Ramp & Transit ATR 72

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RAMP & TRANSIT

ATR 72-100/200

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Introduction:

The ATR 72 is a twin turboprop high wing and short-haul pressurized regional aircraft built by the French-Italian aircraft manufacturer ATR, designed to carry 64 to 74 passengers. It is built in cooperation by EADS(France) and Alenia (Italy).

It was developed from the ATR 42 in order to increase the seating capacity from 48 to 74 by stretching the fuselage by 4.5 m (14 ft 9 in ), increasing the wingspan, adding more powerful engines and increasing fuel capacity by approximately 10 percent.

The ATR 72 was announced in 1986, and made its maiden flight on October 27, 1988.

Exactly one year after that, on October 27,1989 Finnair became the first company to put the airplane into service.

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Dimensions

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Cabin Layout

Several version are available to combine passenger transportation with freight. The cabin layout may range form 64 to 74 passengers.

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Cross Section

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Cockpit

The cockpit can accommodate two pilots and one observer.

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Pilots Seats

The two pilots seats are stowed by sliding laterally against the flight deck side . In flight position, these seats have to move to the center line of the flight deck and then forward. They are also reclined.

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Performance

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Doors

The aircraft passengers door and service door are located at the rear side of the fuselage. The main cargo door is located at left forward side of the fuselage.

• Passengers door – left rear side

• Main Cargo Door -Left forward side

• Service door – right rear side

A door located on after LH side of passengers compartment is provided for access to aircraft.

The entry door is an outward opening, non plug type door with a net of 72cm (28.5”) wide without hand-rail and 1.75m (68.8”) high. The mechanism is essentially composed of two handles, a lifting cam and locking shoot bolts placed on the rear part of the door.

Attached to the integrated stair structure is a folding hand-rail which, by means of a link to the fuselage structure automatically erects when the door is opened.

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Passenger Door

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Cargo Door

The cargo door is an outward opening, non plug type door hinged at its upper edge giving a net clear opening of 1.30 m (51”) wide by 1.57 m (62”) high.

The door is actuated by an electrical actuator. A hold-open strut maintains the cargo in the open position. It also protects the door from wind gusts. The door is opened through the cargo door control panel which is located at the bottom right side of the cargo door. There is also an inside control panel which allow the door to be operated from inside compartment.

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Service Door

The service door is an outward opening, non plug type door with a net opening 69cm (27”) wide and 1.27 m(50”) high. Opened position is forward . Door operation can be performed manually from inside or outside of the airplane.

Door opening from outside:

The door can be opened from outside by rotating the outer handle 90 degrees clockwise from horizontal to vertical position. The door has then to be rotated on hinges by pulling on outer handle assisting it until the handle is next to hold open hook.

Pull the terminal located on the door hinges side push the door and release the terminal so that the handle engages the hold-open hook which holds the door locked in fully open position against the fuselage outer surface.

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Emergency Exit Door

The two emergency exit doors type lll of passenger compartment are hinged on the bottom by (two spigot type hinge) and opens inwards rotating on hinges and falling down.

Five stops are installed on door panel structure, two per side (fwd and aft) and one upper side while a simple locking mechanism serves to protude or retract a shoot bolt upper side of the door lock and unlock the door in the fuselage opening.

The locking mechanism may be operated either from aircraft inside or outside. The outer handle has also the rule of vent door to prevent the cabin pressurization to an unsafe level if the door outer skin periphery is depressed by the fuselage skin surrounding the door cutout when the door is closed. A window is provided in the center section of the door to allow the passengers to look outside.

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Instruments

General Description

The indicating and recording system is divides in two main parts: Indicating and Recording. The indicating system provides:

• Systems monitoring through the CCAS (Centralized Crew Alerting System).

• Indication and control of the system by means of different panels.

• and time display with electrical clocks.

The recording system enables various aircraft parameters, as well as voices in the cockpit, to be recorded.

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Recording System Description

The recording system is composed of :

• A Flight Data acquisition Unit (FDAU).

• A Digital Flight Data Recorder (DFDR).

• A Flight Data Entry Panel (FDEP).

• And an Accelerometer.

