electric motor on the outflow valve actuator. The safety valves prevent cabin pressure from exceeding 8.6 psi or going below 1 psi below ambient pressure. AUTOMATIC PRESSURE CONTROL SEQUENCE Ground Before takeoff and 55 seconds after landing, the outflow valve fully opens to ensure that there is no residual cabin pressure. At touchdown, any remaining cabin pressure is released at a cabin vertical speed of 500 ft/min. An automatic transfer of systems occurs after each landing. Takeoff To avoid a pressure surge or ‘bump’ at rotation, the cabin is prepressurized when the thrust levers are advanced, at a rate of 400 ft/min until ΔP reaches 0.1 psi. At lift off the CPC initiates the climb phase.

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A-320 STUDY NOTES PRESSURIZATION SCHEMATIC

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A-320 STUDY NOTES PRESSURIZATION SCHEMATIC

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A-320 STUDY NOTES

FIRE PROTECTION Fire protection and detection is provided for the engines and APU. Each engine and the APU have dual fire and overheat detection loops installed. The engine loops have overheat sensors located in the pylons, engine cores, nacelles and fan sections. The APU has an overheat sensing element in its compartment. There are two extinguishing agents for each engine and the one for the APU. Smoke detectors are installed in the lavatories, cargo compartment and avionics compartments. The lavatory waste bins are provided with automatic extinguishers and the cargo compartments have fire extinguishing capability. DETECTION LOOPS If an overheat sensing element detects an overheat, a signal is sent via the loops to the FIRE DETECTION UNIT (FDU) and a fire warning is triggered. If a fault or break occurs in one loop, the system is not affected as fire detection is still available by the unaffected loop. Engine or APU fire warning appear under the following conditions: • • • •

Both loops detect a fire One loop detects a fire when the other loop is faulty A break or fault in both loops occurs within 5 seconds of each other A fire test is performed

ENGINE AND APU FIRE WARNING An engine fire warning is indicated by the following: • • • • •

A continuous repetitive chime (CRC) Illumination of the ENG FIRE light on the overhead panel Illumination of the MASTER WARNING light on the glare shield Illumination of the FIRE light on the ENG start and ignition panel ECAM warning and activation

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A-320 STUDY NOTES

An APU fire is indicated by: • • • •

A continuous repetitive chime (CRC) Illumination of the MASTER WARNING light on the glare shield Illumination of the APU FIRE light on the overhead panel ECAM warning and activation

If the APU fire occurs on the ground, a warning horn will sound in the nose wheel well and a APU FIRE light will illuminate in the external service interphone panel. ENGINE AND APU FIRE PROTECTION Each engine is provided with two fire extinguishing agents and the APU with one. The agent DISCH pbs are located on the overhead panel. If the APU fire occurs on the ground, the APU will auto-shutdown and the agent will discharge automatically. In flight the APU must be manually shut down and the agent discharged manually. The APU will auto–shutdown in flight for a number of other reasons: for example low oil pressure, high oil temperature etc., for further details refer to the APU chapter. Each agent has an electrically operated squib which is armed when the FIRE switch is pressed; pressing the SQUIB/DISCH pb will discharge the agent if required.

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A-320 STUDY NOTES Engine fire switch (guarded) When an engine fire switch is pressed the following occurs: • • • • • • •

Cancels the aural warning Arms the squibs Deactivates the generator Closes the fuel LP valve Closes the hydraulic fire shutoff valve Closes the engine bleed Closes the pack flow control valve

APU fire switch (guarded) When an APU fire switch is pressed the following occurs: • • • • • •

Cancels the aural warning Shuts down the APU Arms the squib Closes the fuel LP valve APU fuel pump is turned off APU bleed and crossbleed valves close

LAVATORY PROTECTION AND DETECTION A smoke detector is installed in each lavatory ventilation extract duct. If smoke is detected in a lavatory by the Smoke Detection Control Unit (SDCU), a signal is sent to the flight warning computer and ECAM in the cockpit and a LAV SMOKE warning appears on the forward flight attendant panel. If sufficient heat is detected in or around the lavatory waste bin, the Halon extinguisher will automatically discharge through one of the heat activated nozzles. CARGO PROTECTION AND DETECTION The forward cargo hold has 2 smoke detectors and the aft cargo hold, 4. One detector in the forward hold is connected to one of the two detection loops and the other to the remaining one. In the aft cargo, there are two detectors connected to each loop. The SDCU receives a

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A-320 STUDY NOTES signals from the loops and transmits them to the ECAM, which displays a smoke warning. A smoke warning is signal is sent if both loops detect it or one loop detects it and the other one is inoperative. The cargo isolation valves close automatically if a smoke warning is detected. There is one extinguisher bottle that serves both the FWD and AFT cargo compartments. The bottle has two discharge heads that serve three nozzles; one in the FWD and two in the AFT. If smoke is detected the following occurs: • • • •

A continuous repetitive chime (CRC) Illumination of the MASTER WARNING light on the glare shield Illumination of the SMOKE light on CARGO SMOKE panel ECAM warning and activation

When a crewmember presses the appropriate DISCH pb switch on the CARGO SMOKE panel, the action ignites the relevant squib and the agent is discharged to the appropriate compartment. The SMOKE light may remain on after the agent is discharged because of the effect of the smoke and agent on the smoke detectors. The DISCH light illuminates within 60 seconds of being activated.

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A-320 STUDY NOTES

FUEL There are seven fuel tanks; three in each wing and one in the fuselage. The main wing tanks are divided into two cells, inners and outers. Each wing also has a vent surge tank, that is located outboard of the outer wing tanks to compensate for expanding fuel. When the aircraft has been refueled to maximum capacity, the vent surge tanks allow for 2% expansion or a 20°C rise in fuel temperature. Fuel from the vent tanks drains into the outer cells when capacity is available. The tank in the fuselage is called the centre tank. Fuel cannot be transferred from tank to tank except on the ground during refueling operations. FUEL PUMPS In normal operation, each engine is fed either one pump in the centre tank or two pumps from an inner wing tank. The wing tanks are fitted with sequence valves to ensure that when all pumps are running, the centre tank will supply fuel preferentially. CROSSFEED VALVE The cross feed valve has two electric motors and can be used to balance the fuel load by feeding two engines from one tank or one engine from two tanks. TRANSFER VALVES Two transfer valves are installed to transfer fuel from the outer tanks to the inner tanks. The transfer takes place when the inner tanks are depleted to 750 KGS. When open, the valves remain open until the next refueling, when they close. During steep descents or accelerations/decelerations, the valves may open before the 750 KG threshold. SUCTION VALVES The suction valves are closed when there is normal fuel pressure from the boost pumps. The valves open to allow gravity feed from the wing tanks, in case of electrical failure. The centre tank pumps are not fitted with suction pumps, so gravity feed is not possible from the centre tank.

