ATR 72 - Flight Manual

AIRCRAFT GENERAL1.00.00 P1 CONTENTS 001 SEP 03 AA R 1.00.00 CONTENTS 1.00.10 GENERAL 1.00.20 COCKPIT 1.00.25 COCKPIT DOOR SECURITY SYSTEM (if installed) 30.1 DESCRIPTION 30.2 CONTROLS 30.3 ELECTRICAL SUPPLY/MFC LOGIC/SYSTEM MONITORING 1.00.40 LIGHTING 40.1 DESCRIPTION 40.2 CONTROLS 40.3 ELECTRICAL SUPPLY/MFC LOGIC 1.00.50 WATER AND WASTE SYSTEM 50.1 DESCRIPTION 50.2 ELECTRICAL SUPPLY 1.00.20 AIRCRAFT GENERAL P 2A COCKPIT 001 SEP 05 AA SEAT POSITION SIGHT GAUGE Seat position sight gauge may be used for proper setting seat height and fore/aft position. It assures to the crew a correct view of instrument panels as well as runway environment, especially when flying low visibility instrument approaches. This indicator is composed of three colored balls. Center ball is red and is horizontally shifted compared with the other two white balls. For proper seat position, respective white ball is obscured by the red one. . the failure alert is integrated in the PB which has to be operated for corrective action. PB positions and illuminated indications are based on a general concept with the light out" condition for normal continuous operation according to the basic rule.20 AIRCRAFT GENERAL P 04 001 SEP 04 COCKPIT PHILOSOPHY AA COCKPIT PHILOSOPHY Status and failure indications are integrated in the pushbuttons (PB).. MAN.1. ALTN. PB POSITION BASIC FUCTION IN (DEPRESSED) OUT (RELEASED) ON. SHUT COLOR INDICATION No light illuminated except flow bars BLUE GREEN WHITE AMBER RED Normal basic operation Temporarily required system in normal operation Back up or alternate system selected Selection other than normal basic operation Caution indication Warning indication . the light illuminates to indicate a failure or an abnormal condition.. NORM OFF.00. R Some PB (such as ACW. AUTO. With few exceptions.) are painted in amber to help crew to find them in case of smoke R (fluorescent painting). Whenever possible. 00. the flight crew can authorize entry by unlocking the door which remains closed until it is pulled open. R Mod : 5377 + (5434 or 8330 or 8333) . thereby allowing or denying entry into the cockpit. In normal conditions.Locking and unlocking the electromagnetic locks. following an access request.a Cockpit Door Control Unit located in the forward avionics bay . The Cockpit Door Locking System Switch enables the Cockpit Door Locking System to be switched ON or OFF at the beginning or at the end of the flight and to facilitate maintenance tasks and ground operations. To open the door. upon flight crew action .AIRCRAFT GENERAL 1. the cabin crew may unlock the door by pressing the Emergency push-button of the door call panel installed on cargo compartment side. When the flight crew does not respond to a request for entry. The toggle switch enables the flight crew to lock or unlock the cockpit door. The right door panel can be removed from the cockpit in case of the door becomes jammed by using the quick release hinges that are only accessible from inside the cockpit. The Cockpit Door Control Unit is the system controller in charge of : .a buzzer . intrusion resistant and fully compliant with the rapid decompression requirements.1 Cockpit door description The door has an electromagnetic locking system controlled by the pilots.a Cockpit Door Locking System Switch located on the 121VU The door call panel enables the cabin crew to request access to the cockpit. Upon receiving an entry request.2 Cockpit Door Locking System Description The cockpit door locking system (CDLS) provides a means of electrically locking and unlocking the cockpit door. it remains locked. The system is mainly composed of : .a toggle switch.Alert annunciation to indicate failure of electromagnetic locks .a Door Call panel located on the cargo compartment side on the right door panel . The door is bulletproof. 25.25 P1 COCKPIT DOOR SECURITY SYSTEM 200 SEP 04 AA 25. located on the center pedestal's Cockpit Door Control panel (811VU) . the locking system has to be unlocked. when the door is closed.Activating the access request buzzer and turning on the Door Call panel LED's The buzzer sounds in the cockpit for a minimum of 2 seconds (duration = duration of pushbutton activation + 2 seconds) to indicate that a routine access request has been made or sounds continuously if an emergency access procedure has been initiated. There are two different access request modes : a Routine" access request type and an Emergency" access request. 00. The buzzer will sounds continuously in the cockpit. R Mod : 5377 + (5434 or 8330 or 8333) .3 Door Call Panel The Door Call Panel is used by the cabin crew to request pilots to open the door.AIRCRAFT GENERAL 1. 1 2 DENIED LIGHT illuminates : When the flight crew has denied access. 102/202/212/212A version 101/201/211 version 4 OPEN Green LED illuminates : The door has been unlocked either by a flight crew action or automatically (during 10 seconds) when no flight crew action has been performed during the delay (30 seconds). It is used to initiate the emergency opening of the door when the flight crew does not respond. following an emergency access request. the OPEN green LED on the Door Call Panel will flash. but no action has yet been taken by the flight crew. CALL push-button switch : It is used to sound the buzzer in the cockpit for at least 2 seconds 3 4 EMER push-button switch : It is protected by a rotating plate to prevent inadvertent activation. The door panel can then be pulled open OPEN Green LED flashes : An emergency request to enter the cockpit has been made. the buzzer will sound continuously in the cockpit.25 P2 COCKPIT DOOR SECURITY SYSTEM 200 SEP 04 AA 25. the light remains illuminated for 3 minutes. During this delay any action on the call panel will be inhibited. The door remains locked. If no crew action is taken. the door will unlock at the end of the 30 seconds delay.25 P3 COCKPIT DOOR SECURITY SYSTEM 200 SEP 04 AA 25. CLOSE Position : The door is locked. electrical supply) 3 Buzzer Buzzer sounds : For at least 2 seconds after the cabin crew has request an access through the CALL push-button on the Door Call Panel. the door is locked. or continuously when the cabin crew has started an emergency access procedure R Mod : 5377 + (5434 or 8330 or 8333) . move the handle located on the cockpit side and pull the left panel aft. and emergency access is possible for the cabin crew DENY position : Once the button has been moved to this position. FAULT : This light comes on when a system failure has been identified (Example : electromagnetic lock. Note -If the DENY position has not been used by the pilot for at least 3 minutes. the cabin crew is able to request either the routine or the emergency access. located on the central pedestal 1 Toggle Switch OPEN Position : This position is used to enable the cabin crew member to open the cockpit door.4 Cockpit Door Control Panel The cockpit door opening is controlled by a toggle switch.AIRCRAFT GENERAL 1.00. the cockpit door is automatically unlocked. the buzzer and the Door Call Panel are inhibited for 3 minutes. -The OPEN position overrides and resets any previous selection -In case of an electrical supply failure. To open the left door panel. or not locked OPEN light flashes : The cabin crew has started an emergency access procedure. If there is no reaction from the flight crew. control unit. The switch must be pulled and maintained in the open position until the right door panel is pulled open. EMERGENCY access. 2 Fault/open Indication OPEN light ON : The door is not closed. 00.5 COCKPIT DOOR LOCKING SYSTEM ON/OFF CONTROL The Cockpit Door Locking System ON/OFF Control Switch is located behind the First Officer on the 121VU panel.AIRCRAFT GENERAL 1.6 ELECTRICAL SUPPLY / SYSTEM MONITORING ELECTRICAL SUPPLY EQUIPMENT Cockpit Door Locking System Cockpit Door Fault light R Mod : 5377 + (5434 or 8330 or 8333) DC BUS SUPPLY (C/B) DC BUS 2 ESS BUS 2 (on central pedestrial cockpit door control panel) . ON/OFF Control Switch ON : The cockpit door locking system is operative OFF : The cockpit door locking system is deactivated. If the system is switched off and the aircraft power is available.25 P4 COCKPIT DOOR SECURITY SYSTEM 200 SEP 04 AA 25. 25. The door is unlocked. the door right panel can be opened from the cargo compartment side by pulling it. move the handle located on the cockpit side and pull the left panel. To open the door left panel. the FAULT light on the pedestal Cockpit Door Control panel comes on. 8") high. SERVICE DOOR The service door is an outward opening.00.25).5") wide (without hand-rail(s)) and 1. A latch operated by a knob on the cabin side and a safety key from the cargo side is provided. The mechanism is essentially composed of two handles. smoke doors separate the forward cargo compartment from the cockpit. Mod : 4019 Model : 102Ć202Ć212Ć212 A .30). R Except when a Cockpit Door Security System is installed (see in this case 1. by means of a link to the fuselage structure automatically erects when the door is opened. Opened position is forward. refer to 1. Door operation can be performed manually from inside or outside of the airplane (refer to 1.30).07.27 m (50") high. In case of emergency it can be forced opened in either direction.75 m (68. non plug type door with a net opening of 72 cm (28.00. Four safety pins are provided (two on each side) in order to remove the doors in case of emergency.AIRCRAFT GENERAL 1. Attached to the integrated stair structure is a folding hand-rail which. Note : Remove the pin after closing and install it before opening. INTERNAL DOOR A forward opening hinged door separates the forward cargo compartment and the passenger compartment.30 P1 DOORS AA 160 SEP 03 30.07. non plug type door with a net opening of 69 cm (27") wide and 1.1 DESCRIPTION LOCATION ENTRY DOOR The entry door is an outward opening. a lifting cam and locking shoot bolts placed on the rear part of the door (for door operating. JUN 97 HDL R Model : 102-202-212-212 A . JUN 97 R Model : 102-202-212-212 A . JUN 97 R Model : 102-202-212-212 A . . JUN 97 R Model : 102-202-212-212 A . JUN 97 R or 212 A . JUN 97 R Model : 102-202-212-212 A . . EXIT.RH DOME light with the possibility to switch it off .07). Mod : 5040 Model : 102-202-212-212 A .40 P3 300 LIGHTING JUL 99 AA EMERGENCY LIGHTING Emergency evacuation path marking near the floor is a photoluminescent system. Two sources are available : .DC STBY BUS via a voltage divider. In case of flight with DC STBY BUS only.6V integral batteries charged from the DC STBY BUS with a 10 mn capacity. illuminating when associated door is locked. light will be supplied by DC STBY BUS. If this source fails. the batteries will be utilized automatically.LH three lights located below the glareshield . CEILING and EXTERIOR emergency lights are supplied with 6V DC.AIRCRAFT GENERAL 1. One light is provided in the toilet.00. the cockpit lighting is restricted to : . . Note : Emergency flash lights are provided (see 1.overhead panel light illuminating the pedestal. In case of system activation. . . . . READING LT sw When depressed passenger reading lights are operational. 3 4 5 6 7 8 9 10 Mod : 3795 .40 P8 LIGHTING 050 JUL 99 AA HOSTESS PANEL 1 Galley sw (when installed) Monitor galley lighting. CARGO sw Monitor cargo lighting. 2 LAV sw Monitor lavatory lighting. DIM LT sw Associated light. LAT PASS sw Monitor lateral passengers lighting. VENT sw (when installed) When depressed. ENTRANCE sw Monitor entrance lighting.AIRCRAFT GENERAL 1. passenger ventilation fan operates. CEILING LT SW Monitor the passengers ceiling lighting. are dimmed. EMER LT sw Controls emergency exit light and evacuation path marking causing emergency lights to illuminate (overriding crew switching). when selected. A diffuser switches ON when lavatory latch is closed.00. . SEP 03 R (on lateral panel) . . . 080 JUN 97 Mod : 3973 or 4371 or 4457 . AIR 1.03. : 3973 or 4371 or 4457 .10 P 3/4 GENERAL 080 JUN 97 AA ROFA–01–03–10–003–A080AA SCHEMATIC Mod. It supplies under pressure air for air conditioning. the valve is spring-loaded closed regardless of electrical power supply.Bleed duct LEAK . ISOLATION Downstream of the junction of the LP and HP ducting. with associated control. air is admitted into the duct by a pneumatically operated. At low engine speed when pressure from LP stage is insufficient. Transfer of air is achieved through a pressure regulating valve which is electrically controlled. electrically controlled butterfly bleed valve which acts as a shut off valve.Wing and engine de-icing. air source is automatically switched to the high compressor stage HP. air is therefore bled from HP stage only. The pneumatic system consists of all the systems designed to supply air to the various aircraft systems. pressurization and ice protection system.I. monitoring and indicating components. 7/8) Compressed air is bled from the engine compressors at the LP or HP stages. .Actuation of associated ENG FIRE handle .).AIR 1. AIR BLEED .Engine failure at T0 (UPTRIM signal) .1 DESCRIPTION 070 DEC 97 (See schematic p.03. A protection against overheat due to possible leakage around the hot air ducts is provided. Air is bled from the HP compressor stage. In the absence of air pressure. : 4457 . Transfer of air is achieved by means of a pneumatically operated and electrically controlled butterfly valve. (This may occur on ground and during descent at F. It includes a single solenoïd which locks the valve closed when deenergized. without any recirculation into the engine.  when the HP valve is open.Air conditioning and pressurization The system is designed to :  select the compressor stage from which air is bled. zones or engines. The bleed valve automatically closes in the following cases : .  