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Instruments and Control Panel

The instruments and controls panel, are grouped into 5 main groups:

• Instrument panels in front of the pilots;

• Centre pedestal between the two pilots;

• Consoles on the left and right hand sides of the cockpit;

• Overhead panel above the pilots;

• And miscellaneous panels which do not belong to the groups mentioned above.

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Clock

Time is indicated by two quartz clocks.

The captain's clock is located on the 3 VU panel and the F/O's clock on the 5 VU panel. Both clocks are identical.

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MFC Multi-Function Computer

Many functions are controlled by the Multi-Functional Computers (MFC's). Each MFC is composed of two modules A and B.

Their main purposes are:

• To control and authorize operationof aircraft systems1

• To monitor aircraft systems

• To manage systems failures and flight envelope anomalies and to command triggering of associated warnings.

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Centralized Crew Alerting System CCAS

A centralized crew alerting system (CCAS) continuously monitors all aircraft systems in order to provide the followings functions:

• Alert the crew of the existence of system malfunction or aircraft hazardous configuration with a clear indication of the

urgency of the situation.

• Identify the malfunction or the situation without ambiguity.

• Direct the appropriate corrective action without confusion.

Logic functions are performed by the MFC's B modules which aquire and process system failures and flight envelope protection signals and generate aural and visual warnings.

Two kinds of logic are possible.

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CCAS LOGIC

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CCAS Description

The CCAS is composed of:

• Two Multi-Functional Computers (MFCs)

• A Crew Alerting Panel (CAP).

• Two Master Warning and Master Caution Lights

• An EMER/AUDIO CANCEL switch

• A T.O CONFIG. TEST pushbutton

• A Left Maintenance Panel

The CCAS receives all Warning & Caution signals from various aircraft systems.

The MFCs send an audio warning to the Captain's and F/O's loudspeakers and activate the centralized alarms on the CAP.

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Status and Indication

Status and indications are integrated in the push-buttons (PB). Push-buttons positions and illuminated indications are based on general concept with the light out condition for normal continuous operation according to the basic rule.

With few exceptions the, the light illuminates to indicate a failure or an abnormal condition. Whenever possible, the failure alert is integrated in the push-button which has to be operated for corrective action.

Some push-buttons such as (ACW) are painted in amber to help crew to find them in case of smoke. (Fluorescent Painting).

PUSH-BUTTON BASIC FUNCTION

IN - Depressed

OUT - Released OFF- MAN- ALTN - SHUTON – AUTO – NORMAL

COLOR INDICATION

No light illuminated except flow bar Normal Basic Operation

BLUE Temporarily required system in normal operation

GREEN Back up or alternate system selected

WHITE Selection other than normal basic operation

AMBER Caution Indivation

RED Warning Indication

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MFC Overhead Panel Description

The MFC panel is shown in light test.

During normal operation, when selected OFF, the fault light extinguishes. Four illuminated push-buttons are installed on the overhead panel.

Each of them is dedicated to one module and:

• enables module power supply (push-button pressed)

• indicates if the module is not supplied (OFF legend on)

• indicates if the module is faulty (FAULT legend on)

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MFC Overhead Panel Location

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MFC Maintenance Panel Description

The right side maintenance panel includes a readout display for failures of systems linked to the MFC system. Recording of these failures is performed by the MDC module 1A.

Bite loaded magnetic indicator:

It indicates that at least one failure has been recorded by the MFCs. System Selector Switch:

It is normally placed in “NORM-FLT” position (in all other positions, the “MAINT PANEL” indicator light shows amber on the CAP.

During MFC maintenance memory reading, this selector switch enable various systems to be selected, in order to consult the failures which have affected the system involved. The bite advisory display indicates through illuminated “F” lights, the binary code of the failure recorded. The combination of illumination of these four lights enable up to 14 failures per system coded: The code/failure definitions are given in the concerned Job Instruction Card (JIC) or in the Flight Crew Operating Manual (FCOM). Note that failure code reading is only possible with engines shutdown.

PTA/ERS push-button: When a system is selected, PTA push-button (Push to Advance) enables recorded failures to be run on the failure display. An “FFFF” code indicates the end of list of failures. When ERS position is selected, this push-button serves to erase the system memory, if it is depressed for more than 5 seconds. An “F F” code is then displayed during erasing. But if this push-button is depressed for less then 2 seconds , an ARINC test is performed (Cross Tallk Test). An “F F” code is the when successful.