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A-320 STUDY NOTES ENGINE LP VALVES The LP valves allow the fuel to be cut off form the engines; the valves can be closed by the engine master switches or ENG FIRE pbs. APU FEED The APU has its own fuel pump, used for APU start when fuel pressure is low (due to a loss of AC power). The APU fuel supply is from the left side of the fuel manifold. NORMAL FUEL FEED SEQUENCE All pumps are switched on prior to start up and the MODE SEL is checked to be in AUTO. After both engines are started, if there is fuel in centre tank, the pumps will run for a 2 minute test sequence, and then shutoff until the slats are retracted when airborne or re-selected down; this prevents take-off and landing using the centre tank. The centre tank pumps will then run until 5 minutes after the centre tank is empty. The wing tank pumps run continuously. An ECAM warning will be given during AUTO operation, if the centre tank has more than 250 KGS of fuel and either inner tank has less 5000 KGS. IDG RECIRCULATION COOLING SYSTEM Fuel is directed through the IDG heat exchanger and then back through the fuel return valve to the outer tanks. The fuel return valve is controlled by the FADEC. If the outer tanks are already full, the recirculated fuel spills into the inner tanks though a spill pipe. If the inner tanks are already full, then the centre tank pumps are automatically turned off to allow fuel to burnt from the inner tanks. The inner tanks will supply the engines until approximately 500 KGS of fuel has been used, and then the centre tanks will resume operation.

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A-320 STUDY NOTES IDG Recirculating System

USABLE FUEL QUANTITIES (SG 0.785 kg/l)

VOLUME WEIGHT

LITRES (KG)

USABLE OUTER TKS 880 x 2 691 x 2

FUEL INNER TKS 6924 x 2 5435 x 2

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CENTRE TK

TOTAL

8250 6467

23,858 18,728

A-320 STUDY NOTES

FUEL SYSTEM SCHEMATIC

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A-320 STUDY NOTES

FLIGHT CONTROLS All flight control surfaces are electrically controlled, and hydraulically actuated. Additionally, the stabilizer and rudder can if necessary, be mechanically controlled and hydraulically actuated. The flight control system consists of two sidesticks, two autopilots, two elevator aileron computers (ELACS), three spoiler elevator computers (SECS), two slat flap control computers (SFCC) and two flight augmentation computers (FACS). The basic principle of the fly-by-wire system is shown below:

Commands Autopilot

Electrical orders

Sidestick

Digital computers ELACS (2)

SECs (3)

FACs (2)

Slats/flaps

Mechanical back up

Electro/hydraulic actuators

SFCCs (2)

Hyd. jacks

Elevator Stabilizer Ailerons Spoilers Rudder Slats Flaps

Rudder pedals

In normal operations, ELAC 2 commands the operation of the elevators and horizontal stabilizer, ELAC 1 commands the operation of the ailerons, and SECs 1, 2, and 3 command the spoiler operations. The FACs control the rudder and yaw damper functions, make flight envelope and speed computations and assist with windshear protection. FAC 1 is primary and FAC 2 standby. The SFCCs command the operation of the flaps and slats.

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A-320 STUDY NOTES

A320 flight controls surfaces

Rudder Elevator Slats Aileron

Flaps Trimmable horizontal stabilizer (THS)

Speed brakes Roll spoilers Ground spoilers

Load alleviation function (LAF)

Sidesticks The two sidesticks are not interconnected and provide no feedback, however they do have artificial feel and are spring loaded to neutral. The sidesticks essentially send electric roll and pitch signals to the flight control computers. If both sidesticks are operated simultaneously, the signals are algebraically added, up to a maximum of a single stick deflection. It is therefore very important that only one stick is used at a time. In other words, if one pilot commands a pitch up and the other a pitch down, then there is a possibility of canceling each other out and likewise if both pitch up, the added signals will probably lead to overcontrol; the same goes for roll control. There is a priority pushbutton on either sidestick to help avoid this problem, and an aural warning generated that announces DUAL INPUT. In addition, the priority pushbutton is used to disconnect a failed sidestick or take over control. In this case, the switch must be held down for 30 seconds to disconnect the failed stick. If a pilot presses the priority pushbutton to take over control, the system annunciates this on the glareshield SIDE STICK PRIORITY panel and there is an audio warning, PRIORTIY LEFT or PRIORITY RIGHT. To

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A-320 STUDY NOTES re-instate the failed or deactivated stick, the priority pushbutton on either stick must be pressed. The last person to switch has priority. If a stick fails or is deactivated on the ground prior to take-off, a CONFIG R (L) SIDESTICK FAULT warning is activated. ROLL CONTROL Roll control is achieved by one aileron and four spoilers per wing. Ailerons The green and blue hydraulic systems power two servojacks on each aileron. One servojack is active while the other is in damping mode. If the active jack fails then the other one takes over. Both jacks operate when the load alleviation function (LAF) is active during turbulence. The system automatically selects dual damping mode in the case of a dual ELAC failure, or for loss of both green and blue hydraulics. The ailerons droop 5° when the flaps are extended. Spoilers The four outboard of the five spoilers on each wing, assist with roll control, all five act as ground spoilers, the two outboard panels are used for the LAF and the three middle panels serve as speedbrakes. The five panels are controlled by the three SECs and operated by either the green, blue or yellow hydraulic systems. If a spoiler fails on one wing, the symmetrically opposite panel on the other wing is automatically deactivated. If a SEC fails or there is an electrical failure, then the affected spoilers automatically retract, however if there is a hydraulic failure, then the affected spoiler will remain extended or retract due to aerodynamic force. Ground spoilers Ground spoilers extend automatically during a rejected takeoff when the wheel speed is above 72kts. and either: • Both thrust levers are at forward idle and the ground spoilers are ARMED, or • Reverse thrust is selected on at least one engine with the other thrust lever at idle and the spoilers are not ARMED.

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A-320 STUDY NOTES Ground spoilers extend automatically on landing if both main gear touch down and: • Both thrust levers are at forward idle and the ground spoilers are ARMED, or • Reverse thrust is selected on at least one engine with the other thrust lever at idle and the ground spoilers are not ARMED Ground spoilers will extend partially after landing on one wheel if: • Reverse thrust is selected on at least one engine with the other thrust lever at idle This eases compression of the second landing gear strut when it touches down, and makes for a smoother landing in this situation than otherwise. The spoilers will extend fully on the second wheel touchdown. During a touch and go landing, the spoilers will retract when the thrust levers are advanced beyond 20° TLA (thrust lever angle). After a bounced landing, the spoilers remain extended if the thrust levers remain at idle. Speed brakes A green SPEED BRAKE memo appears on the ECAM when speedbrakes are in use. The memo flashes amber if the speedbrakes are extended and the thrust levers are not at idle. The speedbrakes are inhibited if: • SEC 1 and 3 fail • The L(R) elevator has a fault (spoilers 3,4 only inhibited) • Angle of attack protection is active • Flaps in configuration FULL • Thrust levers are above MCT • Alpha floor is activated If an inhibition occurs with the speed brakes extended, they will retract and stay retracted until the inhibition disappears or the speed brake lever is returned to the RET position for 10 seconds or more. When the aircraft is flying above .75 Mach or 315kts with the autopilot engaged and the speed brake extended, it may take up to approximately 50 seconds for them to retract. This is to avoid activation of the high angle of attack protection. Speed brakes should not be extended beyond halfway when below .75 Mach and above FL 310, again, to avoid activation of high angle of attack protection. TRAINING PURPOSES ONLY - 53 -

A-320 STUDY NOTES

PITCH CONTROL Pitch control is achieved with the elevators and the trimmable horizontal stabilizer (THS). They are both electrically controlled by the ELACs or SECs, and are hydraulically actuated. If both ELACs fail the SECs take over. The stabilizer can be controlled mechanically through cables attached to the cockpit pitch trim wheels, provided there is hydraulic power available. Mechanical pitch trim has priority over electric trim. Elevators The blue and green hydraulic systems power two servojacks on the left elevator and the yellow and blue, two on the right. One jack is active while the other is in damping mode. If the active jack fails then the other one takes over. If electrical control is lost to the jacks, they both go into centering or streamlined mode. If hydraulic control is lost to the jacks, they both go into damping mode. If one elevator is lost, the other will operate at reduced deflection to avoid excessive asymmetric loads on the tail. Trimmable horizontal stabilizer (THS) The stabilizer is electrically controlled by one of three electric motors or mechanically controlled via the trim wheels. A screwjack is hydraulically driven by the green and yellow hydraulic systems to drive the THS. After touchdown the trim returns to 0° automatically.