regulate air pressure in order to avoid excessive pressures Air is generally bled from the low compressor stage (LP). the bleed valves opening is inhibited.20 P1 PNEUMATIC SYSTEM 20. air is directly bled from the LP stage through LP bleed air check valves. Note : During a starting sequence. (HP valve) which remains closed in absence of electrical supply :  when the HP valve is closed. depending on the pressure and/or temperature existing at these stages. Mod. the HP air pressure is admitted into the LP pneumatic ducting and closes the check valves .PROPELLER BRAKE selected ON (for left bleed valve only).Bleed duct OVHT . R . Note : In case of OVHT. They operate at 274_C (525_F) and are controlled by the MFC. For any temperature higher than a preset value : 124_C (255_F) applied to a part of the sensing element. LEAK DETECTION SYSTEM A continuous monitoring system is installed in order to detect overheat due to duct leakage and to protect the structure and components in the vicinity of hot air ducts: .Air conditioning pack area. R . Each sensing element is permanently subjected to the temperature of the compartment it protects. solenoid controlled.AIR 1. it is always open except when both engines are running and propeller brake is disengaged.Upper and lower fuselage floor. one for the RH and one for the LH air duct systems. the associated bleed system may be recovered after a cooling time. The valve is closed with solenoid deenergized. the crossfeed valve is normally closed. The sensing elements comprise a control lead (nickel wire) embedded in an insulating material and are integrated in an inconel tube connected to aircraft ground. These switches. . Mod: 3037 + (3973 or 4371 or 4457) . near the HP compressors exit. This results in an alert signal processed in a control unit which triggers illumination of LEAK It. the associated pack valve. The sensing system includes two single loop detection assemblies. . Note : In case of LEAK.On the ground. In order to ensure rapid leak sensing.20 P2 PNEUMATIC SYSTEM 150 JUN 97 AA CROSSFEED The crossfeed valve installed on the crossfeed duct is designed to connect LH and RH air bleed systems. the resistance of the eutectic mixture rapidly decreases and the central lead is grounded.03. the crew must consider the associated bleed system as inoperative for the rest of the flight. HP valve and BLEED valve (and GRD X FEED valve if the left loop is affected) are automatically latched closed. pneumatic shut off valve.In flight.Wing leading edge and wing to fuselage fairing. After one second time delay. This is a spring loaded closed. which are duplicated for safety. a Kevlar envelope is installed around the major part of the high temperature ducts to collect and direct leaking air to the sensing elements. OVERHEAT CONTROL SYSTEM This system includes switches (thermal resistances) which are installed on the engines. ensure that the bleed valve and the bleed air shut-off valve are closed whenever any abnormal over temperature conditions occur. 03.AIR 1.20 P 2A PNEUMATIC SYSTEM 100 JUL 00 AA LEAK DETECTION SYSTEM (See page 2) The maximum sensing element temperature is 153°C (307°F) instead of 124°C (255°F). Mod : 4584 . OFF (pb released) associated HP and bleed valves are closed. Mod 4584 .AIR 1. Á OVHT light The light illuminates amber and the CCAS is activated when either of the respective bleed duct dual overheat sw operates (T > 274 ºC/525 ºF).03. The valves will open if pressure is available. OFF light illuminates white. ON (pb pressed in) associated HP and bleed valves solenoids are energized. Â LEAK Light The light comes on amber and the CCAS is activated when respective bleed leak detection system signal an alert (T loop > 153 ºC/307ºF).20 P3 PNEUMATIC SYSTEM 100 JUL 00 À ENG BLEED pbs Controls the associated HP valve and BLEED valve. FAULT The light illuminates amber and the CCAS is activated when the bleed valve position disagrees with the selected position. this especially occurs in case of leak or overheat. Ã X VALVE OPEN light The light illuminates amber when the GRD X FEED valve is open. . 100 JUL 00 153°C/307°F) Mod 4584 . . 060 JUL 00 R ROFA–01–03–20–007–A060AA T LOOP Mod : 3973 or 4371 or 4457 T LOOP . AIR 1. dried.30 P1 AIR CONDITIONING 200 SEP 03 AA 30.Pressure control and hence flow control.by two ground turbo fans through turbo shut off valves when : . : PW127 . the right pack supplies only the cabin air conditioned. Note : If one pack is inoperative. (see page 15/16). The left pack supplies the cabin and the cockpit air conditioned . The pack valve is pneumatically operated and electrically controlled.by ram air when IAS > 150 kt. Note : Incorrect position of a turbo fan shut off valve leads to closure of associated pack valve R Eng. IAS
150 kt and landing gear is retracted for less than 10 min.Pack shut off .1 DESCRIPTION See schematics p. the other one supplies both compartments through the mixing chamber. cooling of air is performed : . Hot air from the engines is admitted through pack valves and conditioned (cooled. Normal or high flow are available. The two packs are installed in the main landing gear fairings and operate automatically and independently. . 13/14 and p. Without air pressure and regardless of electrical command. compressed) into the packs. IAS
150 kt and landing gear is extended. This butterfly valve has two functions : . Note : Pack valves will be automatically closed in case of leak detection.03. the turbo fan starts on the opposite side and is running as long as IAS
150 kt and regardless of landing gear position. Note : In case of ENG OIL LOW PRESS. . The selection of the high flow mode increases the pack entrance pressure resulting in conditioning performance improvement. the pack valve is spring-load closed. It will also close without electrical supply.PW127F . 15/16) AIR PRODUCTION The air conditioning system is supplied by air processed through two packs which regulate air flow and temperature as required (see page 13/14). . JUN 97 R Model : 102-202-212-212 A . . SEP 04 R FLT COMPT CABIN . . 200 JUN 97 R Model : 102-202-212-212 A . 200 JUN 97 200 JUN 97 R Model : 102-202-212-212 A . 200 JUN 97 R Model : 102-202-212-212 A . . . . . . SEP 04 R Note : Only for models 101-201-211 Pressuriation AUTO MODE will be inhibited in flight in case on UNLOCKED DOOR alert on ground . . . . . . . . . . the GPS R (if installed) and from some sensors. Roll servo actuator. The computer receives data from the two Air Data computers (ADC). It generates commands to the flight control actuators and to the FD bars. Yaw and pitch servo actuators.1 DESCRIPTION (See schematic p 11/12) The aircraft is provided with an automatic flight control system. the radio-altimeter .AFCS 1.10 P1 001 GENERAL JUL 98 AA 10. control box and computer (cockpit and electronic rack). the two SGU.04. COMPONENT LAYOUT 1 2 3 AFCS advisory display. . Dual microprocessor architecture and digital servo-monitoring technique are used to provide an adequate safety level. the two Attitude R and Heading Reference Systems (AHRS). It achieves : • Autopilot function and/or yaw damper (AP and/or YD) • Flight director function (FD) • altitude alert Main components are : • one computer • one control panel • one advisory display unit (ADU) • three servo-actuators (one for each axis). . . . . . . . Mod.10 P9 GENERAL 070 JUN 97 AA 10.3 ELECTRICAL SUPPLY/MFC LOGIC ELECTRICAL SUPPLY DC BUS SUPPLY (C/B) AC BUS SUPPLY (C/B) DC EMER BUS (on overhead panel CMPTR) . : 4366 .Nil - Servo controls DC STBY BUS (on overhead panel SERVO) .AFCS 1.04.Nil - AP OFF lights + AP DISC circuit DC ESS BUS (on overhead panel WARN) .Nil - DC STBY BUS (on overhead panel ADU) .Nil - EQUIPMENT AP/FD computer + YD DISC circuit + control box + GUIDANCE" indication (*) ADU (*) if installed MFC LOGIC See chapter 1.01. . . . On the ADU. AP and/or YD is disengaged.AFCS 1.GA mode . AP/YD MONITORING RECOVERY When a monitored failure is detected. The AP white arrows extinghish. Action on the RESET pb or the quick disconnect pb clears the warnings and message.Action on the YD pb on control panel or an effort on pedals disengages the YD and AP. Action on the RESET pb or quick disconnect pb clears the warnings and message. . If initial failure condition still exists. PITCH MISTRIM" or AILERON MISTRIM" message is displayed on the ADU.04. or STBY or NORMAL pitch trim switch activation or effort on control column disengage the AP function without disengaging the YD function.LOC or BC modes Mod : 3608 .Loss of AP. Conditions which will inhibit all recovery attempts are : . R Note : If PITCH TRIM ASYM It illuminates on central panel. AUTOMATIC DISENGAGEMENT The warnings and messages are the same as those which occur in case of manual disengagement but AP OFF" light and AP" or AP/YD DISENGAGED" message are flashing. AP automatically disengages and cannot be reengaged. the RESET" pb illuminates amber and the AP/YD DISENGAGED" message is displayed in amber on the second line. If the pilot clears messages displayed on ADU (by using RESET pb) the FGC will attempt a monitor recovery". or GA mode activation.Action on the AP pb on the control panel. The AP/YD can be once again engaged. YD and AFCS controls panel. . Action on RESET" pb clears warnings and messages. the RESET pb illuminates amber and the AP DISENGAGED" message is displayed in amber on the second line.20 P2 AUTO PILOT/YAW DAMPER 050 JUL 01 AA MANUAL DISENGAGEMENT . The AP and YD white arrows extinguish. AP/YD is disengaged again. The AP OFF" It illuminates red and the cavalry charge" aural waning is generated. or quick disconnect pb on each control column. the AP OFF It illuminates red and the cavalry charge" aural warning is generated.Trim inoperative monitor failures . Note : If a failure occurs. The crew has to disengage AP and manually fly the aircraft. the PITCH TRIM FAIL". On the ADU.Any APP mode . . . . . . . . . . . . . . . . 3 ELECTRICAL SUPPLY/MFC LOGIC/SYSTEM MONITORING 10.2 CONTROLS 20.05.00 COMMUNICATIONS P1 030 CONTENTS SEP 02 AA 1.4 SCHEMATIC 1.05.00 CONTENTS 1.05.10 GENERAL 10.3 OPERATION R Mod : 3074 or 3113 or 3625 or 3832 or 5103 or 5146 or 8259 .20 TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM (TCAS) 20.05.1 DESCRIPTION 20.1.2 CONTROLS 10.1 DESCRIPTION 10. GENERAL . 0 to 29.05. HF COMMUNICATION SYSTEM The system provides the pliot access to 40 programmable channels.000 operating frequencies in the 2. In case of audio control panel loss.9999 MHz range.10 P2 GENERAL 030 JUL 00 AA VHF COMMUNICATION SYSTEM ROFA-01-05-10-02-020 Two systems are provided. Mod : (0032 or 1028 or 2267) + (4928 or 5007) .COMMUNICATIONS 1. two AUDIO SEL pbs allow to select one VHF on each side. plus a full 280.000 to 136. Each system has its own transceiver to provide communications on more than 2000 channels from 118. It is controlled by a HF control box through its transceiver. and is controlled by a VHF control box with dual channel selection.975 MHz with 8.33 KHz spacing. 000 to 136. six emergency channel frequencies and all 249 half-duplex ITU maritime radiotelephone channels are programmed. The system has the capability to store 99 user programmed preset channels.05. continuous wave (CW) and optional frequency modulation (FM) mode.999 MHz frequency range in 100 Hz steps. lower sideband (LSB) when allowed. In case of audio control panel loss. amplitude modulation equivalent (AME). Mod : 5285 . Communication is possible in both simplex and half duplex in upper sideband (USB). HF COMMUNICATION SYSTEM The HF system covers the 2. two AUDIO SEL pbs allow to select one VHF on each side.0 to 29.10 P2 GENERAL 050 SEP 02 AA ROFA-01-05-10-02-020 VHF COMMUNICATION SYSTEM Two systems are provided.975 MHz with 25 KHz spacing. Each system has its own transceiver to provide communications on 720 channels (or 760 depending on version) from 118. In addition.COMMUNICATIONS 1. and is controlled by a VHF control box with dual frequency selection. In case of ground crew call MECH pb illuminates blue on overhead panel and a door bell is generated by the MFC. PASSENGER ADDRESS SYSTEM The passenger address system allows the crew and the cabin attendant to make announcements to the passengers. A single chime (or three for an emer call) is generated in the cabin and the CAPT white light illuminates on the cabin attendant panels.To call the ground crew from the cockpit (see 1.The RCAU which allows the cockpit crew to make announcements to the passengers by selecting PA key on audio control panel.To interconnect all stations (ground crew stations + cockpit + cabin attendant) : D Depress INT transmissions key D Use PTT pbs .The cabin attendant call pb in the cockpit .10 p 9) D Depress the ATTND pb (overhead panel) for a normal call. . .10 P3 GENERAL 070 JUL 99 AA INTERPHONE SYSTEM . Passenger address system also generates single chime sound in the cabin.05.The cabin attendant call pb distributed in the cabin for the passengers .20) Mod : 5017 or 5018 or 8214 or 8215 .The cockpit voice recorder . or press the ATTND pb three times successively for an emergency call. . the CAPT It extinguishes. both visual and aural calls will be cancelled.The NO SMOKING/FASTEN SEAT BELTS controls .05.Cockpit crew interphone Using boom set or oxy mask : D Set the RAD/INT selector on "INT" position without use of PTT pb or.15) Generated by TCAS (when installed.To call cockpit from cabin attendant station D Depress INT pb (besides the hostess panel) for a normal call or press the EMER pb for an emergency call. refer to 1. D Set the PTT selector (control wheels) in the forward position regardless of the RAD/INT selector position .02) Generated by GPWS (refer to chapter 1. As soon as the cabin attendant selects "INT".05.Loudspeakers distributed in the cabin (one of them being installed in the toilet) DISTRIBUTION OF AURAL ALERTS Generated by CCAS (refer to chapter 1. both visual and aural calls will be cancelled.To call cabin attendant from the cockpit (see 1.The cabin attendant handset .10 p 9) D Depress MECH pb (overhead panel) A horn call is generated in the nose gear bay. Associated ATTND light illuminates on overhead panel and a door bell is generated by the MFC.COMMUNICATIONS 1.05. By pressing RESET. By pressing RESET. The passenger address system is connected to : . . . an inner knob is used for selection numbers right of the decimal point  Channel display The active channel is displayed on the first line. à XFR/MEM switch This is a three positions spring loaded toggle switch. Á Channel selector In normal use. controls the preset channel display : . . Mod : 0043 + (4928 or 5007) .an outer knob is used for selection of numbers left of the decimal point .05.10 P5 020 GENERAL JUL 00 ROFA-01-05-10-005-020 AA À ON/OFF and volume knob Energizes the control box and the associated VHF.MEM : successive actions cycle the six memory channel through the display.NEUTRAL .COMMUNICATIONS 1. .XFR : exchanges preset and active channel. SQ OFF position disables the receiver squelch circuit. Annunciators are displayed on both lines. The preset channel is displayed on the second line. 128. and first line can directly be tuned from channel selector.05. 129. Push the STO button a second time enters the preset channel. 