Magnetic indicator test push-button serves to check operation of the bite loaded magnetic indicator, when pressed for more than 3 seconds the magnetic indicator is activated. De-activation requires system selector on ERS position and PTA /ERS push-button pressed for more than 5 seconds. Connector outlet enables the optional MTS (Maintenance Test Set) system to be connected. The maintenance team can extract detailed maintenance data for the nine last flights with the TMT(Transportable Memory Terminal) system.

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ATR Structure

The ATR structure is composed of:

• Fuselage

• Wings

• Stabilizers

• Nacelles

The structure is also composed of :

• Windows

• Doors

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Fuselage

The fuselage of ATR 72 is a semi-monocoque type, designed according to fail safe/damage tolerant criteria. For structural and productions reasons, it is manufactured in structural sections as follows:

Section 11 : Fuselage nose section

Section 13 : Fuselage and FWD center section Section 15 : Fuselage center section

Section 16 : Fuselage Rear Center Section Section 18 : Fuselage Tail Section

Structure Details:

Semi-monocoque structure consist of frames and panels. The fuselage frame are built in 7075-T6 bent or rolled sheet with a z-profile (except in flight compartment). The frames are shear tired to the skin for load introduction and

damage tolerance of the panels (in the lower and side part of the fuselage). Stringers are tied to the frames by stringer clips. The floor panels capability are:

On cargo compartment, cabin entrance and corridor – 400kg/m sq (88lb/ft sq). On cabin under seats – 200 kg/m sq (44 lb/ft sq).

On entry passenger compartment, galley,lavatories and after cargo – 400kg/m sq (44lb/ft sq).

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Composite Materials

Here's shown composite structural components of the ATR 72.

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Wings

The ATR is equipped with a high wing composed of three main elements:

Rectangular center wing box, two trapezoidal outer wing box with a dihedral of 2.5° and a central wing box. The surface area of the wing is 61m² (656.4 sq. ft) and its span is 27m (88 ft 7 inches).

The following secondary structures are attached to the main elements: Wing tips Leading edges Trailing edges Ailerons Spoilers Fairings

Center wing box:

The center wing box structure is made of light alloy and includes: Front spar

Rear spar

13 sheet metal ribs 14 machined ribs single- piece lower skin Four element upper skin

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Stabilizers:

The aircraft is provided with T – type stabilizers bolted to the fuselage tail section.

The T – type design has been chosen because of its good performance specially when used in conjunction with turbo-prop engines. The main components are:

• horizontal stabilizer

• elevator

• vertical estabilizar

• rudder

• horizontal to vertical stabilizer fairings

Elevator horns and rudder horn accommodate part of anti-ice system.

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Electrical Systems

The electrical systems consists of three separate subsystems:

• DC (Direct Current)

• AC ( Alternating Current), constant frequency

• ACW ( Alternating Current Wild) variable frequency

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Electrical

The electrical power generation is provided by the following sources:

• Main and emergency batteries

• Two engine-driven DC starter / generators

• Two AC wild frequency generators

• Two external power units AC / DC.

In addition two static inverters supplied by DC system provides constant frequency AC power.

The ACW electrical system can also supply DC electrical system through a transformer rectifier unit (TRU). The electrical distribuition is ensured by busses which feed equipments.

Two separed networks (left and right) run individually and can be connected in case of generation failure due to the (BTC) Bus Tie Contactor.

With the battery switch off and no GPU connected, these buses are powered:

HOT EMER BAT BUS

+

-HOT MAIN BAT BUS

+

-Hot Main Battery Hot Emergency Battery HOT EMER BAT BUS

+

-HOT EMER BAT BUS HOT EMER BAT BUS

+

-HOT MAIN BAT BUS HOT MAIN BAT BUS

+

-+

-Hot Main Battery Hot Emergency Battery

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Generation

Electrical power necessary aboard the aircraft is given by two engine-driven DC starter generators and two AC wild frequency (ACW) three-phase generators driven by the reduction gearbox of the engine-propeller.