YAW CONTROL Yaw control is achieved by the rudder. Yaw orders for turn coordination and yaw damping are computed by the ELACs and transmitted to the FACs. Rudder The rudder is electrically controlled by the trim motors or mechanically by the rudder pedals. In mechanical control hydraulic pressure is still required to move the surface. Three independent servojacks, operating in parallel, operate the rudder. In automatic operation (turn coordination and yaw damping) a green hydraulic servo actuator drives all three servojacks. A yellow system actuator is always synchronized and takes over if there is a failure. TRAINING PURPOSES ONLY - 54 -

A-320 STUDY NOTES Rudder travel is limited as a function of speed. If both FACs fail, maximum rudder deflection will be obtained when the slats are extended. The rudder trim is operated by the number one electric motor, controlled by FAC 1, number 2 motor is in synchronized in backup with FAC 2. Electric trim with the autopilot engaged operates at 5°/second up to a maximum of 20°. Engine failure compensation is calculated by the FACs and FMGC as a function of speed, engine power available, bank angle and yaw. Manual rudder trim operates at 1°/second up to 20°. Manual trim is not available when the autopilot is engaged. Yaw damper The yaw damper receives inputs form the ELACs and FACs. The information is sent to the yaw damper servo for damping and turn coordination. There is no feedback from the rudder pedals during these functions.

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A-320 STUDY NOTES

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A-320 STUDY NOTES FLAPS AND SLATS The five slats and two flaps on each wing are electrically controlled and hydraulically actuated via inputs form the flap lever. Signals from the flap lever are transmitted to the two flap slat control computers (SFCC). The slats are operated by the green and blue hydraulic systems and the flaps by the green and yellow systems. If one hydraulic system fails its associated surfaces will operate at half speed. If one SFCC fails, all the slats and flaps will operate at half speed. There four hydraulically operated wing tip brakes (WTB)to stop the slats or flaps moving in case of asymmetry, overspeed, uncommanded movement and symmetrical runaway. The wing tip brakes cannot be rest in flight, and furthermore, their circuit breakers (red collared) must never be pulled in flight. If the slats are locked out by the wingtip brakes, it is still permissible to operate the flaps and vice versa. There is also a flap disconnect system, that operates if there is excessive differential movement detected, between the inner and outer flap on each wing; this might occur if one of the surfaces attachments fail and the system is therefore designed to stop further damage. If the slats are locked out by the wingtip brakes, it is still permissible to operate the flaps and vice versa. There is an alpha lock system that stops the flaps from being retracted at excessive angles of attack or low airspeed. The inhibition is removed when the angle of attack is decreased or the speed is increased. Flaps 1 + F When flaps are selected to position 1 or CONFIG 1 on the ground, they will extend to position 1+F, which is Slats 18° and Flaps 10°. After takeoff, if the airspeed goes above 210kts, with the aircraft still in CONFIG 1, the Flaps will automatically retract to 0° and the Slats will remain extended until they are retracted, by physically moving the flap lever to position 0. During flight, when the flap lever is selected to CONFIG 1, only the Slats will extend to 18°. The first Flap movement will be at CONFIG 2, where the flaps will extend to 15°. After flap retraction, CONFIG 1 + F is not available until the speed is below 100kts, unless flaps 2 or more has been selected and the flaps are retracted again i.e. during a go-around.

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A-320 STUDY NOTES

POSITION

SLATS

FLAPS

1 1+F 2 3 FULL

18 18 22 22 27

O 10 15 20 40

MAX SPEED 230 215 200 185 177

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REMARKS Initial App. Takeoff Takeoff/App. Takeoff/App./ Ldg. Landing

A-320 STUDY NOTES

FLIGHT CONTROL LAWS NORMAL LAW In normal law there are three modes: • Ground mode • Flight mode • Flare mode Ground mode In the ground mode, the aircrafts flight control characteristics are very similar to those of a conventional aircraft, where there is a direct relationship between sidestick deflection and control surface deflection. Pitch Pitch trim is set manually prior to takeoff and there is no autotrim until airborne, however after landing the pitch trim automatically resets to 0°. During takeoff, and above 70kts, the maximum elevator deflection is reduced from 30° to 20°. Once airborne, the flight mode is progressively blended in. Flight mode The flight mode is active from lift off until the flare mode engages at 50’ AGL. Pitch The normal law flight mode is a load factor demand law, with autotrim and full flight envelope protection. The aircraft maintains 1g in pitch (corrected for pitch attitude) with the sidestick in neutral and wings level. Once a turn is established the pilot does not have to make any trim corrections in pitch because there is autotrim, both in manual and automatic flight up to 33° of bank. Essentially, forward or aft movement of the sidestick commands the elevators and stabilizer trim to achieve a load factor proportional to the amount of stick deflection. The same amount of stick movement produces the same load factor regardless of the speed, assuming the same bank angle.

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A-320 STUDY NOTES Automatic pitch trim is inhibited in the following circumstances: • • • • • • •

Radio altitude below 50’ (100’ with the autopilot engaged Load factor less the 0.5g Load factor greater than 1.25g (nose-up trim is inhibited) High speed/Mach protection is active 33° of bank is exceeded When angle of attack protection is active (nose-up trim is inhibited) Manual trim being used (inhibited until trim wheel is released)

PROTECTIONS

Load factor limitation

The load factor is automatically limited to: • +2.5g to -1.0g flaps retracted • +2.0g to 0.0g flaps extended

Pitch attitude protection

Pitch attitude is limited to: • 30° nose up with flaps 0-3 (progressively reduced to 25° at low speed) • 25° nose up with flaps full (progressively reduced to 20° at low speed) • 15° nose down The flight director bars are removed when pitch attitude exceeds 25° up or 13° down.

High angle of attack protection Consists of three angle of attack functions; alpha floor ( floor), alpha protection ( prot) and alpha maximum ( max). The associated speeds vary with the weight and configuration and are referred to as V floor, V prot and V max. The elevator changes from normal mode to protection mode when the angle of attack exceeds V prot. When the aircraft speed reduces, and as a result the angle of attack increases into the prot range, the autopilot trips off, the speedbrakes retract, and the sidestick directly max., the controls angle of attack. If the speed further decays to system will not allow a higher angle of attack, even with full back-stick. If the sidestick is released in this condition, the aircraft will automatically TRAINING PURPOSES ONLY - 60 -

A-320 STUDY NOTES fly to prot and maintain that speed until corrective action is taken; the sidestick must be pushed forward to re-enter normal mode. This function provides protection against stall and windshear and will override any other protection. On takeoff, V prot = V max for 5 seconds. Alpha floor occurs between V prot and V max. At the predetermined speed, the autothrust engages automatically and TOGA thrust is provided, regardless of thrust lever position. • • •

floor protection is not available below 100’ RA altitude. floor protection is only available in Normal Law If the autothrust disconnect pb is pressed for more than 15 seconds, autothrust and its associated functions are lost, including floor, for the remainder of the flight.