128.) TX illuminates when the VHF is transmitting. second line displays dashes. Æ TEST button Is used to initiate the radio self-test diagnostic routine. Ç Annunciators Three types of messages can appear in this location.the MEM swith may be used to advance through the channel numbers. Mod : 0043 + (4928 or 5007) . É Light sensor Automatically controls the display brightness.400 .For 5 seconds.390 . Å ACT button Allows to change the active channel .When depressed. ACT flashes if the actual channel is not identical to the channel in the active channel display.COMMUNICATIONS 1. È Compare annunciator ACT signal illuminates when channels are being changed.g.When depressed. CAUTION: Untimely squelch triggering may accur on the following channels 124.990 . MEM illuminates when a preset channel is being displayed on the second line. RMT illuminate when the VHF is remotely tuned (by an FMS e. Returns to the initial configuration when depressed a second time.10 P6 GENERAL 020 JUL 00 AA Ä STO button Allows entering six channels in the memory .990 without disturb emission and reception. the upper window displays the channel number of available memory (CH1 to CH6).990 and 134. ALT Normal operating position. (7) .Photo cell Automatically controls the display brightness.Annunciators . Transponder replies with flight level information.Power and mode switch OFF ATC control box and transceiver are deenergized.1. (3) . causes the transponder to transmit IDENT" signal.Inner knob controls digits.ACT Compare annunciator ACT is displayed during code changes. FID Elementary Surveillance.TX is displayed when the ATC replies to an interrogation.10 COMMUNICATIONS P7 GENERAL 240 SEP 04 AA ATC CONTROL BOX (1) . Mod : 5487 .Outer knob selects digits. The stored code can be recalled by momentarily pressing the PRE button again. STBYATC system is under power. ON ATC Transponder Mode S replies to both Mode A and Mode C interrogations but without from ground or air.ATC Code and FID Display Display selected ATC code and Flight ID. . (4) . . ACT flashes when the actual reply code is not identical to the code shown in the active code display. (2) . FID : . Flight IDentification (FID) display and selection. (5) . but does not transmit replies. .IDENT button When depressed.Inner knob controls the two rigth-hand digits.Test button Press the TEST button to initiate the radio self test routine.RMT is displayed when the ATC is remotely tuned. (9) .05.Code select knobs ATC : .Outer knob controls the two left-hand digits. (6) .PRE button (Preset) Push and hold the PRE button while turning the code select knobs to select a preset code for storage. (8) . . NORM (pb depressed) RCAU functions normally. the volume of both loudspeakers is muted.10 P9 GENERAL AA 070 JUL 99 LOUDSPEAKERS VOLUME KNOBS Communication reception over cockpit loudspeakers is controlled by an individual knob for each of the two cockpit loudspeakers.COMMUNICATIONS 1. . CALLS PB(s) See 1.05. Volume is adjusted by affected loudspeaker volume control. Note : On the affected side PA. interphone and other VHF can not be used any longer. ALTN (pb released) affected crew station is connected directly to : VHF 1 if CAPT station is affected or VHF 2 if F/O station is affected. FAULT illuminates amber and the CCAS is activated when when an associated RCAU processing board failure or power loss is detected. Note : In case of aural alert : .during any transmission.normal volume is always available regardless of knobs position.05. AUDIO SEL PB(s) Controls functioning of associated RCAU processing board.10 p 3 Mod : 5017 or 5018 or 8214 or 8215 . The antenna is located in the fairing ahead of the stabilizer fin. This system includes its own battery. the mechanical horn is triggered too.Net work X MIT ALERT illuminates amber during 2 seconds. or to test the emergency beacon. . AUTO transmission is made automatically on 121.Failure X MIT ALERT It flashes during 15 seconds.10 P 10 GENERAL 050 JUL 98 AA EMERGENCY BEACON The transmitter is located in the ceiling of the cabin between the passengers entry door and the toilet door. MAN allows commanded operation (X MIT ALERT It illuminates amber).5 MHz. Two cases are possible for the test : . R Aircraft on ground (and electrically supplied). AUTO TEST RST is used in case of undue alert (resert).05. when the emergency beacon is triggered R after 30 seconds. Mod : 4080 . 243 MHz and 406 MHz when deceleration exceeds 5 g (X MIT ALERT It illuminates amber).COMMUNICATIONS 1. CAUTION : The test must not be performed in MAN mode. . The MODE field uses two alphanumeric characters to display the RF emission mode. The VOL or SQL levels will be displayed at the same time in the FREQ/CHAN field. MODE. selects built-in test (BIT). The OPR field will display FLT (fault) or WRN (warning) should a fault or warning occur in the system.10 P 11 GENERAL 050 SEP 02 AA HF CONTROL BOX (1) CURSOR .Maritime mode . PGM . The OPR field will momentarily display VOL (volume) or SQL (squelch) when either the VOL or SQL control settings are changed. allows scanning (reception) of user-programmed channels. MAN . or PWR) by rotating CURSOR control.05. allows selection of a discreet frequency and mode. selects external system control.Manual mode . TST .Test mode .External mode . EXT .COMMUNICATIONS 1.Scan mode . SCN . MAR . allows selection of built-in maritime channels.Standby mode. (2) OPR . SBY .The OPR field displays the system operating mode. (3) MODE . A segmented line display is positioned under field to be changed (OPR.The cursor control moves the cursor left or right to select the field to be changed. The emission mode is selected using the VALUE control. EMR . The system makes available the following RF emission modes.Channel mode . The available modes include the following : CHN . FREQ/CHAN. UV LV UD LD AM CW FM Mod : 5285 Upper sideband voice Lower sideband voice (where allowed) Upper sideband data Lower sideband data (where allowed) Amplitude modulation equivalent Continuous wave Frequency modulation (option) .Program mode .Emergency mode . used to program channels. allows selection of preprogrammed emergency channels. allows selection of user-programmed channels. . Rotating CHAN control sequentially changes (increments or decrements) the current channel number regardless of cursor position. (5) PWR . (6) VALUE .Disables the squelch circuit (11) SQL .The CHAN control provides selection of all preprogrammed (maritime. S Channel selection is accomplished using either the CHAN or VALUE control. Mod : 5285 .. Press FREQ/LD switch when EMR or MAR is displayed in OPR field to view emergency (receive and transmit) or maritime (receive) frequency. the ADF bearing pointer will be parked in 90_ position (ADF flag). or PWR) selected by the cursor. Press the FREQ/LD switch when CHN is displayed in OPR field to recycle the channel and frequency display.10 P 12 050 GENERAL SEP 02 AA (4) FREQ/CHAN S The frequency is displayed using all six digits and a decimal point.medium power All three bars . user-programmed) channels.05.COMMUNICATIONS 1. emergency. press the FREQ/LD switch when PGM is displayed in OPR field to load desired receive-transmit emission mode. MODE. Change the frequency using the VALUE control.The VALUE control changes (increments or decrements) the field (OPR. CAUTION : Collins ADF system is connected to HF9000 PTT to inhibit erratic bearing deviation during transmission. Should a permanent stuck-on PTT failure occur during more than 15 to 30 sec. For half-duplex operation. (9) CHAN . and channel data into non-volatile memory of the receiver-transmitter for the 99 user-programmable channels. (7) See (1) (8) FREQ/LD . press both the FREQ/LD switch and the microphone key to load transmit data.For simplex operation.low power Bottom two bars .high power Use the VALUE control to select the power level. Press FREQ/LD switch while keying microphone to view maritime transmit frequency. (10) DSBL .The power field is a 3-level bar indicator that displays the selected RF output power as follows : Bottom bar .Allows to reduce background noise when not receiving a signal. frequency. FREQ/CHAN. . . . 05.COMMUNICATIONS 1.10 P 16 GENERAL AA LEFT INTENTIONALLY BLANK 001 JUN 97 . . . . 05. The TCAS will not detect aircraft without transponders. The TCAS is a single system installation consisting of : . The RA displays a threat resolution in the form of a vertical maneuver if the potential conflict is projected to occur. and issue vertical resolution advisories on the pilot's TCAS vertical speed indicator (TCAS VSI) to avoid conflict. Outputs from the TCAS System are voice messages and visual displays on the TCAS VSI's for Resolution Advisories (TCAS RA's) and Traffic Advisories (TCAS TA's). .1 DESCRIPTION The TCAS is an on-board collision avoidance and traffic situation display system with computer processing to identify and display potential and predicted collision targets. R Mod : 3074 or 3625 or 8259 or 5146 or 3832 or 5103 . The TA is informative and indicates potential threats. TCAS will determine the threat using standardized algorithms.two high resolution bearing antennae (one top mounted and one bottom mounted) . Threat aircraft with mode A transponders will not provide altitude information . .1. therefore TCAS will not issue resolution advisories for these threats. From this. . .two overhead speakers for voice messages and associated wiring.20 COMMUNICATIONS P1 TCAS 050 SEP 02 AA 20. TCAS determines relative altitude.One TCAS processor. From the transponder replies.two mode S transponders.two modified TCAS VSIs each integrating trafffic advisory display and vertical speed information.one pylon mounted TCAS control box. range. and bearing of any aircraft equipped with a mode C or S transponder. AUTO : Normal operating mode of the TCAS. .If altitude reporting is off or no valid ModeS transponder is selected. TCAS will automatically go into the TA only mode when the TCAS equipped aircraft is below the RA descent altitude and in a climb inhibit configuration.Disables the RA mode of operation. automatic (AUTO). . Traffic Advisory mode or Resolution Advisory) are not operative. TA only .If appropriate. TCAS will be in standby (RA OFF on TCAS VSI). .May be selected but should be used only to prevent unnecessary resolution advisory when operating near closely spaced parallel runways or in the cases TCAS could command Climb maneuvers resulting in an unsafe situation for the aircraft (see limitations on 2. preparation.20 P2 TCAS 200 SEP 02 AA 20.01.05.The TCAS should be tested by pressing the TEST" button during cockpit preparation. Note : . . or traffic advisories only (TA ONLY) mode of operation. 2 R TCAS test function .05). STBY : TCAS system is under power.2 CONTROLS TCAS CONTROL BOX 1 TCAS rotary selector Enables TCAS to be set to standby (STBY).COMMUNICATIONS 1. but TCAS functions (intruder visualisation.The mode S ATC transponder will no function during some portion of the self-test sequence.Use of the self-test function in flight will inhibit TCAS operation for up to 20 seconds depending upon the number of targets being tacked. Mod : 0043 + (3074 or 3113 or 3625 or 8259 or 3832 or 5146 or 5103) . COMMUNICATIONS 1.. Threat : A target that has satisfied the threat detection logic and thus requires a resolution advisory. Mode S : Type of secondary surveillance radar (SSR) equipment which provides replies to mode A and Mode C interrogations and discrete address interrogations from the ground or air. Corrective resolution advisory : A resolution advisory that advises the pilot to deviate from current vertical speed. e.20 P3 TCAS 170 SEP 02 AA DEFINITIONS A B C D E F G H I R Advisory A message given to the pilot containing information relevant to collision avoidance. Preventive resolution advisory : A resolution advisory that advises the pilot to avoid certain deviations from the current vertical speed because certain vertical speed restrictions exist. Proximate traffic : Nearby aircraft within $850 ft and 6NM which are neither an RA nor a TA.g.05. The information contains no resolution information. Resolution advisory (RA) : Aural and visual information provided to the flight crew to avoid a potential collision. Intruder A target that has satisfied the TCAS threat detection logic and thus requires a traffic advisory. Traffic advisory (TA) : Information given to the pilot pertaining to the position of another aicraft in the immediate vicinity. Mod : (3074 or 3625) + 3832 . CLIMB when the aircraft is levelled. and assists the pilot in achieving visual acquisition of the threat aircraft. CLIMB.3 OPERATION The TCAS provides two levels of threat advisories : If the traffic gets between 20 and 48 seconds (depending on aircraft altitude) of projected Closest Point of Approach (CPA). This level provides a recommanded vertical maneuver using modified TCAS VSI's and voice messages to provide adequate vertical separation from the threat aircraft. of CPA.20 P4 TCAS 150 SEP 02 AA 20. or prevents initiation of a maneuver that would place the TCAS aircraft in jeopardy. DESCEND. The TCAS resolution advisories are annunciated by the following voice messages. it is then considered an intruder. "CLEAR OF CONFLICT" : (Range is increasing. This level calls attention to a developing collision threat using the traffic advisory display and the voice message. It permits mental and physical preparation for a possible maneuver to follow. CROSSING DESCEND. CROSSING DESCEND" : (Descend at the rate depicted by the green (fly to) arc on the TCAS VSI) safe separation will best be achieved by descending through the intruder's flight path. CLIMB. Mod : (3074 or 3113 or 3625) + 3832 . If the intruder gets between 15 and 35 seconds (depending on aircraft altitude).) "MONITOR VERTICAL SPEED. and separation is adequate. DESCEND. TRAFFIC TRAFFIC". CROSSING CLIMB. and an aural and visual traffic advisory is issued. "REDUCE DESCENT-REDUCE DESCENT" : (Reduce vertical speed to a value within the illuminated green arc).COMMUNICATIONS 1. it is considered a threat and an aural and visual resolution advisory is issued. MONITOR VERTICAL SPEED" : (Spoken only once if issued after a previous corrective advisory). Safe separation will best be achieved by climbing through the threat's flight path. Assure that vertical speed is out of the illuminated TCAS VSI red arc.).05. DESCEND" : (Descend at the rate depicted by the green (fly to) arc. REDUCE CLIMB" : (Reduce vertical speed to a value within the illuminated green arc). CLIMB" : (Climb at the rate depicted by the green (fly to) arc on the TCAS VSI). "CLIMB. CROSSING CLIMB" : (Climb at the rate depicted by the green (fly to) arc on the TCAS VSI). "DESCEND. "DESCEND. return to assigned clearance). "REDUCE CLIMB. as appropriate : A B C D E F G H R CLIMB. Received after a DESCENT" resolution advisory and indicates a reversal in sense is required to achieve safe vertical separation from a maneuvering threat aircraft. DESCEND-DESCEND NOW" : (descend at the rate depicted by the green (fly to) arc on the TCAS VSI).05. INCREASE CLIMB" : (climb at the rate depicted by the green (fly to) arc on the VSI). "INCREASE CLIMB. Received after DESCEND" advisory. The tone and inflexion must constate increased urgency. "DESCEND-DESCEND NOW. and indicates additional climb rate required to achieve safe vertical separation from a maneuvering threat aircraft. INCREASE DESCENT" : (descend at the rate depicted by the green (fly to) arc on the TCAS VSI).20 P5 TCAS 150 SEP 02 AA The following voice messages annunciate enhanced TCAS maneuvers when the initial TCAS RA does not provide sufficient vertical separation. Received after CLIMB" advisory. "CLIMB-CLIMB NOW.below 900 ft AGL when aircraft is descending R Mod : (3074 or 3113 or 3625) + 3832 . Received after a CLIMB" resolution advisory and indicates a reversal in sense is required to achieve safe vertical from a maneuvering threat aircraft.COMMUNICATIONS 1.below 1100 ft AGL when aircraft is climbing . A B C D INCREASE DESCENT. All TCAS aural alerts are inhibited : . and indicates additional descent rate required to achieve safe vertical separation from a maneuvering threat aircraft. CLIMB-CLIMB NOW" : (climb at the rate depicted by the green (fly to) arc on the TCAS VSI). below 1100 ft AGL during a climb . In this case.