In addition two static inverters, supplied by DC main buses, give AC constant frequency (AC) power and two batteries, give power in flight, to the critical load, when all main DC sources are lost. Main battery is also used for engine starting.

The aircraft has two external power receptacles to allow an electrical supply, on ground, using DC and AC ground power units. The aircraft electrical network comprises:

• 28 VDC from starter generators

• 115/200 VAC WF (341-488 Hz) three-phase from ACW generators;

• 115 VAC CF (400 Hz) single-phase from static inverters;

• 26 VAC CF (400 Hz ) single-phase from static inverters;

• 24 VDC from batteries.

Each DC starter generator and ACW generator is associated with a Generator Control Unit (GCU), connecting it to its related channel and providing protection and fault detection. Two Bus Power Control Units (BPCU), one connected to DC GCU's and the other one connected to the ACW GCU's control main buses-tie, load shedding and external power.

The AC distribution is given by the following buses:

• 115 and 26 VAC inverter buses 1 and 2 - 400 Hz

• 115 and 26 VAC standby (STBY) buses - 400 Hz

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AC / DC Feeders

DC and ACW generation are provided by engine driven generators.

AC and DC feeders run in the wing leading edge and under the ceiling roof to the electrical racks.

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Static Inverter

Constant frequency is provided by two inverters which convert DC into AC.

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Inverters Location

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Electrical Racks

The electrical components are located in the electrical racks, except for batteries which are located on the floor of avionics racks. AC and DC electrical controls are located on the overhead panel.

Circuit breakers are located on the overhead panel and behind the first officer.

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Control Panels

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Interfaces

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Electrical Subsystems

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Hotel Mode

In Hotel mode, only DC and AC constant frequency power supply is available. ACW power supply is not available since ACW. Generator is linked to the propeller Reduction Gearbox. The hotel mode allows the aircraft to be free from a Ground Power Unit. The main electrical network is supplied by DC GEN 2.

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DC Hotel Mode Operation

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DC Cooper Cables

DC copper cables are divided into two aluminum-alloy cables to save weight.

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Weak Battery

If the main battery is weak, the ground power unit cannot be connected to the A/C electrical network, thus no external power will be available.

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DC Distribution

The aircraft DC distribution consists of:

• Two main busses ( DC Bus 1 and DC Bus 2 )

• a Hot Main Bat Bus

• a Hot Emer Bat Bus

• a DC Emer Bus

• a DC Ess Bus

• a DC STBY BUS

• 2 Utility Buses (Utilily Bus 1 and Utilily Bus 2)

Additionally:

• A DC SVCE BUS provides supply for aircraft services

• A GND-HDLG BUS provides supply for ground handling operations

DC EMER BUS , DC ESS BUS and DC STBY BUS provides supply equipment required to fly. This equipment remains supplied even after dual DC generator loss, by batteries.

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Safety and Precautions

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Batteries

The emergency battery avoids power transient on critical equipment during engine start and supplies emergency power after

main battery has been discharged.

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Multi Function Computer

Multi Function Computer (MFC) 1 and 2 provide controls for battery's own operation and transfer contactors controls.

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When any of the following doors are opened, the Ground Handling Bus is energized through the Battery Transfer Relay:

• Cargo Door Access Panel

• Refueling Panel

• Main Cabin Door

• Auxiliary Hydraulic Pump

HOT EMER BAT BUS

+

-Hot Main Battery Hot Emergency Battery HOT EMER BAT BUS

+

-HOT EMER BAT BUS HOT EMER BAT BUS

+

-HOT MAIN BAT BUS HOT MAIN BAT BUS

+

-+

-Hot Main Battery Hot Emergency Battery GND HDLG BUS GND HDLG BUS

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DC Service Bus

If DC service Bus push-button is depressed on the flight attendant panel, then the DC Service Bus is powered through the Service Bus Transfer Relay.