PFD speed-tape indications

“SPEED SPEED SPEED” The low energy warning is activated and repeated aurally 5 times, when the speed is decreasing and approaching floor. Thrust must be added to recover to a safe angle of attack. The FACs compute the warning threshold based on the configuration (flaps 2, 3 or full), the rate of deceleration and angle of attack. The warning is inhibited when: • TOGA is selected • Below 100’ RA and above 2000’ • When floor or GPWS is triggered • In alternate or direct law • If both RAs fail •

Flaps ≤ 1 TRAINING PURPOSES ONLY - 61 -

A-320 STUDY NOTES

High speed protection

If the speed exceeds VMO + 6kts or Mmo + .01, the high speed protection is triggered and a pitch up command is sent to the elevator to prevent a further increase in speed. After activation the pitch trim is frozen, and the autopilot disconnects; additionally, positive spiral static stability is introduced so that if the sidestick is released the aircraft will roll, wing level to 0° bank angle instead of the normal law, 33° (see bank angle protection). If the sidestick is held fully forward in this regime, the speed will increase to a maximum of VMO + 30 kts or Mmo + .07, but will then reduce automatically to VMO + 16 kts or Mmo + .04. The speed will return to VMO when the stick is released and the protection is deactivated.

Bank angle protection

The aircraft maintains positive static spiral stability up to 33° of bank. The aircraft will maintain its bank angle up to this threshold with the side stick in neutral. For bank angles up to and including 67°, the aircraft rolls back to 33° when the stick is released. Full lateral control in normal law is limited to 67° which equates to the normal 2.5g limit. If the angle of attack protection or high speed protection is active, then the bank angle is limited to 45°, and the aircraft rolls wings level if the stick is released to neutral. Auto trim will also be frozen in this condition. In flight, if the bank exceeds 45°, the FD bars disappear and the autopilot disengages; the FDs return again when the bank angle is reduced to 40°. Auto trim is inoperative when bank angle protection is operating i.e. beyond 33° bank angle.

Load alleviation function (LAF)

The LAF relieves wing structural loading in turbulent conditions. The ELACs and SECs monitor the total load factor and deflect the ailerons and spoilers 4 and 5 symmetrically upwards to compensate if necessary. The LAF is inhibited in the following conditions: • Flap lever not in 0 • Airspeed below 200 kts or above VMO + 10 • Slats or flaps wing tip brakes engaged • Pitch in alternate law, without static stability protection • Pitch in direct law

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A-320 STUDY NOTES

Flare mode The flight mode changes to flare mode at 50’ RA. The system memorizes the pitch attitude at 50’ as a reference and at 30’, gradually feeds in a nose down pitch; reducing it to 2° nose down over 8 seconds. The pilot is therefore required to feed in nose-up authority to flare the aircraft.

ALTERNATE LAW If multiple failures occur the flight control system might revert to Alternate or Direct Law. There is no alternate law for roll control; it reverts to direct law always. In addition, when the aircraft is in alternate law, pitch control will always revert to direct law when the gear is extended; this is because alternate law does not have a flare mode, therefore the pilot must flare the aircraft conventionally like any other non - fly by wire aircraft. When the aircraft goes onto alternate law, the ECAM failure message will be: F/CTL ALTN LAW (PROT LOST) Ground mode The ground mode is exactly the same as in normal law Flight mode The flight mode differs to normal law in that some of the flight envelope protections are lost or degraded. Pitch In alternate law, all pitch protections are lost except for load factor maneuvering limits. VMO is reduced to 320 kts due to reduced high speed protection.

Low speed stability

This replaces the angle of attack protections of normal law. The low speed stability protection is activated if the speed approaches 5–10 kts above the stall warning. A gentle nose down pitch command is initiated which attempts to keep the speed from reducing further. This command

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A-320 STUDY NOTES can be overridden by the pilots and the aircraft can be stalled in

alternate law.

The stall warning consists of a ‘cricket’ audio warning signal and a STALL STALL synthesized voice warning. In alternate law the PFD speed scale is modified, where alpha prot and alpha max are replaced with a red and black barber pole; the top of which is the stall warning speed (Vsw).

High speed stability

If the aircraft exceeds Vmo/Mmo in alternate law, a gentle nose up command is initiated to keep the speed from increasing any further. This command can also be overridden by the pilots. Flare mode There is no flare mode in alternate law. Roll Roll control always degrades to direct law; there is no roll alternate law. Yaw In alternate law turn coordination is lost but yaw damping is available.

ALTERNATE LAW WITHOUT SPEED STABILITY After certain failures, the flight control capabilities will be degraded to alternate law, without speed stability, which means that both the low and high speed stability functions are lost. Only load factor functions and yaw damping is available, however yaw damping is lost in the case of a triple ADR failure. DIRECT LAW If the aircraft is in alternate law, then it will revert to direct law when the landing gear is lowered. It is also possible for the aircraft to go straight to direct law from normal law for example, with a triple IRS failure. There are no protections available in direct law, except for the aural stall warning and overspeed warning.

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A-320 STUDY NOTES Both roll and pitch control in direct law give a direct stick to control relationship, where the input is directly proportional to the surface movement. An amber USE MAN PITCH MESSAGE appears on the PFD when the aircraft is in direct law. ABNORMAL ATTITUDE LAW This law is activated if the aircraft enters an unusual or extreme attitude, for example a bank angle of 125º or more, airspeed greater than 400 kts or less than 60 kts, a pitch attitude of 50º nose up or 30º nose down. When the aircraft is in abnormal attitude law, the following will occur: • • •

Pitch in alternate law without speed stability Roll in direct law Yaw in alternate law without yaw damping

MECHANICAL BACK-UP In the event of a complete loss of electrical signals to the flight control systems, the aircraft will enter the mechanical back-up mode, where pitch is controlled through the manual stabilizer trim wheel and lateral control provided by the rudder pedals. There must be hydraulic power available to utilize this capability. A red MAN PITCH TRIM ONLY appears on the PFD in this situation.

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FLIGHT CONTROL LAWS SUMMARY NORMAL LAW

Ground Mode

Flight Mode

Flare Mode

• •

Active when aircraft is on the ground.

• •

Is active until shortly after liftoff.

• •

Becomes active shortly after takeoff and remains active until shortly after touchdown.

• • • • • • •

With sidestick neutral and wings level, system maintains a 1 g load in pitch.

• •

Transition to flare mode occurs at 50' RA during landing.

•

In the event of a go-around, transition to flight mode occurs again at 50' RA.