DESCEND" RA's are inhibited : . In non altitude crossing encounters for which a CLIMB" RA is posted.below 1650 ft AGL during a climb .below 1000 ft AGL in approach during a descent
INCREASE DESCEND" RA's are inhibited .20 P6 180 TCAS SEP 02 AA TCAS OPERATING CHARACTERISTICS
NON ICING CONDITIONS of inhibition CONFIGURATION FLAPS 0 FLAPS 15 TO FLAPS 15 Approach FLAPS 30
RA CLIMB RA INCREASE CLIMB AUTHORIZED AUTHORIZED AUTHORIZED AUTHORIZED AUTHORIZED INHIBITED AUTHORIZED INHIBITED ICING CONDITIONS of inhibition CONFIGURATION FLAPS 0 Z < 20 000 ft Z > 20 000 ft FLAPS 15 TO FLAPS 15 Approach FLAPS 30 RA CLIMB RA INCREASE CLIMB AUTHORIZED INHIBITED AUTHORIZED AUTHORIZED INHIBITED INHIBITED INHIBITED INHIBITED INHIBITED INHIBITED
The increase climb" RA is inhibited for certain above conditions.05. the CLIMB" RA will remain posted for the duration of the encounter.below 900 ft AGL during a descent
There can be a case where the threat aircraft track on altitude information is lost during an RA.below 1450 ft AGLduring a descent
All RA's are inhibited : . the RA sense will be reversed and a DESCEND" RA will be posted. If the threat never crosses through. Since the increase climb" RA is inhibited. the threat may maneuver or accelerate toward own aircraft and cause a reduction in vertical separation despite the RA.COMMUNICATIONS 1. the RA will terminate without a CLEAR OF CONFLICT" annunciation. Mod : 5103 or 5146 or 8259 .below 1200 ft AGL during a climb at take off . As soon as the threat passes throught own aircraft's altitude. the climb RA remains posted. . . . JUN 97 R Mod : 1603 Model : 102-202-212-212 A . . . GND HDLG BUS The GND HDLG BUS supplies the DC loads required for airplane servicing on the ground even with BAT sw selected OFF. from HOT MAIN BAT BUS provided : D Cargo door operating panel door is open (micro switch). this bus may be recovered by selecting OVERRIDE pb.06. the cabin attendant controls the DC SVCE BUS supply from a sw located on the cabin attendant control panel. the DC power transfer is achieved by automatic opening and/or closure of electrical contactors according to the particular electrical conditions.ELECTRICAL SYSTEM 1. The master sw is the DC SVCE/UTLY BUS pb. * DC EMER BUS is supplied from the emergency battery or from the TRU.On ground D When EXT PWR is connected (p. . 21/22) * The BTC closes (BTC green flow bar illuminates). DC STBY BUS are supplied from the main battery. * The engine driven generator 2 supplies the DC BUS 2. the GND HDLG BUS is deenergized in flight. 19/20) * The engine driven generator 1 supplies the DC BUS 1. from DC SVCE BUS.DC BUS 1. if selected ON. or from the TRU. and on ground during airplane servicing operations. Only the cabin attendant pb has control. Since these loads are not required during flight. or D Refueling panel is open (micro switch). or D Entry door is open (micro switch). . D If both engine driven generators fail (p. if selected ON. . The GND HDLG BUS can be supplied : .In flight D Both engine driven generator operating (p.20 P4 DC POWER 220 JUL 99 AA R DC SVCE BUS The DC SVCE BUS supplies power in flight.EXT PWR. The supply of the BUS may be performed with batteries switched OFF. D If both engine driven generators fail and TRU is inoperative (p. * The entire electrical network is supplied by the remaining engine driven generator.When EXT PWR is not available. Mod : 1603 Model : 102-202-212-212 A . 15/16) Note : The electrical power transfer is achieved in the same way as in flight as long as EXT POWER is not connected. The DC SVCE BUS can be supplied by : . 26) With all switches in normal position.When EXT PWR is available. 25/26) * When DC STBY BUS reaches undervoltage (amber UNDV light comes ON). * The BTC is open. 15 to p. D If one engine driven generator fails (p. . TRANSFER (see schematics p. When selected on. 23/24) * DC ESS BUS. . . . . . . . . . ELECTRICAL SYSTEM 1.20 P 15/16 DC POWER Mod : 1603 220 JUN 97 Model : 102-202-212-212A .06. JUN 97 R Mod : 1603 Model : 102-202-212-212 A . JUN 97 R Mod : 1603 Model : 102-202-212-212 A . JUN 97 R Mod : 1603 Model : 102-202-212-212 A . JUN 97 R Mod : 1603 Model : 102-202-212-212 A . JUN 97 R Mod : 1603 Model : 102-202-212-212 A . Output voltage 115 V ± 4V and 26 V ± 1V .06. Note : Two AC electrical networks are supplied by the inverters : 115 VAC and 26 VAC. The inverters are rack mounted and cooled by forced air with provisions for natural convection cooling.Type single phase The two inverters are powered respectively from DC BUS 1 and DC BUS 2. or by HOT EMER BAT BUS in OVRD configuration or by TRU when selected ON.ELECTRICAL SYSTEM 1. Only one is shown on the schematics. but corresponding AC BUS is supplied by AC BTR (BTC pb).30 P1 AC CONSTANT FREQUENCY 020 SEP 05 AA 30.1 DESCRIPTION GENERATION R The source of constant frequency (400 Hz) AC power consists of two static inverters (INV). The static inverter design characteristics are as follows : . In event of both DC BUS power loss. corresponding inverter is not supplied.Power 500 VA . The input voltage range is between 18 VDC and 31 VDC for satisfactory operation.Frequency 400 Hz ± 5 Hz . Mod : 1603 . In event of one DC BUS loss. INV1 is automatically supplied by HOT MAIN BAT BUS. The maximum power available on each 26 VAC BUS is 250VA. ELECTRICAL SYSTEM 1. AC STBY BUS is automatically supplied from INV 2.AC BUS 1 . . provided the BTC pb is not in ISOL position.AC STBY BUS 1 INV 2 normally supplies : .AC BUS 2 In event of inverter failure or input power loss the associated AC BUS is isolated from affected inverter and. In event of INV 1 failure or input power loss.06.30 P2 AC CONSTANT FREQUENCY 001 DEC 96 AA DISTRIBUTION (115 and 26 V) INV 1 normally supplies : . The AC BUS 1 and 2 are automatically tied together. SEP 05 R FAULT . . . . . . . . . . . . . CONSOLE.CAPT LTS : DOME.NAVIGATION (BackĆup of DC SVCE BUS) .F/O PANELS .06.60 P1 DISTRIBUTION EQUIPMENT LIST AA 070 JUN 97 Note : *" = option DC BUS 1 ATA SYSTEM FUNCTION 21 AIR CONDITIONING .DUCT/COMPT Cockpit and cabin Temperature IND .LEFT STICK SHAKER 28 FUEL .EXTRACT FAN PWR SPLY (BackĆup of DC BUS 2) 23 COMMUNICATIONS * .ANNUNCIATOR LT TEST .ELECTRICAL SYSTEM 1. : 4366 CAPT STATIC PORTS STBY STATIC PORTS LH SIDE WINDOW ANTI ICING RH WINDSHIELD HTG IND .BEACON (BackĆup of DC SVCE BUS) Mod.MFC 1B (Primary) 33 LIGHTS .TURBOFAN SOV 1 CTL . CHARTHOLDER.Automatic Pressure CTL .SPOILERS IND .TANK TEMP IND 30 ICE AND RAIN PROTECTION - 31 INDICATING/REĆ CORDING .F/O DOME (Normal) .LP VALVE 1 (Normal) .STORM .STICK PUSHER PWR and CTL .HF 1 when two HF are installed * .FLIGHT INTERPHONE and AUDIO CONTROL PANEL OBSV .HF FERRY 27 FLIGHT CONTROLS . READING .GENERAL ILLUMINATION : LEFT LATERAL RAMP (1 FLUORESCENT LIGHT OUT OF 2) .SEL CAL * . ELECTRICAL SYSTEM 1.06.60 P2 DISTRIBUTION EQUIPMENT LIST 170 JUN 97 AA ATA SYSTEM 34 NAVIGATION - FUNCTION WEATHER RADAR RADIO ALTIMETER GPWS - G/S IND STBY ALTIMETER VIBRATOR DME 1 AHRS 2 (auxiliary) BackĆup - DC BUS 2, IN FLIGHT (Primary) - DC EMER BUS, ON GROUND (Auxiliary) 36 PNEUMATIC - BLEED LEAK IND - CROSS FEED VALVE and IND - HP VALVE 1 61 PROPELLERS - OVSPD TEST ENG 1 - AFU 1 (Normal) - BALANCE TEST 73 ENGINE FUEL and CTL - 79 OIL - PRESS, TEMP IND 1 Mod. : (3973 or 4371 or 4457) + 4366 FUEL FLOW, FUEL USED IND 1 FUEL TEMP IND 1 FUEL CLOG IND 1 EEC 1 GROUND IDLE SOLENOID SPLY ELECTRICAL SYSTEM 1.06.60 P3 170 DISTRIBUTION EQUIPMENT LIST JUN 97 AA DC BUS 2 ATA SYSTEM 21 AIR CONDITIONING FUNCTION - Landing elevation IND - TURBOFAN SOV 2 CTL - EXTRACT FAN PWR SUPPLY (Primary) 23 COMMUNICATION - VHF 2 26 FIRE PROTECTION - NAC 1 (when installed) and 2 OVHT DET 27 FLIGHT CTL - PITCH TRIM STBY COMMAND (BackĆup of DC EMER BUS for NORMAL command) - RIGHT STICK SHAKER 28 FUEL - LP VALVE 2 (Normal) 29 HYDRAULIC POWER - DC AUX HYD PUMP NORM CTL, IND and PWR in flight 30 ICE and RAIN PROTECTION - 31 INDICATING/REĆ CORDING - F/O CLOCK LANDING GEAR - WOW 2 CTL - Secondary IND 32 Mod. : 4116 + 4366 DE ICE VALVES ENG 2 BOOTS A and B (Normal) WINGS and EMPENNAGE BOOTS B (Normal) F/O WIPER F/O STATIC PORTS F/O PROBES IND LH WINDSHIELD HTG INDICATOR RH SIDE WINDOW ANTI ICING - MFC 2B (Primary) ELECTRICAL SYSTEM 1.06.60 P4 DISTRIBUTION EQUIPMENT LIST 070 JUN 97 AA ATA SYSTEM 33 LIGHTS FUNCTION - PASSENGER SIGNS - WING LIGHTS - F/O LTS : CHARTHOLDER, CONSOLE, READING - UTILITY SPOT and FLOOD - NORMAL INSTRUMENTS SPLY and LABELS INTEGRATED LT CTL - TAXI and TAKE OFF CTL - GENERAL ILLUMINATION : RIGHT LATERAL RAMP (1 FLUORESCENT LIGHT OUT OF 2) 34 NAVIGATION 36 PNEUMATIC - HP VALVE 2 52 DOORS - ALERTS 61 PROPELLERS - OVSPD TEST ENG 2 - AFU 2 (Normal) 73 ENGINE FUEL and CTL - 79 OIL - Press, Temp IND 2 Mod. : 4366 * - ATC 2 * - DME 2 - VOR/ILS 2 * - ADF 2 - CAPT RMI - SGU 2 - F/O EADI - AHRS 1 (Auxiliary) (BackĆup of DC EMER BUS) - AHRS 2 (Primary) - F/O EHSI FUEL FLOW, FUEL USED IND 2 FUEL TEMP IND 2 FUEL CLOG IND 2 EEC 2 (Normal) IDLE GATE FAIL IND ELECTRICAL SYSTEM 1.06.60 P5 DISTRIBUTION EQUIPMENT LIST 070 JUN 97 AA HOT EMER BAT BUS ATA SYSTEM 24 ELECTRICAL POWER FUNCTION - DC EMER BUS AND DC STBY BUS CTL (BUSSES REMAIN SUPPLIED BY DC BUS 1) - EMER BAT AMMETER - EMER BAT VOLT IND - EMER BUS and INV 1 ON EMER BAT IND (ARROW) - TRU CTL and IND (when installed) 31 INDICATING/REĆ CORDING - MFC 1 MOD A (Auxiliary) (BackĆup of DC ESS BUS) 34 NAVIGATION - STBY HORIZON BackĆup - ADC 1 (BackĆup) - ADC 2 (BackĆup) Mod. : 4366 ELECTRICAL SYSTEM 1.06.60 P6 DISTRIBUTION EQUIPMENT LIST 250 SEP 04 AA HOT MAIN BAT BUS ATA SYSTEM 24 ELECTRICAL POWER - R FUNCTION DC EXT PWR CONTACTOR CTL MAIN BAT AMMETER ESS BUS and INV 1 ON MAIN BAT IND (ARROW) DC GND/HDL XFR BUS SPLY (Back-up of EXT PWR) MAIN BAT VOLT IND MAIN and EMER BAT CHGE INHIBIT DC ESS BUS and INV 1 CTL (REMAIN SUPPLIED BY DC BUS 1) ENG FIRE EXTINGUISHING CTL and IND (Back-up of DC EMER BUS) DC AUX HYD PUMP GND SPLY, CTL and IND (Back-up of DC BUS 2) MFC 2A (Auxiliary) (Back-up of DC EMER BUS) 26 FIRE DETECTION - 29 HYDRAULIC POWER - 31 INDICATING/ RECORDING - 33 LIGHTS - ENTRANCE 61 PROPELLERS - A/F AUX PUMPS PWR Mod : 3552 Model : 102-202-212-212 A ELECTRICAL SYSTEM 1.06.60 P7 DISTRIBUTION EQUIPMENT LIST AA 390 JUL 00 DC EMER BUS ATA SYSTEM 21 AIR CONDITIONING 22 AUTOĆFLIGHT 23 COMMUNICATIONS 24 ELECTRICAL POWER 26 FIRE DETECTION 27 FLIGHT CONTROLS 28 FUEL 29 HYDRAULIC POWER 30 ICE and RAIN PROTECTION R Mod. : 4366 + (4373 or 8167) FUNCTION - OVERBOARD and UNDERFLOOR VALVES CTL and IND and AIR COOLING HIGH FLOW IND - PRESSURE IND and EXCESS ALTITUDE IND - PNEUMATIC OUTFLOW VALVES - AP/FD COMPUTER and GUIDANCE IND (when installed) - VHF - F/O COCKPIT AMPLIFIER - RCAU - GCU 1 DC (BackĆup) - AC BUS OFF 1 and 2 IND - ACW BUS OFF 1 and 2 IND - INV FAULT 1 IND - DC BUS OFF 1 and 2 IND - BPCU DC (BackĆup) - DC STBY BUS IND (UNDV - OVRD) - EMER BAT : CHG IND - DC STBY BUS CTL (BUS REMAIN SUPPLIED BY DC BUS 1) - AC 1 BUSSES CTL (BUSSES REMAIN SUPPLIED BY INV 2) - ENG FIRE EXTINGUISHING CTL and IND (NorĆ mal) - FIRE HANDLE IND ENG 1 and 2 - FIRE DETECTION ENG 1 and 2 - PITCH TRIM NORMAL COMMAND - RUDDER TRIM - AILERON TRIM - AILERON LOCKING IND - LP VALVES 1 and 2 and IND (BackĆup of DC BUS 1 - DC BUS 2) - BLUE PUMP CTL and IND - GREEN PUMP IND - AAS IND and ALERTS - AAS CTL - WING, EMPENNAGE BOOTS A - ENG 1 BOOTS A and B - PROPELLERS 1 and 2 ANTIĆICING CTL and IND - WING, EMPENNAGE BOOTS B and ENG 2 BOOTS A and B (BackĆup of DC BUS 2) - HORNS ANTIĆICING IND and CTL