ON

GND SVCE BUS PASS LAV CAPT LAV CARGO

EMER LT CALLS SMOKE GALLEY LAV LAT PASS DIM UPR PASS ENTR AFT CARGO READ LT VENT AC DC ON GND SVCE BUS PASS LAV CAPT LAV CARGO

EMER LT CALLS SMOKE GALLEY LAV LAT PASS DIM UPR PASS ENTR AFT CARGO READ LT VENT AC DC EXTER NAL POWER EGHR GND HDLG BUS HOT EMER BAT BUS + -HOT MAIN BAT BUS + -Hot Main Battery Hot Emergency Battery BTR EXTER NAL POWER EXTER NAL POWER EXTER NAL POWER EGHR GND HDLG BUS GND HDLG BUS HOT EMER BAT BUS + -HOT MAIN BAT BUS + -Hot Main Battery Hot Emergency Battery HOT EMER BAT BUS + -HOT MAIN BAT BUS + -HOT EMER BAT BUS HOT EMER BAT BUS + -+ -HOT MAIN BAT BUS HOT MAIN BAT BUS + -+ -Hot Main Battery Hot Emergency Battery BTR

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DC Electrical Generation System Indicating

The DC measuring system gives voltage and current indication by a voltmeter and an ammeter, under a control selector. The two DC instruments and the selector are installed on LH maintenance panel 101 VU.

The selector controls the voltage and current if it is set to each single DC generating source as follows:

• Current indications are given, by two hall effect current sensors through a separate output.

• Another hall effect current sensor ( for the external power channel) gives current indications relevant to the DC external power.

• Voltage indications, for each channel, are given by the related POR installed at the load end of each generator feeder.

• In addition an ammeter, with a center zero, is installed on MAIN ELEC PWR panel. It shows the charge / discharge current, in true

amperes of battery selected by BAT SEL switch on the same panel. This measurement occurs through the related charge/ discharge shunt 22 PA (MAIN BAT) and 39 PA (EMER BAT).

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DC Electrical System Indicating Panel

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Batteries Voltage

Batteries voltage is measured at HOT MAIN BAT BUS and HOT EMER BAT BUS level respectively for main and emergency batteries. Voltage is displayed on the DCV indicator located on the LH maintenance panel. Load is measured from shunts, and indicated on the DC AMP indicator located on the main electrical panel power control panel. When selected on the proper position, the “ELEC IND” selector allows the relevant parameters to be displayed. Selecting the “BAT SEL” switch allows the batteries load to be checked ( charged or discharge).

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GPU Connect

To connect the GPU to the network, external power must be available.

This is indicated by the “AVAIL” light on the overhead panel and by the “DC CNTD” and “DC PWR NOT USED” white lights located on the external power control panel.

When all the conditions are met, the BAT switch and the EXT PWR push-button have to be selected ON. Then the 11 PG contactor closes, supplying DC BUS 1 and DC BUS 2 through the BTC 16 PU.

As the contactor closes, the DC PWR NOT USED light extinguishes on the external power control panel.

DC PWR NOT USED also extinguishes when the SERVICE BUS and the GROUND HANDLING BUS are directly supplied by external power.

One of the conditions to connect external power is “aircraft on ground”. So, when the aircraft is on jacks, this condition is lost. To restore this condition, the WEIGHT ON WHEELS switch, located on the RH maintenance panel has to be selected to FLT. to connect the GPU to the network, external power must be available.

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DC / AC GROUND CONNECTING

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DC External Power Receptacle Location

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Lights

This chapter deals with normal and emergency aircraft lighting for the cockpit and the cabin, as well as exterior lights and ground service lighting.

For aircraft lighting, different system are installed:

• controlled from the cockpit

- cockpit lighting - cabin signs lighting - emergency lighting - exterior lighting

• controlled from the cabin attendant panel

- cabin lighting - emergency lighting

- rear cargo compartment lighting

• controlled from the exterior panel lights

- Fwd cargo compartment

- Main and nose landing gear wheels - Fwd and Aft avionics compartment - refueling panel

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Cabin and Cargo Lighting

Normal cabin lights consists of two rows of fluorescent lights . In addition, separate lighting is installed in the rear cargo compartment. These lights are operated from the cabin attendant panel. The forward cargo bay is lit from a switch located outside, on the cargo door operating panel. A switch located RH of the entrance door, provides 2 min illumination of two emergency lights and the cockpit entrance light. Another switch located at the cockpit entrance panel (40 VU) provides the same functions .