Direct relationship between the sidestick deflection and deflection of the flight control surfaces. After touchdown, ground mode is reactivated and resets the stabilizer trim to zero

Sidestick deflection and load factor imposed on the aircraft are directly proportional, regardless of airspeed. No requirement to change pitch trim for changes in airspeed, configuration, or bank up to 33º. Sidestick roll input commands a roll rate request. Roll rate is independent of airspeed. A given sidestick deflection always results in the same roll rate response. Turn coordination and yaw damping are computed by the ELACs and sent to the FACs. No rudder pedal feedback for the yaw damping and turn coordination functions.

System memorizes pitch attitude at 50' and begins to progressively reduce target pitch to 2º, forcing pilot to flare the aircraft

Load factor Limitation

•

Prevents pilot from overstressing the aircraft even if full sidestick deflections are applied.

Attitude Protection • Pitch limited to 30º up, 15º down, and 67º of bank. Pitch limits change as a function of configuration and speed • These limits are indicated by green = signs on the PFD. • Bank angles in excess of 33º require constant sidestick input. • If input is released the aircraft returns to and maintains 33º of bank. • FD command bars disappear at 45º of bank, re-appear at 40º. High Angle of Attack Protection (alpha): Protections

•

When alpha exceeds alpha prot, elevator control switches to alpha protection mode in which angle of attack is proportional to sidestick deflection.

•

Alpha max will not be exceeded even if the pilot applies full aft deflection

High Speed Protection:

• • •

Prevents overspeed. Adds a pitch up load factor demand at VMO + 6 kts M MMO + .07. The pilot cannot override the pitch up command. Pitch trim is frozen, autopilot disconnects, positive static spiral stability is introduced

Low Energy Warning:

• •

Available in CONF 2, 3, or FULL between 100' and 2,000' RA when TOGA not selected. Produces aural "SPEED SPEED SPEED" when change in flight path alone is insufficient to regain a positive flight path (Thrust must be increased).

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ALTERNATE LAW If Multiple Failures occur, the flight controls revert to Alternate Law. The ECAM displays the message: ALTN LAW: PROT LOST Ground Mode

Flight Mode

Protections

The ground mode is identical to Normal Law.

•

In pitch alternate law the flight mode is a load factor demand law similar to the Normal Law flight mode, with reduced protections.

•

Pitch alternate law degrades to pitch direct law when the landing gear is extended to provide feel for flare and landing, since there is no flare mode when pitch normal law is lost.

• • •

Automatic pitch trim and yaw damping (with limited authority) is available.

• • • •

All protections except for load factor maneuvering protection are lost.

•

The PFD airspeed scale is modified: o V LS remains displayed o V ALPHA PROT and V ALPHA MAX are removed o They are replaced by a red and black barber pole, the top indicating the stall warning speed V SW

Turn coordination is lost. When pitch law degrades from normal law, roll degrades to Direct Law - roll rate depends on airspeed.

The load factor limitation is similar to that under Normal Law. Amber XX's replace the green = attitude limits on the PFD. A low speed stability function replaces the normal angle-of-attack protection o System introduces a progressive nose down command which attempts to prevent the speed from decaying further. o This command CAN be overridden by sidestick input. o The airplane CAN be stalled in Alternate Law. o An audio stall warning consisting of "crickets" and a "STALL" aural message is activated. o The Alpha Floor function is inoperative.

•

A nose up command is introduced any time the airplane exceeds V MO /M MO to keep the speed from increasing further, which CAN be overridden by the sidestick.

• • •

Bank angle protection is lost. Certain failures cause the system to revert to Alternate Law without speed stability. Yaw damping is lost if the fault is a triple ADR failure.

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ABNORMAL ALTERNATE LAW Abnormal Alternate Law is activated if the airplane enters an unusual or extreme attitude

• • •

Pitch law becomes Alternate (without autotrim or protection other than Load Factor protection).

•

There is no reversion to Direct law when the landing gear is extended.

Roll law becomes Direct law with mechanical yaw control. After recovery from the unusual attitude, the following laws are active for the remainder of the flight: o Pitch: Alternate law without protections and with autotrim. o Roll: Direct law o Yaw: Alternate law

DIRECT LAW Direct law is the lowest level of computer flight control and occurs with certain multiple failures.

•

Pilot control inputs are transmitted unmodified to the control surfaces, providing a direct relationship between sidestick and control surface.

• • •

Control sensitivity depends on airspeed and NO autotrimming is available.

• •

There are no protections provided in Direct Law, however overspeed and stall aural warnings are provided.

An amber message USE MAN PITCH TRIM appears on the PFD. If the flight controls degrade to Alternate Law, Direct Law automatically becomes active when the landing gear is extended if no autopilots are engaged. If an autopilot is engaged, the airplane will remain in Alternate Law until the autopilot is disconnected. The PFD airspeed scale remains the same as in Alternate Law. MECHANICAL BACKUP

In case of a complete loss of electrical flight control signals, the aircraft can be temporarily controlled in mechanical mode.

• • • •

Pitch control is achieved through the horizontal stabilizer by using the manual trim wheel. Lateral control is accomplished using the rudder pedals. Both controls require hydraulic power. A red MAN PITCH TRIM ONLY warning appears on the PFD.

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NAVIGATION FLIGHT MANAGEMENT AND GUIDANCE SYSTEM (FMGS) The FMGS provides automatic navigation and flight guidance, map displays, autothrust management and thrust limits, and performance optimization. On the ground, a flight plan is entered into the FMGS along with certain required performance criteria to compute the optimum flight profile from departure to arrival. The system provides automatic aircraft guidance and computes the current and predicated progress along the flight plan.

Flight Management and Guidance Computers (FMGC)

The A-320 has two FMGCs; each has three functions: • Flight management (FM): This function computes the aircraft position, provided map display, selects and auto tunes navigation radios, and calculates performance. • Flight guidance (FG): this function provides commands to the autopilots, flight directors, and for autothrust. • Flight augmentation: This function provides rudder and yaw damping inputs, flight envelope and speed computations and windshear protection.

Multifunction Control and Display Units (MCDU)

The MCDU is the interface unit between the pilot and the FMGC. The MCDU supports insertion of the navigation and performance information, and also provides an interface with ACARS and AIDS (aircraft integrated data system) for maintenance purposes.

Flight Control Unit (FCU)

Located on the glareshield, this unit provided the capability of short term interface between the pilot and the FMGC. It is used to modify flight parameters and enable the selection of operation modes for the autopilot, flight directors and autothrust. Confirmation of engaged modes is through the FMA display on the PFDs.

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A-320 STUDY NOTES

Thrust Levers

The thrust levers activate the flight directors during takeoff and goaround by their selected position. The selection of FLEX or TOGA positions on takeoff will activate the flight directors and the SRS system; they will also trigger the updating of the FMGC position at the beginning of the take off roll. The go-around mode of the flight directors is triggered when the thrust levers go to the TOGA position during this maneuver.