Model : 102Ć202Ć212Ć212A ELECTRICAL SYSTEM 1.06.60 P8 DISTRIBUTION EQUIPMENT LIST 570 JUN 97 AA ATA SYSTEM 31 INDICATING/REĆ CORDING FUNCTION - CAPT CLOCK - FDAU ON GROUND - MFC 1B (Auxiliary) (BackĆup of DC BUS 1) - MFC 2A (Primary) - MFC 2B (Auxiliary) (BackĆup of DC BUS 2) 32 LANDING GEAR - ANTISKID OUTBOARD - NOSE WHEEL STEERING - WOW 1 CTL 33 LIGHTS - 34 NAVIGATION - ATC 1 - AHRS 2 (ON GROUND, Auxiliary) (BackĆup of IN FLIGHT : - DC BUS 2, Primary) - DC BUS 1, Auxiliary) - AHRS 1 (Primary) - ADC 1 (Primary) - ADC 2 (Primary) 61 PROPELLERS - 73 ENGINE FUEL and CTL - EEC 1 and 2 PWR and IND (BackĆup of - DC BUS 1) - DC BUS 2) 76 ENGINE CTL - CL FIRE IND 1 and 2 Mod. : 4366 CAPT PANELS PYLON STBY COMPASS LAVATORY (EMERGENCY) F/O DOME (BackĆup of DC BUS 1) AFU 1 and 2 (BackĆup of DC BUS 1 - DC BUS 2) A/F AUX PUMPS CTL TORQUE IND 1 and 2 PEC 1 and PEC 2 (Normal), associated PVM and PIU. Eng : PW127F ELECTRICAL SYSTEM 1.06.60 P9 DISTRIBUTION EQUIPMENT LIST 350 JUL 00 AA DC ESS BUS ATA SYSTEM 21 AIR CONDITIONING - 22 AUTOFLIGHT 23 COMMUNICATIONS 24 ELECTRICAL POWER 26 FIRE DETECTION - 27 FLIGHT CONTROLS 28 FUEL - R Mod. : (3529 or 3530) + (4373 or 8167) - FUNCTION PACK 1 and RECIRC FAN 1 IND PACK 1 VALVE EXTRACT FAN CTL PACK 2 and RECIRC FAN 2 IND PACK 2 VALVE LANDING ELEVATION IND (ALPHANUMERIC DISPLAY) COCKPIT and CABIN AUTOMATIC and MANUAL TEMPERATURE CTL and IND AP OFF IND AP DISC BY QUICK DISCONNECT FLIGHT INTERPHONE CAPT and F/O AUDIO CONTROL PANELS CAPT and F/O CAPT COCKPIT AMPLIFIER PASSENGER ADDRESS MECHANIC CALL COCKPIT and CABIN CREW CALL CVR GCU 2 DC (BackĆup) GCU 1 and 2 ACW (Back-up) BPCU ACW (Back-up) DC SVCE and UTLY BUSSES 1 and 2 CTL MAIN BAT CHG IND INV 2 FAULT IND AC 2 and STBY BUSSES CTL (BUSSES REMAIN SUPPLIED BY INV 1) TOILETS SMK DET AVIONICS SMK DET FWD and AFT COMPT SMK DET AFT COMPT and TOILETS DET FANS CTL and IND - FWD COMPT DET FANS CTL and IND CLUTCH REENGAGEMENT SYSTEM AILERON LOCKING CTL FQI and 2 CROSS FEED VALVE STARTING PUMP 1 and 2 and MOTIVE FLOW VALVES 1 and 2 Model : 102-202-212-212A 160 JUN 97 R Mod : (3973 or 4371 or 4457) + 4116 EMERGENCY 34 NAVIGATION Mod.ADU 27 FLIGHT CONTROLS .INTERCONNECTING VALVE 32 LANDING GEAR - 33 LIGHTS .06.60 P 11 DISTRIBUTION EQUIPMENT LIST 070 JUN 97 AA DC STBY BUS ATA SYSTEM 22 AUTO FLIGHT FUNCTION .AP/FD SERVOS .ELECTRICAL SYSTEM 1. : 4366 * - PRIMARY IND LANDING GEAR CTL (HYDRAULIC VALVE) ANTISKID INOP and BRAKE OVTEMP IND R and L INBOARD and OUTBOARD BRAKES TEMP XMITTERS VOR/ILS/MKR 1 SGU 1 CAPT EADI RMI F/O ADF 1 OMEGA CAPT EHSI .FLAPS CTL 29 HYDRAULIC POWER .PRESS TRIPLE IND . . . . SEP 04 R Model : 102-202-212-212 A . . . . . . . . . . . . JUN 97 R Model : 102-202-212-212 A . CREW MEMBERS (See schematic p 9/10) The quick donning masks are stowed in readily accessible containers adjacent to each crew member seat. R Oxygen system capability allows to supply 25 % of the passengers with a continuous diluted flow for a duration of 30 mn in case of pressure drop (4 minutes to descend from 25 000 ft to 13 000 ft. A mask mounted diluter demand regulator provides dilution and emergency pressure control : . . Each of these bottles is equipped with a diluted oxygen mask. A high pressure cylinder of 2180 l (77 cu. It controls flow to oxygen masks installed under the hat racks.07.25 cu ft) portable oxygen bottles are stowed under the cabin attendant seats. with the dilution control in the 100 % position. diluted oxygen is provided to 3 cockpit cew for a duration of 120 mn at demand flow (10 minutes to descend from 25 000 ft to 13 000 ft and continuation of flight R between 13 000 ft and 10 000 ft for 110 minutes). 26 minutes to continue the flight between 13 000 ft and 10 000 ft) in addition to the cockpit crew 120 minutes consumption.EMERGENCY EQUIPMENT 1. PASSENGERS The main system provides passenger protection through a PAX SUPPLY valve. The mask harness inflates automatically when the mask is pulled out of the container and it can easily be donned with one hand.In case of smoke or noxious gas emission. Mod : 4411 . They permit a continuous diluted flow to both cabin attendants at 13 000 ft for a duration of 30 mn.ft) capacity.Protective breathing equipments are located in the baggage box of the cabin.1 DESCRIPTION (See schematic p 7/8) The system consists of a main system supplying the cockpit crew and the passengers and two portable units for the cabin attendants. oxygen is provided to 3 cockpit crewr for a duration of 15 minutes at demand flow. CABIN ATTENDANT PORTABLE UNITS R . supplies oxygen to the distribution system. with the dilution control in the N (normal) position.20 P1 OXYGEN SYSTEM 050 JUL 99 AA 20.In case of pressure drop.two 120 l (4. Nominal charge pressure of 1850 PSI is reduced to 78 PSI in the distribution system. R . A discharge part is provided to face overpressures. quantity must be checked to be adequate for intended flight (refer FCOM 2. OFF illuminates white. pb pressed in : The valve is open. 3 PAX SUPPLY pb ON : pb released : (pb pressed in) Passengers supply valve is open. If preflight pressure is below 1400 PSI. LO PR : illuminates amber and the CCAS is activated when a low pressure (below 50 PSI) is detected in the low pressure distribution circuit. 2 MAIN SUPPLY pb Controls the low pressure supply solenoïd valve.05).20 P2 OXYGEN SYSTEM 001 DEC 96 AA 20. OFF : (pb released) the valve is closed. The scale is marked by a red arc from 0 to 85 PSI and by a green arc from 85 to 2025 PSI. ON illuminates blue.2 CONTROLS OXYGEN PANEL 1 HP ind.07.EMERGENCY EQUIPMENT 1.01. . Passengers supply valve is closed. low pressure oxygen is supplied to the cockpit crew oxygen masks. Oxygen bottle pressure is displayed in PSI X 1000. . . . 20 P 7/8 OXYGEN SYSTEM 001 SEP 04 AA 20.4 SCHEMATICS R (number depends on configuration) .EMERGENCY EQUIPMENT 1.07. . EMERGENCY EQUIPMENT 1.07.30 P1 EMERGENCY EVACUATION 370 SEP 03 AA EMERGENCY EVACUATION IN CASE OF DITCHING R A EMERGENCY EXIT B SERVICE DOOR C PASSENGER/CREW DOOR Note : Be sure that safety pin is removed. : 4019 + 4358 Model : 102-202-212-212 A . Mod. : 4019 + 4358 Model : 102-202-212-212 A .07. Mod.EMERGENCY EQUIPMENT 1.30 P2 EMERGENCY EVACUATION 370 JUL 99 AA EMERGENCY EVACUATION IN CASE OF FORCED LANDING A EMERGENCY EXIT B SERVICE DOOR C PASSENGER/CREW DOOR Note : Be sure that safety pin is removed. . FIRE illuminates on CAP in case of : . When right nacelle temperature exceeds 170°C. CARGO AND TOILETS SMOKE DETECTION SYSTEM R Forward cargo and after cargo are each equipped with one optical smoke detector. Ambient transmittance is monitored by reflection measurement.Two identical detection loops (A and B) mounted in parallel.30 p.Fire signal detected by one of the 2 loops if the other one is selected OFF. SMOKE" signal is sent to CCAS through the MFC. Red ENG. Smoke detection between the avionics compartment and the extract fan activates a "ELEC SMK" red alert on CAP. 3) The avionics extract air duct is provided with a smoke detection device. The detection principle is based on the variation of resistance and capacitance of the detection cable (fire signal).Extinguishing for : • each engine • cockpit.Detection for : • each engine fire • right nacelle overheat (on ground only) • each cargo compartment and toilets smoke • avionics compartment smoke .FIRE PROTECTION 1. If there is only a change in resistance.A fire detection control unit.10 P1 GENERAL 001 JUL 99 AA 10. associated loop will be declared failed by the fire detection control unit (fault signal). . AVIONICS SMOKE DETECTION (See schematic 1.Fire signal detected by both loops A and B or.08.03. and the CCAS is activated. . cabin and each cargo compartment • toilets waste bin ENGINE FIRE DETECTION SYSTEM Each engine is equipped with a fire detection system which consists of : . NAC2 OVHT red alarm is triggered on CAP. Toilets are equipped wih one photoelectric smoke detector.1 DESCRIPTION The fire protection system is provided in order to ensure : . In case of smoke detection. linked to the CCAS. . R RIGHT NACELLE OVERHEAT DETECTION SYSTEM (on ground only) Right nacelle is equipped with an overheat detector. JUN 97 R Model : 102-202-212-212 A . . LOOP amber light illuminates on CAP.FUEL SO light illuminates in associated CL if CL isn't in fuel shut off position .FAULT lights of both LOOP A and LOOP B pb illuminate . ENG FIRE red light illuminates on CAP.CCAS is activated. The OFF light illuminates white. . R CL FUEL LT Illuminates red in case of fire signal from associated engine. FIRE .08. LOOP amber light illuminates on CAP.ENG FIRE red light illuminates in associated fire handle .CCAS is activated. Extinguishes after CL is set at fuel shut off position or if fire detection signal terminates. TEST sw Spring Loaded in neutral position Spring loaded in neutral position allows a test of the detection of fire and fault signals when both LOOP pbs are selected ON : FAULT . FAULT : The light illuminates amber and the CCAS is activated when the associated pb is selected ON and a fault signal is generated by the fire detection control unit.FIRE PROTECTION 1. 5 OFF : (pb released) Aural and visual alerts are inhibited for the related loop.10 P4 001 GENERAL JUL 98 AA 4 LOOP pb Allows activation of aural and visual alerts when a fire signal (FIRE) or a fault signal (LOOP) is generated by the fire detection control unit for the related loop. pb pressed in : Aural and visual alerts are activated when a fire or a fault signal is generated by the fire detection control unit for the related loop. LOOP amber light illuminates on CAP. JUN 97 R Model : 102-202-212-212 A . FIRE PROTECTION 1.10 P6 GENERAL AA LEFT INTENTIONALLY BLANK 001 JUN 97 .08. . JUN 97 R Model : 102-202-212-212 A . 3 GUST LOCK DESCRIPTION CONTROLS ELECTRICAL SUPPLY/MFC LOGIC/SYSTEM MONITORING R Mod.60 60.4 YAW DESCRIPTION CONTROLS ELECTRICAL SUPPLY/MFC LOGIC SCHEMATIC 1.2 50.09.50 50.5 FLAPS DESCRIPTION CONTROLS ELECTRICAL SUPPLY/MFC LOGIC/SYSTEM MONITORING LATERAL MAINTENANCE PANEL SCHEMATIC 1.4 30.1 50.1 40.10 GENERAL 1.3 30.1 20.3 ROLL DESCRIPTION CONTROLS ELECTRICAL SUPPLY 1.2 40.09.09.00 P1 060 CONTENTS AA 1.3 40.2 60.FLIGHT CONTROLS 1.20 20.40 40.4 50.09.30 30.2 20.5 PITCH DESCRIPTION CONTROLS ELECTRICAL SUPPLY/MFC LOGIC/SYSTEM MONITORING LATERAL MAINTENANCE PANEL SCHEMATIC 1.2 30.09.3 50. : 4373 or 8167 JUL 00 .1 60.09.09.1 30.09.00 CONTENTS 1. On each wing. Spoilers are hydraulically actuated (blue system). On each wing flaps are provided in two parts (inboard and outboard) mechanically linked and hydraulically activated (blue system).a locking mechanism on pitch and roll axes .A rudder (spring-tab equipped) Ailerons.a damping unit on yaw axis limiting rudder excessive travel speed.Two elevators (servo-tab equipped) .10 P1 GENERAL 060 JUL 00 AA The control of the aircraft is achieved on the three axes by : . Wind protection is achieved on ground by : . elevators and rudder are mechanically actuated.FLIGHT CONTROLS 1. : 4373 or 8167 . one aileron (spring tab equipped) and one spoiler .09. R Mod. 20 P1 ROLL 060 JUL 01 AA 20. Full roll trim travel requires about 30 s. SPRING TAB A spring tab provides a flexible compensation which automatically increases with the aerodynamic loads applied on the ailerons. It is electrically controlled from a twin control sw through an electrical actuator. R LH aileron trim controlled tab travel : 6. thus ensuring a reduction of the pilot's efforts. 