As soon as passenger door is open, Ground Handling Bus is available (entrance light power supply).

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Exterior Lighting

Exterior Light include:

• A: Navigations lights

• B: Taxi / take off lights

• C: Landing lights

• D: Wing scan lights

• E: Beacon anti collision lights

• F: Strobe lights

• G: Logo lights

• H: Emergency lights

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Light Panels Location

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Compartment Lights

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Instrument Light Panel

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Exterior Lights Panel Description

The position of the aircraft is indicated by navigation lights installed on the wing tips and tail cone and controlled by a switch. The system consists of:

• A red navigation light on the LH wing tip

• A green navigation light on the RH wing tip

• A white rear navigation light on the tail cone

• One NAV/EXT /LT switch installed on the overhead panel EXT LT section 27 VU in the flight compartment.

Each navigation light is equipped with one 28 VDC 50W lamp . The lights are supplied by 28 VDC SVCE bus or by 28 VDC bus 1. Note: Ice evidence probe light illuminates when navigation light switch is set to NAV position.

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Exterior Lights Panel

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Navigation Lights and Ice Evidence Probe Light

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Location

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The emergency lighting system is independent from the main lighting system. It consists of: Exterior Lighting

• 3 flood lights, located on the fuselage for ground lighting.

• 1 stair integrated ground floodlight.

Interior Lighting

• 4 dome lights located in the ceiling in the passenger compartment axis.

• 2 EXIT signs located at each end of the passenger compartment and showing the location of emergency exits.

• 4 EXIT signs located above or next to the emergency exit.

• 4 EXIT signs located next to the emergency exits in the vicinity of the floor.

• 1 EXIT path located at the floor level , on the left side , below the armrests enabling aisle lighting.

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Communications

The communications systems comprises:

• Radio communications between aircraft and ground stations.

• Passenger address system (PA).

• Audio integrating system, which manages all audio signals.

• Cockpit voice recorder which records all messages to and from the cockpit

The communication system also provides:

• An emergency locator transmitter (ELT)

• Navigation source identifiers reception

• Aural alerts generated by the CCAS, GPWS, and TCAS through the cockpit loudspeakers only.

Note that sound level for CCAS, GPWS, and TCAS alerts is preset and is not adjustable.

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The passenger address system is used to broadcast announcements and audio signals (chime) via the cabin and galley loudspeakers. The announcements can be made from the cockpit or from the cabin attendant station.

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RCAU Remote Audio Control Unit

The heart of all electronic cockpit communications is the Remote Control Audio Unit (RCAU). The RCAU contains an electronic processor for each cockpit crew member.

Each processor integrates audio controls as selected on the respective Audio Control Panel at each cockpit crew member position.

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CVR Cockpit Voice Recorder

The Cockpit Voice Recorder (CVR) enables recording of conversations received and transmitted by the crew members, announcements

made to passengers as well as aural warnings. Recording is made in a tape with a recording capacity of 30 minutes. The CVR panel provides test and erase facilities.

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Cockpit location

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Interfaces

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Audio Control Panel

The audio integrating system manages all the audio signals between aircraft and ground communications and navigation stations. It allows intercommunication between crew members(pilots and cabin attendant) and with ground mechanics. It also enables selection transmissions channels and provides adjustment audio level reception.

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RCAU Failure

In case of RCAU failure, two AUDIO SEL push-buttons are provided on both sides of the front panel to directly connect a VHF channel to the respective pilot. By pressing its AUDIO SEL push-button, the captain will be connected to VHF1 and the first officer to VHF 2. On the affected side, passenger address, interphone and other VHF (or HF if installed ) can not be use any longer.

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CVR Recording After Engine Shutdown

Once the aircraft is on ground, the cockpit voice recorder continues to record for ten minutes after both engine shutdown if the battery selector is set to on and the external power is not available. This delay is controlled by the MFC internal logic.

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CVR Erase

The CVR erase procedure can only be performed when the aircraft is on ground and the parking brake is set to the parking position.

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Safety and Precautions

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VHF System Description

Here is the VHF system block diagram . Two systems, VHF 1 and VHF 2 , are provided. Each system has its own transceiver to provide communications om 720 channels (or 760 depending on version) from 118.00 to 136.975 MHz with 25 KHz spacing.