FMGS Operation

Each FMGC is linked to its’ own MCDU, radio management panel (RMP) and electronic flight instrument system (EFIS). There are three possible modes of operation: dual, independent and single. The normal mode of operation is the dual mode, where each FMGC makes its own computations and exchange of data via the cross-talk bus. One FMGC is always the master and one the slave. • If one autopilot is selected on, the related FMGC is the master • I f two autopilots are selected on, FMGC 1 is the master • If neither is selected on: o FMGC 1 is the is the master for autothrust operation when both FDs are turned off o FMGC 1 is the master whenever the Captain’s FD is switched on. If a significant discrepancy occurs between FMGCs, they will go into independent operation with no cross-talk. Raw data must then be used to confirm navigation accuracy. Single mode is automatically activated if one FMGC fails; both MCDU are available and data is transferred to the operating system. In this mode, the NDs must be set to the same mode and range to access the map on either side. A message: SET OFFSIDE RNG/MODE will appear on the ND, if the offside mode and range are different to the onside or working side.

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FMGS Position Computation

Each FMGC receives position information from the three IRS units. They then compute the average position, called the MIX IRS position. In flight, the FMGC computes a radio position based on information sourced from, DME/DME, VOR/DME or ILS/DME. This radio positon is used to update the IRS MIX position. The system keeps track of the difference between the two positions or bias. If radio positioning is lost, the FMGC constantly computes the present position, taking into account the last know bias along with IRS MIX position until radio updating is available again.

Navigation Accuracy

If there difference between the FMGC computed position and the radioposition exceeds certain pre-determined limits, the navigation accuracy changes from HIGH ACCURACY to LOW ACCURACY; Raw data must then be used for navigation. The tolerances vary according to the phase of flight; for example: enroute accuracy should be less than approximately 3 nm, whereas for approach, this changes to approximately 0.4 nm.

Flight Guidance

The flight guidance part of the FMGS utilizes the autopilots, the flight directors, and the autothrust system to provide flight guidance. There are two basic modes of flight guidance: Selected and Managed. The selected modes are engaged by ‘pulling’ the appropriate knob on the FCU. The managed modes are armed or engaged by ‘pushing’ the appropriate knob. The only exception is that NAV mode engages automatically if the DIR TO is selected on the MCDU. Managed modes are used for vertical, lateral and speed profiles, as determined by the FMGS. These modes are considered ‘long term’ modes and are modified by action on the MCDU. Selected modes are used for vertical, lateral and speed profiles, as determined by the crew’s actions on the FCU. These modes are considered ‘short term’ commands and will override the managed modes.

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Windshear alerting system

The FACs provide windshear detection during take-off and approach. Detection is provided from liftoff till 1300’ when the flaps are at CONF 1 or greater, and from 1300’ to 50’ on approach. A red WINDSHEAR message appears on the PFDs and an aural WINDSHEAR, WIDSHEAR, WINDSHEAR is repeated. The FDs provide guidance for escape maneuvering. The autopilot if engaged will fly the escape maneuver if required.

ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) The EFIS has six display units (DUs); two are the primary flight displays (PFDs), two are the navigation displays (DUs) and two are the electronic centralized aircraft monitoring (ECAM). The ECAM DUs consist of an, engine/warning display (E/WD) and a system display (SD). Three identical display management computers (DMCs), process the inputs from the various aircraft sensors and computers to generate the display images. In the case of a DU failure, either automatic or manual switching allows the display to be transferred to another DU.

Primary flight displays

The PFD indications are: • Attitude indication • Flight director commands • Glide slope, localizer • Airspeed scale • Vertical speed scale • Barometric altitude scale • Radio Altitude • Heading/track information • FMGS modes on the flight mode annunciator (FMA) • Altimeter setting • ILS identifier • Marker beacons • TCAS and windshear recovery commands

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A-320 STUDY NOTES

Flight Directors (FDs)

The flight directors display the FMGS and FCU commands on the PFD. The basic mode is HDG/VS, which will engage if no other modes or armed or if certain reversions take place. A TRK/FPA mode is also available; it displays the aircraft horizontal track and vertical trajectory. If the FDs are on and the TRK/FPA pb is pressed, the Flight Path Director (FPD) appears in place of the cross-bars.

Navigational radios

Navaids are normally auto-tuned by the FMGC, but may be manually selected through the MCDU or radio management panels (RMPs).

VOR

Two VOR receivers are installed. They can be displayed on both NDs and the DDRMI. Since DME/DME is the preferred method of radio updating the FMGS, the VORs displayed on the NDs are not necessarily the navaids being used by the FMGS for radio updating.

ILS

Two ILS receivers are installed. Number 1 ILS is usually displayed on the CAPT PFD and FO ND and number 2 on the FO PFD and CAPT ND.

DME

Two, five channel DMEs are installed. Channels 1 and 2 are for normal VOR/DME tuning. Channels 3 and 4 are used for radio updating of the FMGC and channel 5 is for the ILS DME. VOR DME is displayed on the NDs and DDRMI. Once again, the DME stations that are being used for radio updating, are not necessarily those being used for display.

NDB Two NDBs are installed. Information is displayed on the NDs and DDRMI. NDBs are only autotuned if an NDB approach has been selected in the FMGS.

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Windshear alerting system

The FACs provide windshear detection during take-off and approach. Detection is provided from liftoff till 1300’ when the flaps are at CONF 1 or greater, and from 1300’ to 50’ on approach. A red WINDSHEAR message appears on the PFDs and an aural WINDSHEAR, WIDSHEAR, WINDSHEAR is repeated. The FDs provide guidance for escape maneuvering. The autopilot if engaged will fly the escape maneuver if required.

ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) The EFIS has six display units (DUs); two are the primary flight displays (PFDs), two are the navigation displays (DUs) and two are the electronic centralized aircraft monitoring (ECAM). The ECAM DUs consist of an, engine/warning display (E/WD) and a system display (SD). Three identical display management computers (DMCs), process the inputs from the various aircraft sensors and computers to generate the display images. In the case of a DU failure, either automatic or manual switching allows the display to be transferred to another DU.

Primary flight displays

The PFD indications are: • Attitude indication • Flight director commands • Glide slope, localizer • Airspeed scale • Vertical speed scale • Barometric altitude scale • Radio Altitude • Heading/track information • FMGS modes on the flight mode annunciator (FMA) • Altimeter setting • ILS identifier • Marker beacons • TCAS and windshear recovery commands

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A-320 STUDY NOTES

Navigation displays

The navigational displays show navigational information in one of five modes: • ARC (map mode) • Rose NAV (compass rose map mode) • Rose VOR (compass rose VOR mode) • Rose ILS (compass rose ILS) • Plan

Air data/inertial reference system (ADIRS)

Three identical air data/inertial reference units (ADIRU) are installed. Each ADIRS combines an air data reference (ADR) system with a laser gyro inertial reference system (IRS) in a single unit. Failure of one system will not affect the other. The ADR provides airspeed, Mach number, barometric altitude, angle of attack, temperature and overspeed warnings. The IRS provides attitude reference, flight path vector, heading , track, acceleration, deceleration, groundspeed and aircraft position. The IRS essentially provides position input to the FMGCs for navigation computation.

ADIRS control panel

The ADIRS control panel is located on the overhead panel. The present position (PPOS) must be entered in during the alignment process before use; this can be done through the MCDU or on the ADIRS control panel. Normal ‘full’ alignment takes 10 minutes, however a quick align can be performed by selecting the ADIRU 1, 2, and 3 OFF for less than five seconds and then back to ON, assuming of course that the system had been previously aligned. This facility is useful on short ground intervals. After a quick alignment, the system is ready to navigate after 3 minutes. GROUND PROXIMITY WARNING SYSTEM (GPWS) The GPWS monitors the flight path for dangerous or potentially hazardous conditions. The system provides both aural and visual alerts if the flight path enters certain predetermined conditions.