6.7° up.7° down.1 DESCRIPTION Roll control is achieved through control wheels. 14° down ROLL TRIM Aileron trim is performed by varying the neutral position setting of the left aileron spring tab with respect to the aileron.09.FLIGHT CONTROLS 1. R Wheel travel :  87° Ailerons travel : 14° up. Mod : 4373 or 8167 . . . . . . . . . . . . . . . . . . . SEP 03 R . ELEVATOR MECHANICAL LOCKING DEVICE The system immobilizes the control column in pitch and therefore control surfaces. Each locking device is electrically actuated through switches installed on the gust lock lever. Note : Ailerons may be locked slightly beyond the neutral position.60 P1 GUST LOCK 060 JUL 00 AA 60. R Mod : 4373 or 8167 . This system provides protection against take off with gust lock engaged. AILERON ELECTRO MECHANICAL LOCKING DEVICE The system is composed of two electro-mechanical locking devices immobilizing one aileron each. Therefore the control wheel may be tilted (5°). This device is controlled by a control lever located on the pedestal and mechanically operated through cables and gears.09. according to the actual position of ailerons. This system includes an elevator mechanical locking device and an aileron electrico-mechanical locking device.1 DESCRIPTION A gust lock system is provided to protect the pitch and roll flight controls on ground and to limit the PL travel slightly below FI. or too high power setting when in hotel mode.FLIGHT CONTROLS 1. AIL LOCK lt Illuminates amber and the CCAS is activated through the MFC whenever one of the locking actuators is in disagreement with the gust lock lever position (Lock or unlock position).60 P2 GUST LOCK 060 JUL 00 AA 60.2 CONTROLS GUST LOCK LEVER When the gust lock is engaged. R Mod : 4373 or 8167 .FLIGHT CONTROLS 1. the PL travel is limited slightly below FI to provide protection against take off and two high power setting when in hotel mode. The gust lock handle can be put into the locking notch whatever the position of the flight controls but these controls must be brought to neutral to positively engage the locking devices.09. Aural alert is Continuous Repetitive Chime (CRC) . . "FLT CTL" amber light illuminates on CAP .O. .Disagree between Aileron locking actuators and gust lock control (Temporized alert 8 sec). "CONFIG" red light illuminates on CAP .Disagree between Aileron locking actuators and gust lock control during the T. Aural alert is Continuous Repetitive Chime (CRC) R Mod : 4373 or 8167 . "FLT CTL"amber light illuminates on CAP . "MW" flashing red .60 P3 060 GUST LOCK JUL 00 AA 60. SYSTEM MONITORING The following conditions are monitored by visual and aural alerts. "CONFIG" red light illuminates on CAP .Aileron locking actuators not fully retracted and PL on TO position . "MW" flashing red .3 ELECTRICAL SUPPLY/MFC LOGIC/SYSTEM MONITORING ELECTRICAL SUPPLY EQUIPMENT DC BUS SUPPLY (C/B) Aileron lock ind. "AIL LOCK" amber light illuminates on the pedestal . "MC" flashing amber . CONFIG TEST. Aural alert is single chime (SC) R R R R R R R . "FLT CTL"amber light illuminates on CAP . . DC EMER BUS (on lateral panel AIL TRIM & AIL LOCK WARN) Aileron lock CTL DC ESS BUS (on lateral panel GUST LOCK AIL) MFC LOGIC See chapter 1.01.09.FLIGHT CONTROLS 1. . other systems : AHRS 2. ADC 1 supplies : .total air pressure provided by its specific pitot probe.10 P1 AIR DATA SYSTEM 030 SEP 02 AA 10. Each computer is supplied with : . vertical speed ind. From this data. FDAU. AFCS.true air speed (TAS).other systems : AHRS 1.static air pressure provided by its specific static ports.CAPT flight instruments (altimeter. . FDAU. STANDBY SYSTEM The standby system consist of : ..a pitot probe. each ADC computes : . AFCS ATC 1 and ATC 2 through TCAS controller box and TCAS through ATC 1 and ATC 2 (if installed and mode S only). MFC.).static air temperature (SAT). .vertical speed. TAT/SAT/TAS indicator and GPS (if installed) are supplied either by ADC 1 or ADC 2 according to ADC selector on capt panel.one standby system.pressure altitude.1 DESCRIPTION (See schematic p 13/14) The flight environment data are provided by three independent air data systems : .F/O flight instruments (altimeter.total air temperature (TAT).total air temperature provided by its specific TAT probe.10.indicated air speed (IAS). . ADC 2 supplies TCAS through ATC 1 and ATC 2. vertical speed ind. . . pressurization. airspeed ind. . ATC 1 and ATC 2. pressurization.).two static ports. . . .FLIGHT INSTRUMENTS 1.two main systems. Note : If ATC 2 mode S is installed. MAIN SYSTEMS Aircraft is equipped with two independent AIR DATA COMPUTERS (ADC). Probes and ports are located on the LH and RH side of the fuselage and are electrically heated. airspeed ind.. GPWS. EEC's. . and standby altimeter are directly supplied by raw data. MFC. ATC 1. Standby airspeed ind. . ADC 2 supplies : . R Mod : 3074 or 3113 or 3625 or 3832 or 5103 or 5146 or 8259 . . . . 000 70 200 R 20. Altitude pointer One revolution of pointer represents 1000 ft altitude change.10. Note : Allowable deviation between normal altimeter indications and between normal and standby altimeter indications : 3 4 FL (ft) NORM/NORM (ft) NORM/STBY (ft) R 0 55 70 R 5. A black and white flag marks the LH drum (ten thousands) when altitude is between 0 and 9999 ft.000 100 260 R 25. Altitude counter The digital counter is equipped with three drums indicating ten thousands.10 P5 001 AIR DATA SYSTEM JUL 01 AA STANDBY ALTIMETER 1 Baroset value is displayed in millibars (875 to 1 050 mb).000 60 150 R 10. 2 Baroset knob Sets barometric reference on mb counter.FLIGHT INSTRUMENTS 1. thousands and hundreds of fett. An orange and white flag marks the two LH drums (ten thousands and thousands) when altitude is below 0 ft.000 120 300 R R . low altitude climb.Select the 12 nautical mile range for high altitude cruise.Select the 6 nautical mile range for take-off. (2) . (6) . The range selected has no effect on the TCAS system logic used to determine TA's and RA's. 1.Light sensor (8) . R Mod :3074 or 3113 or 3625 or 3832 or 5103 or 5146 or 8259 . viewing of traffic from 2700 ft below to 2700 ft above.1. (5) . (3) . approaches and landings .05.Extended altitude surveillance status When selected ABV or BLW . indicated vertical speed range.20) (1) . (7) .BLW viewing of traffic from 2 700 ft above to 9 900 ft below.Test When depressed. In normal position.10 FLIGHT INSTRUMENTS P6 030 AIR DATA SYSTEM SEP 02 AA TCAS VERTICAL SPEED INDICATOR (TCAS VSI ) (cf. (4) .ABV viewing of traffic from 2 700 ft below to 9 900 ft above .Fixed aircraft mock-up The fixed aircraft mock-up is surrounded with a 2 mn loop.Vertical speed pointer Indicates rate of climb/descent from 0 to "6 000 ft/mn.10. indicator will display a test pattern.Vertical speed prohibited arc (red) Red arc indicates that pilot is advised to fly out of.Vertical speed recommended arc (green) Green arc indicates vertical speed range to fly in.Display range selection The following ranges for the sextant TCAS indicator are recommended : . or not enter. 3 Test When depressed. Display accuracy is
40 ft/mn. 4 Light sensort 5 Not available Mod : 4541 . indicator will display a test pattern. the vertical speed pointer disappears when V/S flag appears. 2 Vertical speed flag Appears if the indicator is not able to display vertical speed information In that case. From 0 to 1000 ft/mn the scale is graduated in 100 ft/mn increments.FLIGHT INSTRUMENTS 1. and from 1000 to 6000 ft/mn in 500 ft/mn increments.10 P6 AIR DATA SYSTEM 040 DEC 97 AA TCAS VERTICAL SPEED INDICATOR (TCAS FUNCTION NOT AVAILABLE) 1 Vertical speed pointer Indicates rate of climb/descent from 0 to
6000 ft/mn.10. TA ONLY" indication .in case of self test activation. Resolution advisory flag .10.Proximity : filled diamond (Cyan) . the vertical speed pointer disappears when V/S flag appears.Value : two digits (color of the associated symbol) .Unit : ft x 100 .in case of TCAS fails.positive = the intruder is above .TCAS TA : filled circle (amber) .Appears if the indicator is not able to display vertical speed information .FLIGHT INSTRUMENTS 1.when the TCAS is in STBY mode.Arow to the top : intruder climbing . Intruder relative altitude .This flag appears if the TCAS is in TA ONLY" mode.negative = the intruder is below Relative vertical speed indicator .Sign : . TCAS FAIL" appears or .Appears only if the indicator is not able to display RA's or vertical speed.10 P7 AIR DATA SYSTEM 030 SEP 02 AA 9 10 11 12 13 14 15 R Intruder symbol . Mod : 3074 or 3113 or 3625 or 3832 or 5103 or 5146 or 8259 .Arrow to the bottom : intruder descending Vertical speed flag . TD FAIL" appears or .Others : blank diamond (Cyan) Center of the symbol shows the intruder relative position. TEST" appears.In that case. TCAS OFF" appears or .TCAS RA : filled square (red) .If the indicator is not able to display intruder's. Traffic function flag . . . 10.FLIGHT INSTRUMENTS 1.10 P 10 AIR DATA SYSTEM Mod : 3973 or 4371 or 4457 050 JUN 97 . 05. and vertical speed ind.Nil - 26 VAC BUS 2 (on overhead panel ASI VSI) F/O altimeter . .12 Mod.Loss of ADC D See ADC FAULT procedure in chapter 2.FLIGHT INSTRUMENTS 1.Nil - 26 VAC STBY BUS (on overhead panel ALTM) F/O airspeed ind.05. and vertical speed ind.Nil - Standby altimeter vibrator SYSTEM MONITORING The following conditions are monitored by visual alerts : .3 ELECTRICAL SUPPLY/SYSTEM MONITORING ELECTRICAL SUPPLY EQUIPMENT ADC 1/ADC 2 DC BUS SUPPLY (C/B) HOT EMER BAT BUS (BackĆup on overhead panel ADC 1/2 HOT) DC EMER BUS (primary on overhead panel ADC 1/2 EMER) AC BUS SUPPLY (C/B) .12 . .Nil - CAPT airspeed ind. TAS/Temperature ind.Incorrect ADC switching D See ADC SW FAULT procedure in chapter 2.Nil - 26 VAC STBY BUS (on overhead panel ASI VSI ALTM) CAPT altimeter + recording FDAU . : 4366 .10.Nil - 26 VAC BUS 2 (on overhead panel ALTM) DC BUS 1 (on overhead panel STBY ALTM) .10 P 11/12 070 AIR DATA SYSTEM JUN 97 AA 10. SEP 02 R Mod : 3074 or 3113 or 3625 or 3832 or 5103 or 5146 or 8259 . 4_.Two attitude . three gyrometers and three accelerometers are aligned with the aircraft axes as a strapdown system.ASCB bus . (AHRU) .FLIGHT INSTRUMENTS 1.CAPT RMI (heading) .heading) . weather radar and FDAU.SGU 2 (attitude and heading) . The IMU components. latitude or variation insertion. is used to compute gyro erection.SGU 1 (attitude and heading) .Standby instruments AHRS AHRS consists of : . AHRS 1 supplies : . Vertical accuracy remains within $1.Radar (attitude) . Each AHRU receives inputs from its associated flux valve. TAS.ASCB bus AHRS 2 supplies : . a microprocessor and electronic controls.heading reference units. Earth rotation and gyro drift are computed without requiring heading. AHRU sends altitude and heading signals to indicators. fed by both ADC.F/O RMI (heading) .Two main systems (AHRS) .One dual remote compensator Each AHRU includes an inertial measurement unit (IMU).10. AFCS.Two flux valves .1 DESCRIPTION The attitude and heading data are provided by : .20 P1 ATTITUDE HEADING REFERENCE SYSTEM 001 DEC 96 AA 20. heading accuracy within $2_.FDAU (attitude . Note : However. Relative bearing info is lost. STANDBY COMPASS A retractable standby magnetic compass with internal lighting is provided under glareshield. two pointers with rabbit ears" switching to present either VOR or ADF bearings. . .20 P2 ATTITUDE HEADING REFERENCE SYSTEM 001 DEC 96 STANDBY INSTRUMENTS STANDBY HORIZON A stand-by electrical horizon is provided on the central panel. RADIO MAGNETIC IND (RMI) An RMI is installed on each pilot's panel coupled to the opposite AHRS.In case of navigation system indication failure or data supply failure. validity of these information should be confirmed. .ADF needle displays only relative bearing to station (without indication of magnetic bearing). Each includes a compass rose.In case of RMI internal failure or AHRS supply loss : .10. the associated pointers move to 3 o'clock position except the double pointer when ADF is selected (9 o'clock position).VOR needle displays magnetic bearing to station on rose card (no metter when the card is frozen). .FLIGHT INSTRUMENTS 1.RED OFF" flag appears . showing magnetic heading. 