Each system is controlled on the pedestal by a VHF control unit with dual frequency selection.

Two antennas are installed on the fuselage, one for each VHF system. Each VHF system is connected to the audio integrating and to the FDAU, and via the Audio Integrating to the Cockpit Voice Recorder .

In case if audio control panel loss , two AUDIO SEL push-buttons select one VHF to each side.

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VHF Operation

When the VHF1 control unit is ON, it displays the active and preset frequencies stored in memory when the equipment was last turned off. The inner knob of the frequency selector is used to change the 2 right digits by 50 Khz increments or by 25 Khz increments for the first two increments when the direction of rotation is reversed .When a transmission is performed, the TX amber annunciator appears on the VHF 1 control unit.

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Transceiver

The VHF 1 and VHF 2 transceivers are located in the electronics rack (82 VU and 83 VU).

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VHF Control Unit

Both control units are located on the centre pedestal, one on the left side, the other on the right side.

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VHF Antenna

Two VHF antenna are mounted above of the aircraft fuselage.

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Static Discharges

The aircraft is equipped with a total of 25 static dischargers:

• Four on each aileron and one on each aileron horn

• Four on the rudder and one on the top of the vertical stabilizer

• Three on each elevator and one on each elevator horn

• Two on the tail cone.

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ELT Emergency Locator Transmitter

An automatic Emergency Locator Transmitter (ELT), must be attached to the aircraft in such a manner that, in the event of crash, the probability of the ELT transmitting a detectable signal is maximized.

The ELT system comprises:

• a transmitter in the pressurized ceiling, close to the toilet door.

• An antenna in the fairing ahead of the stabilizer fin.

• A Remote Control Unit in the flight compartment

The ELT is able to transmit a distress signal on 121.5 MHz and 243.0 MHz frequencies and a geographical position sign on 406.026 MHz frequency. The system includes its own battery.

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ELT Auto Test

An auto test of the Emergency Locator Transmitter can be performed through the Remote Control Unit in the flight compartment. This test must not be performed in manual mode. So check that the MAN/AUTO selector is in “AUTO” position and guarded. The AUTO TST RST switch is used of undue alert (reset) or test the emergency beacon . Press Auto TST RST switch.

The X MIT ALERT caution light illuminates amber for 2 seconds and then extinguishes.

By setting VHF 1 frequency to 121.5 MHz , check that no distress signal is heard in the headset. In case of test failure, the X MIT ALERT caution light will flash amber.

When aircraft is on ground (and electrically supplied) , if the emergency beacon is triggered after 30 seconds, note that the mechanical horn will be triggered too.

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Center Pedestal

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Navigation System Description

The aircraft navigation system provide the crew with information required for a flight in compliance with safety requirements. This data can be divided in 8 groups:

• Flight environment data

• Attitude and direction

• Electronic flight information system

• Navigation

• Air traffic

• Aircraft internal aids

• Landing

• Flight management computer

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________________________________________________________________________________________________________________________________ ATR72-100/200 Navigation Rev: 1 Issue: 15/12/2010 For Training Purpose Only Page: 113

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EGPWS

The purpose of the Enhanced Ground Proximity Warning System (EGPWS) is to prevent CFIT (Controlled Flight Into Terrain) accidents.

The followingadditional modes are incorporated:

• Terrain Clearance Floor (TCF)

• Terrain Awareness and Display (TAD)

• Mode 6 update: Altitude and call-outs

• Mode 4 update: Envelope and Aural Alerts To operate, EGPWS system requires data supplied from ADC1, ILS2, Radio Altimeter, AHRS1,

GNSS, WX Radar, flaps position transmitter and gear lever position switch. The system provides SGU1 and 2 with terrain data to display on EHSI.

The Ground Proximity Warning System (GPWS) provides alerts in case of dangerous flight path conditions which would result in inadvertent ground contact if maintained. To achieve this function the GPWS generates visual (warning or caution light) or aural (synthetic voice)

warnings or cautions by processing signals supplied by other aircraft systems (air data system, radio altimeter, flaps, landing gear…).

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