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A-320 STUDY NOTES There are seven basic modes of operation:

Mode 1 – Excessive Descent Rate

In this mode, the GPWS monitors for excessive descent rate, irrespective of configuration and will initially give a repeated voice warning of SINK RATE, along with the amber GND/PROX/G/S inhibit light switch. If the descent rate continues, a voice alert of WHOOP, WHOOP, PULL UP, sounds in conjunction with the red master warning lights illuminating and a red PULL UP message on the PFDs.

Mode 2 – Excessive terrain closure

This mode monitors for excessive terrain closure rate with gear and flaps not in the landing configuration (mode 2A), or with flaps in the landing configuration (mode 2B) Mode 2A provides a voice warning of TERRAIN, TERRAIN and the GND/PROX/G/S inhibit light illuminates. If the closure rate continues, the aural warning changes to WOOP, WOOP, PULL UP in conjunction with the red master warning lights illuminating and a red PULL UP message on the PFDs. Mode 2B provides a repeated TERRAIN, TERRAIN the GND/PROX/G/S inhibit light illuminates. If the closure rate continues below 700’ RA and the landing gear is not down, the aural warning changes to WOOP, WOOP, PULL UP in conjunction with the red master warning lights illuminating and a red PULL UP message on the PFDs.

Mode 3 – Altitude loss after takeoff or go-around

If the aircraft descends after takeoff or go-around, when it is below 700’ RA, a repeated DON’T SINK alert occurs along with the amber GND/PROX/G/S light illuminating.

Mode 4 – Unsafe terrain clearance

This mode monitors for unsafe terrain clearance with the gear not down (mode 4A) or the flaps not in landing configuration (mode 4B). Mode 4A provides a repeated voice alert, TOO LOW GEAR, or TOO LOW TERRAIN, depending on the speed and altitude. Mode 4B provides a repeated voice alert, TOO LOW FLAPS or a TOO LOW TERRAIN, depending on the speed and altitude. Sub-modes 4A and 4B can be inhibited by the crew for approaches in nonnormal configurations.

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A-320 STUDY NOTES

Mode 5 - Deviation below glide slope

If the aircraft deviates more than 1.3 dots below the glide slope, a voice alert of GLIDE SLOPE is provided; initially the warning is a low volume, increasing in intensity as the deviation increases. This mode can be inhibited for intentional deviations by pressing the GPWS G/S pb on the overhead panel. All other GPWS modes can preempt this mode as appropriate.

Mode 6 – Altitude advisories

This mode provides voice callouts or radio altitude at 2500, 500 (only when an ILS is tuned and the aircraft is not within 2 dots of the glide slope), 50, 30, and 10 feet. This function may be customized by the airline.

Mode 7 - Windshear

This mode monitors flight conditions for excessive downdrafts or tailwinds. If an excessive condition is detected, there is a voice alert WINDSHEAR, WINDSHEAR, and red WINDSHEAR message appears on the PFD. When the windshear warning is active, all other GPWS modes are inhibited until the condition ceases or the escape maneuver is initiated. The FMGS provides windshear guidance by means of the normal TOGA, pitch and roll modes.

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A-320 STUDY NOTES ELECTRONIC CENTRALIZED AIRCRAFT MONITORING (ECAM) The ECAM presents aircraft engine and system data on two of the six identical CRTs on the forward instrument panel. The two central displays are for the ECAM. The upper display is the engine/warning display (E/WD) and the lower, the system display (SD).

Engine/warning display

• The E/WD is divided into four section:

• The primary engine instruments and fuel quantity indications • Flap/slat position information • Warning and caution information • Memo messages concerning aircraft system status

System display

The system display has 12 different system pages that can be displayed either automatically by flight phase or system degradation, or manually by the pilots.

E/WD switching

The E/WD has priority over the SD, so if the upper display fails, the E/WD is automatically transferred to the lower display. If this occurs, the crew can manually select a system screen by pressing and holding the required key on the ECAM control panel (this leads to, hot fingers!) The SD can also be transferred to the Captain’s or First Officer’s navigational display (ND) by using the ECAM/ND XFER switching facility.

ECAM colour code

The system uses the following colour coding: • Red – requires immediate action • Amber – requires awareness but not immediate action • Green – normal long-term operation • White – titles and remarks

• Blue – actions to be carried out • Magenta – special messages (i.e. TO INHIBIT AND LDG INHIBIT)

If a system parameter requires monitoring by the crew, the ECAM will automatically call up the relevant page, and the affected parameter will pulse green to direct the pilots’ attention.

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ECAM warning and caution classification

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ECAM memo display

The E/WD memo section (lower left display) lists systems that are being temporally used (e.g. LDG LT). The memos are displayed in green, amber or magenta. Durong the takeoff and landing phases; TO INHIBIT and LDG INHIBIT are displayed in magenta to alert the crew that most warnings are suppressed below 1500’ for TO and 750’ for landing. This is provided to avoid distraction of the crew. However, the following warnings are not inhibited: ENGINE FIRE APU FIRE ENG OIL LO PR L + R ELEV FAULT A/P OFF CONFIG FWC 1 + 2 FAULT (amber caution with no aural warning) A (T.O) memo appears 2 minutes after the second engine start or with one engine running, when the CONFIG TEST pb is pressed. The memo disappears when FLEX or TOGA power is set. The landing memo appears below 1500’ RA if the gear is down or below 800’ if the gear is up.

T.O. CONFIG warnings/cautions

The following warnings or cautions will appear after the T.O CONFIG pb is pressed, and the aircraft is not properly configured for takeoff: • SLATS/FLAPS NOT IN TO RANGE (red) • PITCH TRIM NOT IN TO RANGE (red) • SPEED BRAKES NOT RETRACTED (red)

• SIDESTICK FAULT (BY TAKE OVER) (red) • HOT BRAKES (amber) • DOOR NOT CLOSED (amber)

The following occur only when takeoff power is applied: • PARK BRAKE ON (red)

• FLEX TEMP NOT SET (amber) (not displayed if thrust levers are in

TOGA detent).

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ECAM E/WD and SD sequence The following items occur when the ECAM detects a failure: • E/WD shows the warning or caution message • MASTER WARN or MASTER CAUT lights illuminate (except for level

1 cautions) • Audio warning is triggered (except for level 1 cautions) • SD presents the affected system page • CLR key illuminates on ECAM control panel • The lower left portion of the E/WD (memo section) is replaced with the primary or independent failure information, including the required action steps. The lower right side of the ECAM shows memos and secondary failure information. • A warning light may appear on the affected systems control panel, depending on the failure.

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ECAM SD SYSTEMS SYNOPTIC PAGES

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ECAM SD SYSTEMS SYNOPTIC PAGES

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ICE AND RAIN The ice and rain system allows the aircraft to be operated in heavy rain and icing conditions. Hot bleed air is used for wing and engine anti-icing and, electrical heating is used for the windshields, sensors, pitot probes, static ports and drain masts. The windshields also have windshield wipers and rain repellent.