2 CONTROLS AHRS ERECT PB Iluminates amber when the associated AHRS loses the TAS signal from the ADC.20 P3 ATTITUDE HEADING REFERENCE SYSTEM 001 DEC 96 20. Â VOR/ADF selectors Select the stations (VOR or ADF) associated to the bearing pointers. the pb remains illuminated as long as the TAS signal is lost RMI À Compass card. If the aircraft is stabilized (unaccelerated level flight) a gyro fast erection may be performed by depressing the associated pb for 15 s. The AHRS will continue to operate without auto-erect capability. When released. Á Bearing pointers Indicate the magnetic bearing to the station selected by the associated VOR/ADF selector. Ã Red OFF" flag Appears in case of RMI internal failure or AHRS supply loss. .FLIGHT INSTRUMENTS 1. Displays heading information on a rotating heading dial graduated in 5 degree increments.10. 20 P4 ATTITUDE HEADING REFERENCE SYSTEM LEFT INTENTIONALLY BLANK 001 DEC 96 .FLIGHT INSTRUMENTS 1.10. Roll angle is given by a scale marked at 10. 60 and 90 degrees. causes a rapid erection if the instrument is powered. . 30. The compass rose is graduated in 10 degree increments. STANDBY COMPASS Hidden in up position. represents the aircraft position on the attitude sphere. or when gyroscope speed becomes insufficient. Ã Setting knob When pulled. Â Red/black flag Appears when electrical supply is lost. Compass control should be place on DN for use. to $ 80 degrees. Á Aicraft Symbol Orange.20 P5 ATTITUDE HEADING REFERENCE SYSTEM 001 DEC 96 STANDBY HORIZON À Attitude Sphere Marked every 5 degrees of pitch axis. 20.10.FLIGHT INSTRUMENTS 1. 10. S See AHRS A/ERECT FAIL procedure in chapter 2.Nil - DC BUS 2 (on overhead panel AUX) .Nil - CAPT RMI DC BUS 2 (on overhead panel 28 VDC) 26 VAC BUS 2 (on overhead panel 26 VAC) F/O RMI DC STBY BUS (on overhead panel 28 VDC) 26 VAC BUS (on overhead panel 26 VAC) Standby horizon power supply DC ESS BUS (on overhead panel NORM STBY HORIZON) .12.Nil - Standby horizon aux power supply HOT EMER BAT BUS (on overhead panel AUX STBY HORIZON) .05.12.05.12. . S See EFIS COMP procedure in chapter 2.05.One AHRS loses TAS input from both ADC.Loss of AHRS S See AHRS FAIL procedure in chapter 2.FLIGHT INSTRUMENTS 1.Nil - DC BUS 2 (on overhead panel NORM) . .3 ELECTRICAL SUPPLY/SYSTEM MONITORING ELECTRICAL SUPPLY DC BUS SUPPLY (C/B) AC BUS SUPPLY (C/B) DC EMER BUS (on overhead panel NORM) .Nil - AHRS 2 aux power supply in flight DC BUS 1 (on overhead panel FLT) .Nil - EQUIPMENT AHRS 1 power supply AHRS 1 aux power supply AHRS 2 power supply SYSTEM MONITORING The following conditions are monitored by visual and aura alerts : .20 P6 001 DEC 96 ATTITUDE HEADING REFERENCE SYSTEM 20.AHRS disagree. .Nil - AHRS 2 aux power supply on ground DC EMER BUS (on overhead panel GND) . . . . . . . . . . 10. At power up.Outer knob (ADI DIM) is used to select EADI ON/OFF and to set brightness. all failure messages appear on EFIS. Depressing it enables a test of the EFIS system and radio altimeter : .Inner knob (DH TST) is used to set decision height from . . This pb is inoperative in composite mode. Automatic setting is also performed when ambient brighness changes. Groundspeed is displayed.30 P 10 EFIS 001 DEC 97 AA EFIS CONTROL PANEL (ECP) 1 FULL/ARC pb Repetitive action on this pb alternately selects FULL mode and ARC mode on EHSI. RA indication displays 100 ft on EADI.Outer knob (HSI DIM) is used to select EHSI ON/OFF and to set brightness. On OFF position. . the RA test provides the radar with altitude information which trigger undue GPWS alerts. On OFF position. . R R R CAUTION : In flight.10 to 990 ft. blue pointer disappears from EHSI. At power up. EFIS test is performed only on ground. . 6 N° 2 BRG (◊) selector To select green bearing pointer to VOR 2 or to ADF 2. Radio altimeter test is performed in flight as well as on ground. FULL mode is automatically displayed. 4 HSI/DIM/WX/DIM knobs .FLIGHT INSTRUMENTS 1. 5 N° 1 BRG (0) selector To select blue bearing pointer to VOR 1 or to ADF 1. and to set average brightness in relation to other traces. green pointer disappears from EHSI. 2 GSPD/TTG pb Repetitive actions on this pb alternately selects Groundspeed (GSPD) and Time to go (TTG) on EHSI display. 3 ADI/DIM/DH/TST knobs .Inner knob (WX DIM) is used to select ON/OFF weather radar traces. Automatic setting is also performed when ambient brighness changes. . . . . Nil - 26 VAC STBY BUS (on overhead panel RMI) SGU 1 NAV Reference CRS1/HDG panel RMI 2 SGU 2 power supply ALT switching indication .30 P 15/16 070 EFIS JUN 97 AA 30.Nil - DC STBY BUS (on overhead panel EFIS SG 1) CAPT EADI DC STBY BUS (on overhead panel EADI) .Nil DC BUS 2 (on overhead panel EFIS SG 2) F/O EADI DC BUS 2 (on overhead panel EADI) .Nil - F/O EHSI DC BUS 2 (on overhead panel EHSI) .10.Nil - 26 VAC BUS 2 (on overhead panel RMI) .Nil - SGU 2 NAV Reference ALT/CRS 2 panel CRS1/HDG panel (HDG 2 reference) RMI 1 Mod.Nil - CAPT EHSI DC STBY BUS (on overhead panel EHSI) .3 ELECTRICAL SUPPLY EQUIPMENT DC BUS SUPPLY (C/B) SGU 1 Power supply ALT switching indication AC BUS SUPPLY (C/B) .Nil - . : 4366 .FLIGHT INSTRUMENTS 1. . . . The FDAU also receives data from a Flight Data Entry Panel (FDEP) located on the pedestal. They can be energized by selecting ON the RCDR pushbutton.a Cockpit Voice Recorder. They should transmit a signal on 37. depending on version) are retained. CVR R All crew communications transmitted through the RCAU are recorded. The detection range is 3. The beacons actuate immediately after immersion. DFDR Various aircraft parameters are sent to a Flight Data Acquisition Unit (FDAU) which converts them into digital data. Each recorder is equipped with an underwater acoustic beacon which is used to locate the recorder in the event of an aircraft accident over the sea. located below the overhead panel.10.000 yards). DFDR. Cabin crew announcements are also recorded. The recorders are automatically energized as soon as the aircraft is on its own electrical supply and are switched OFF automatically TeN minutes after engines cut. In addition.50 P1 FLIGHT RECORDERS 001 SEP 04 AA 50. and . The data are recorded by the DFDR which stores them on a magnetic tape.5 kHz for 30 days.FLIGHT INSTRUMENTS 1. CVR.5 km (4.a Digital Flight Data Recorder. a CVR microphone. recorders are OFF until one engine is started. All recording may be erased by pressing ERASE pushbutton provided the aircraft is on ground and the parking break is set. acquires cockpit conversation and aural alerts for recording. . Only the last 30 minutes (or 120 minutes. The 25 last hours of flight are retained. When the aircraft is on external power. and deenergized by pushing the RESET pushbutton.1 DESCRIPTION The aircraft is equipped with : . • First sequence : hours and minutes .first left thumbwheel of Data entry panel must be on 9 position. or .the DFDR is failed.UPDATE pb depressed. reset the flight number on data entry panel. correction is taken into account and is displayed for 5 seconds.FLIGHT INSTRUMENTS 1. • Third sequence : year. the tape records are marked to identify a special event. month. Repeat first sequence and insert year. Note : Once data are inserted.10. • Second sequence : month and day Repeat first sequence and insert month and day. the display flashes . UPDATE pb Data displayed are updated as following : . day.insert hour and minutes on data entry panel . flight number and maintenance data. year. or .UPDATE pb depressed. Data entry panel Enables (through 4 thumbwheels) to insert different data : hour. The following sequence must be initiated during these 5 seconds. STATUS FDAU light Illuminates amber when the FDAU is failed. . STATUS SYST light Illuminates amber when : .2 CONTROLS FLIGHT DATA ENTRY PANEL (FDEP) 1 2 3 4 5 6 R R R Data display Date and time may be displayed and selected through the UPDATE pb 3 (successive pressures) and the Data entry panel 2 (except when 8 and 9 position of its first left thumbwheel is selected). Events pb When momentarily depressed.the DFDR or QAR (if installed) electrical power is lost.50 P2 FLIGHT RECORDERS 001 JUL 00 AA 50. minutes.QAR (if installed) 80% full. . 10.FLIGHT INSTRUMENTS 1.50 P4 FLIGHT RECORDERS AA RECORD PANEL R Mod : 3420 or 3552 050 JUL 00 . . . . At the fuel outlet of each tank a fuel LP valve.11. When low level is reached in one tank. an electrical pump and a jet pump are installed. the electrical pump is only used to start the engine. In the feeder compartment. The jet pump is activated by HP fuel from the engine HMU and is controlled by a motive flow valve. In this case. It's possible to use only one fuel tank by switching off the opposite pump pb. jet pump takes over automatically. a blue FUEL X FEED" light comes on memo panel. After start. is installed. the electrical pump is automatically activated to supply the engine. Note : Each electrical pump is able to supply one engine in the whole flight envelope. . allowing control of an unbalance situation. each engine is supplied from its associated wing tank.1. R In normal operation. A crossfeed valve. Fuel flow/fuel used ind. When the crossfeed valve is open. If jet pump pressure drops below 350 mbar (5 PSI). its electrical pump is automatically actuated (160 kg remaining fuel in the tank).10 FUEL SYSTEM P3 GENERAL 001 SEP 05 AA ENGINE FEED (See schematic p 13/14) In normal conditions. allows both engines to be fed from one side or one engine to be fed by either tank. allow the crew to monitor fuel consumption for each engine. the two electrical pumps are automatically actuated. controlled by an electrically operated actuator. controlled by the associated fire handle. Each tank is fitted with a 200 I feeder compartment always full of fuel protecting the engine feed system against negative or lateral load factors. . 5 PSI) pressure switch. flow bar illuminates green and crosses the system flow line. The valve is open. jet pump motive flow valve is controlled closed. CROSS LINE The valve is closed. IN LINE (pb pressed in) The flow bar illuminates green in line. X FEED pb Controls the operation of the fuel crossfeed valve. electrical pump is automatically switched off. released) electrical pump is deactivated. TANK FUEL TEMP. Each valve is controlled by its associated fire handle.FUEL SYSTEM 1. flow bar is extinguished. flow bars are extinguished.11. IN LINE Flow bar illuminates green.10 P5 GENERAL 550 JUN 97 AA 10. • jet pump motive flow valve is controlled open but will remain closed until a sufficient pressure is available. Both electrical pumps are automatically actuated. OFF It illuminates white. Note : During transient phases (opening or closing). FEED LO PR light The light illuminates amber and the CCAS is activated when the fuel delivery pressure drops below 300 mbar (4 PSI).b. OFF (p. (cf description p7) Mod. : PW127F . LP VALVE position ind. IND. : 4457 Eng. The position of the fuel LP valve is displayed.  