Wing anti-ice

The three outboard slats are anti-iced with hot air form the pneumatic system. The wing anti-ice pb on the ANTI ICE panel controls the opening and closing of the pneumatic valves in each wing. On the ground, the valves will open for testing purposes but will close again after 30 seconds. The valves close automatically if a leak is detected, after landing and if there is an electrical failure.

Engine anti-ice

The engine anti-ice valves are electrically controlled and pneumatically operated. Each engine nacelle is anti-iced by an independent air bleed from the high pressure compressor. The engine anti-ice valves are operated by their respective switches on the ANTI ICE panel. The engines must be running for the system to work, and in the event of an electrical failure, the valves will open and remain open until electrical power is restored. When either anti-ice valve is opened the following occurs: • • •

Maximum N1 is limited Continuous ignition is applied Minimum idle RPM is increased to provide adequate bleed air pressure

Window heat

Electrical heating is used for anti-icing each windshield and for demisting the cockpit side windows. Two window heat computers (WHC); one on each side, automatically regulate the temperature and provide overheat protection. Window heat automatically comes on when: • •

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A-320 STUDY NOTES Low heat is provided on the ground and normal power in the air. The windows have a constant heat output with no changeover.

Probe Heat

The pitot probes, static ports, angle of attack sensors, total air temperature (TAT) are all electrically heated. Three independent probe heat computers (PHC) automatically control and monitor the: • Captain probes • F/O probes • STBY probes The probes are heated automatically when: • At least one engine is running or the aircraft is in flight • Manually before engine start, if the PROBE/WINDOW HEAT pb is pressed. On the ground the TAT probes are not heated, and the pitot heating operates in low heat until the aircraft is airborne.

Drain masts

The drain masts are electrically heated whenever there is electrical power on the aircraft. On the ground they operate at low heat and normal levels in flight.

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ENGINES The A-320 has two CFM 56-5A engines. Each engine is equipped with a full authority digital engine control (FADEC) system that provided full engine management. The FMGC calculates N1 values for all operating conditions and computes the maximum N1 limits in the case of FADEC failure. The engine interface unit (EIU) transmits to the FADECs the data it uses for engine management.

FADEC

The FADEC performs the following functions: • Fuel metering • Engine limits • Engine parameter transmission and monitoring indications • Automatic and manual starting sequences • Thrust reversers • Fuel circulation • Detection, isolation and recording of failures • Self cooling

for

cockpit

The FADECs have two channel redundancy, with one channel active and one in standby. They are powered by their own magnetic alternator when the N2 increases above 10% or by the aircraft electrical system when N2 decreases below 15%.

Start sequence During an automatic start sequence, the ENG page is displayed on the SD

when the ENG MODE selector is positioned to IGN/START, and the pack flow control valves close. When the ENG MASTER switch is selected to ON, the FADEC will control the start in the following sequence: • • • • • •

LP fuel valve opens Start valve opens If used, APU speed increases Ignition starts at 16% N2 (in flight, immediately) HP valve opens at 22% N2 (in flight, 15%) Start valve closes at 50% N2 and: o Ignition off (on the ground only) o Pack valves will re-open 30 seconds after start o APU speed returns to normal TRAINING PURPOSES ONLY - 123 -

A-320 STUDY NOTES Note: If the ENG MASTER switch has not been selected ON within 30 seconds of selecting the ENG MODE selector to the IGN/ START, the pack valves will re-open. On the ground with the N2 below 50%, the FADEC monitors engine starting parameters and if necessary, aborts the start. The FADEC will then automatically crank the engine to clear out fuel vapors. The FADEC will abort a start for the following: • Hot start • Hung start • Engine stall • No light up The flight crew can interrupt the start sequence at any time by turning off the MASTER switch. The manual start sequence is partially under the control of the FADEC; it will open the start valve when the ENG MAN START switch is selected ON, and will open the HP fuel valve when the ENG MASTER switch is selected ON as well activating both igniters. Additionally it will close the start valve at 50% N2, however it will not automatically abort starts.

Ignition system

The ignition system is provided for engine start on the ground and engine restart in flight. It also provides flame-out protection during icing conditions, turbulence etc. The system consists of two identical and independent circuits, channels A and B. During normal automatic starts, one igniter is used for each engine; the FADEC will alternate channels between starts. During in-flight or manual ground starts, two igniters are used on each engine. When the ENG MODE selector is in NORM and the engines are running, continuous ignition is automatically provided during the following: • • • • •

FLEX or TOGA thrust is selected on the ground TOGA is selected in flight ANG ANTI ICE switch is ON Engine surge or stall occurs in flight FLAPS lever position is other than 0 in flight

Continuous ignition can be selected at anytime by manually by selecting the ENG MODE selector to IGN/START.

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A-320 STUDY NOTES

Thrust levers

The thrust levers can only be moved manually. There are five detented positions: reverse idle (REV IDLE), forward idle (FWD IDLE), climb (CL), FLX/MCT and TOGA. The thrust lever position is transmitted to the FADEC which then computes the thrust rating limit and required N1. If the thrust lever is in detent the FADEC sets the limit according to the position; if the levers are between detents, the FADEC sets the higher limit.

Autothrust

The autothrust A/THR system, receiving inputs from the FMGC, controls the thrust dependent on the speed, altitude and configuration. The autothrust system does not move the thrust levers, instead if the autothrust is armed, the levers are moved manually to obtain the required thrust. If the autothrust is engaged, and the thrust lever is in a detent, then the system will control the thrust or speed depending on the autopilot/FD inputs. Autothrust is armed on the ground by: • Setting the thrust levers to either FLEX or TOGA with at least one FD on. (will not arm with both FDs OFF) • Pushing the A/THR pb on the FCU when the engines are not running, however the autothrust will disconnect after start. Autothrust is armed in flight by: • Pressing the A/THR pb on the FCU when the thrust levers are not in the engagement range (see note below) • Setting the thrust levers above the climb detent with two engines running • Setting the thrust levers above the MCT detent with one engine running Autothrust is engaged in flight by: • Setting the thrust levers in the engagement range when A/THR is armed • Pressing the A/THR pb on the FCU when the thrust levers are in the engagement range • Activation of alpha floor, regardless of the thrust lever position or arming status.

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A-320 STUDY NOTES Note: The thrust levers are in the engagement range when: • Both thrust levers are above IDLE but not above CL detent when both engines are running • The thrust lever of the operating engine is above IDLE, but not above MCT detent, when only one engine is running Autothrust can be disengaged by: • Pressing the disconnect switches on the thrust levers • Moving both thrust levers to the IDLE detent • Pressing the A/THR pb on the FCU, when the autothrust is engaged • Loss of arming signals • Setting both thrust levers above MCT detent when both FDs and autopilots are off and if below 100’ RA.

Warning

The autothrust system does not need to be armed or engaged for the alpha floor function to engage, however if the autothrust disconnect switch is pressed and held for more than 15 seconds, the system is disconnected for the remainder of the flight; this would then preclude alpha floor protection.

Notes: • The correct method of autothrust engagement in flight is to manually move the thrust levers until the white donuts line up with the actual N1, and then press the instinctive disconnect pb on the thrust levers. • After an alpha floor engagement, the thrust is locked in TOGA (TOGA LK) Thrust control is regained by pressing the instinctive disconnect push-buttons on the thrust levers or by selecting idle thrust.

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