30 seconds after HP fuel pressure is available and normal jet pump functioning is sensed by the 600 mbar (8. The valve is open. This indicates pump failure or fuel starvation. The valve is closed. PB pressed in :  When jet pump delivery low pressure is detected (engine not running or jet pump pressure drop) : • electrical pump is automatically activated. Note : During transient phases (opening or closing).2 CONTROLS FUEL PANEL 1 2 3 4 5 PUMP pb Controls electrical pump and jet pump motive flow valve in each tank. Permanent extinguishing of both bars indicates a valve fault. RUN illuminates green when electrical pump is activated. CROSS LINE (pb released) The flow bar illuminates green and crosses the system flow line. FF indication The mass fuel flow to the engine is indicated by a pointer on a scale graduated in kg/h X 100 FU counter On the digital read out. if the functioning is normal. FU reset knob The fuel used counter is reset to 0 by pulling associated ind.) may be tested by the overhead panel ANN LIGHT switch on TEST position. is provided for each engine. SC is heard. the corresponding electrical pump is automatically actuated. 1 R 2 3 A fuel flow/fuel used ind. Note : All the digits (on the FU counter as well as on the FUEL QTY ind.11. CCAS is activated. At the same time. Mod : 3596 .10 P6 GENERAL 050 JUL 98 AA FUEL QTY PANEL 1 2 3 FUEL QTY indications Fuel quantity in each tank is displayed in kg. fuel used is indicated in kg. display all 8's. in addition. Test pb Pressing the test button will check both measurement channels and. reset knob.FUEL SYSTEM 1. This value is computed by integration of the fuel flow parameter. FF/FU IND. LO LVL amber lights Each light illuminates amber and the CCAS is activated when quantity of the concerned display becomes lower than 160 kg . MC flashes amber. TANK FUEL TEMPERATURE INDICATOR R A temperature measuring device is installed in the left feeder compartment. Yellow sector : -54_C to 0_ C Green sector : 0_C to 50_C Yellow sector : 50_C to 57_C Red dash : -54_C and + 57_C FUEL CLOG LIGHT Light illuminates amber when fuel pressure loss in the corresponding HP pump fuel filter exceeds 45 PSI.10 FUEL SYSTEM P7 060 GENERAL AA SEP 04 X FEED ADVISORY LIGHT Illuminates blue on memo panel when the crossfeed valve is selected open. FUEL TEMP IND FUEL TEMP indication Fuel temperature is displayed.1.11. indicating that the filter is blocked and by-passed. Mod : 3973 or 4371 or 4457 . Temperature is displayed on the overhead panel. . . . : PW127F 500 JUN 97 .FUEL SYSTEM 1.11.10 P 11/12 GENERAL AA Eng. 5 SCHEMATIC Eng.11. : PW127F 500 JUN 97 .10 P 13/14 GENERAL AA 10.FUEL SYSTEM 1. . . . LO PR : The light illuminates amber and the CCAS is activated through the MFC when the associated pump delivered fluid pressure drops below 1500 PSI (103.HYDRAULIC SYSTEM 1. pb pressed in : pump is energized OFF : (pb released) pump is deactivated. XFEED pb Controls opening and closure of the crossfeed valve. LO PR the light illuminates amber and CCAS is activated when auxiliary pump outlet pressure is detected lower than 1500 PSI and functioning conditions are met. Both hydraulic circuits are separated.ACW blue pump pressure below 1500 PSI and. .10 P3 GENERAL 001 DEC 97 AA 10.12. OFF illuminates white. .at least one engine running OFF (pb released) . AUTO (pb pressed in) pump runs as soon as the following conditions are met : . Both hydraulic circuits are connected. The XFEED automatically closes. ON : (pb pressed in) crossfeed valve is selected open. ON lt illuminates white. .2 CONTROLS HYD PWR PANEL 1 2 3 R 4 5 Main pumps pbs Control activation/deactivation of ACW electric motor driven pumps.gear handle selected DOWN and.67 USgal).propeller brake released and. auxiliary pump is deactivated. OVHT lt The lt illuminates amber and the CCAS is activated when pump case drain line overheat is detected (T > 121° C/250° F) LO LEVEL lt The lt illuminates amber and the CCAS is activated when associated tank compartment fluid quantity drops below 2. pb released : crossfeed valve is closed.5 I (0. Auxiliary pump pb Controls operating mode of DC auxiliary pump.5 bars). . OFF lt illuminates white. . JUN 97 . . . . . . . . . reminding the crew of stall alarm threshold being lower in icing conditions. ICING flashes amber when ice accretion is detected and horns anti icing and/or airframe de icing are not selected ON.20 P3 ANTI ICING ADVISORY SYSTEM 180 JUL 00 AA 20. DE ICING INDICATOR Illuminates Flashes Blue on memo panel when the airframe deicing system is selected ON.ICE DET FAULT illuminates. : (3973 or 4371 or 4457) + 5008 . . .ICING amber light flashes on central panel (with associated warning) if system works correctly. Press and hold test button for 3 seconds. ICE DET PTT The push to test pb is used to check the ice detector correct operation. Blue on memo panel when the airframe deicing system is still selected ON five minutes after last ice accretion detection. loss of power supply).13. In this case. .ICING AOA lt illuminates green as soon as one horns anti icing Pb is selected ON. provided both horns anti icing are selected OFF. stall alarm threshold recovers the values defined for flight in normal conditions.ICE AND RAIN PROTECTION 1. Mod. (with associated central warnings) if an ice detector failure is detected.2 CONTROLS ICE DETECTOR PANEL 1 R R R R 2 3 ICE DET INDICATION LIGHT ICING illuminates steady amber when ice accretion is detected.ICING AOA lt can only be extinguished manually by depressing it. ICING AOA pb . FAULT illuminates amber when a system failure is detected (detector fault. provided both horns anti icing and airframe de icing are selected ON. . . 13.30 P2 ENGINE AND WING PROTECTION 001 JUL 01 AA TIME SEQUENCE DIAGRAM NORMAL MODE (PILOTED BY MFC) R R BEGINNING OF THE FOLLOWING SEQUENCE AT : . boots inflate according to a separate timer and MFC is totally by passed.ICE AND RAIN PROTECTION 1. .60 sec (FAST MODE (SAT > -20_C) -180 sec (SLOW MODE (SAT < -20_C) OVRD MODE (SEPARATED TIMER AND FAST MODE ONLY) R BEGINNING OF THE FOLLOWING SEQUENCE AT60 SEC (FAST MODE) Note : When de icing OVRD mode is selected. But airframe de-icing is never available.30 P3 ENGINE AND WING PROTECTION 080 SEP 04 AA 30.Air temperature upstream of the de-ice valves exceeds 230°C. Pb pressed in Normal operation. OFF (pb released) OFF light comes on white. : 4050 . ON (pb pressed in) Signal is sent to the MFC in order to initiate a de-icing cycle depending on MODE SEL pb.ICE AND RAIN PROTECTION 1. Both DE ICE and isolation valves are closed.Associated distribution valve output has been controlled closed but a downstream pressure is detected. AIRFRAME pb Controls the outputs A and B of both wings and stabilizers distribution valves.Air pressure downstream of the de-ice valves stays below 14 PSI for more than 10 seconds. However engine de-icing may be used (engine de-icing selected ON will open de-ice valve).13. . Both DE ICE and ISOLATION VALVES are open. : . Associated boots stay deflated. Mod. or . ON light illuminates blue. Pb released In normal operation.2 CONTROLS ENGINE/WING DE ICING PANEL 1 R 2 R AIRFRAME AIR BLEED pb Controls both de ice and isolation valves. The alert is inhibited when pb is released. FAULT The light illuminates amber and the CCAS is activated when : .Associated distribution valve output has been controlled open but no downstream pressure has been detected. FAULT The light illuminates amber and the CCAS is activated when inflation sequencing of airframe boots A or B is not correct. NORM (pb released) Normal operation OVRD (pb pressed in) The emergency de-icing activation is selected (timing cycle = 60 s). Mod.Associated distribution valve output has been controlled closed but a downstream pressure is detected. This position is used when the associated FAULT light illuminates. Also controls associated de-ice valve in closed position. ON (pb pressed in) De-ice valve is controlled open even if Airframe Airbleed is not selected ON. .30 P4 ENGINE AND WING PROTECTION 080 JUN 97 AA 3 ENGINE pbs Control de-ice valves.13.ICE AND RAIN PROTECTION 1.SLOW light illuminates blue. FAULT Light illuminates amber and CCAS is activated when : . FAULT The light illuminates amber when both MFC modules associated to air intake boots control fail resulting in an incorrect inflation sequencing. the light illuminates white and all de-icing lights extinguish. 5 DE-ICING OVERRIDE guarded pb Controls the emergency de-icing operation. ON light illuminates blue. 4 DE ICING MODE SEL pb Controls the selection of wings/engines boots inflation cycles when MAN is selected on MODE SEL AUTO pb 6 FAST (pb released) timing cycle = 60 s SLOW (pb pressed in) timing cycle = 180 s . and a signal is sent to the MFC in order to initiate a cycle. . or . Pb released Associated boots stay deflated. The control panel enables control of all double valves (ENG and AIR FRAME). after Airframe Airbleed FAULT and ENG FAULT. : 4050 .AIRFRAME AIRBLEED pb selected OFF and air temperature upstream of the de-ice valve exceeds 230°C.Associated distribution valve output has been controlled open but no downstream pressure has been detected. as well as the outputs A and B of respective engine distribution valves.Inflation sequencing of engine boots A or B is not correct. MAN illuminates white. MFC1B. In this case the FAST mode is automatically activated. ADC2 and MFC2B FAULT Illuminates amber and the CCAS is activated when MFC (1B or 2B) and/or ADC failure occurs. MAN (pb pressed in) The DE-ICING MODE SEL pb 4 is operative and allows the crew to select the appropriate timing cycle depending on SAT. The DE-ICING MODE SEL pb 4 is inoperative. The DE-ICING MODE SEL pb 4 is inoperative. The cycle selection is provided ADC1.ICE AND RAIN PROTECTION 1. Mod. : 4050 .13.30 P 4A ENGINE AND WING PROTECTION 080 JUL 01 AA 6 R R MODE SEL AUTO pb Pb released Normal operation (automatic operating mode). . . . : 4050 .13.09.09.05.30 P 7/8 ENGINE AND WING PROTECTION 080 JUL 01 AA SYSTEM MONITORING The following conditions are monitored by visual and aural alerts : . S See DE ICING MODE SEL FAULT procedure in chapter 2.05. .09. Discrepancy between outputs S See MODE SEL AUTO FAULT procedure in chapter 2.05.ICE AND RAIN PROTECTION 1. .05.MFC 1B or 2B and/or ADC failure. cycle beginning.Power loss on a horn anti-icing unit S See HORNS ANTI ICING FAULT procedure in chapter 2.Distribution valve output controlled open but no downstream pressure detected or controlled closed but downstream pressure detected. Mod.05. R .Boots do not operate following MFC failure or both boots A and B of the same engine are supplied 200 sec after eng cycle beginning or Boots A (B) of both engines are supplied while boots B (A) are not supplied 20 sec after eng.09.09. S See AIRFRAME AIR BLEED FAULT procedure in chapter 2. S See AIRFRAME DE ICING or ENG DE or ANTI-ICING FAULT procedure in chapter 2.LOW pressure in the de icing common air manifold (P < 14 PSI and t > 6s) or over temperature (T > 230_C) upstream the pressure regulating valve. . 13. TIME SEQUENCE DIAGRAM MODE SEL : NORMAL OPERATION LOW POWER CYCLE MODE SEL : ON HIGH POWER CYCLE Eng.ICE AND RAIN PROTECTION 1. the heat elements are electrically connected in three blades (every) other blade). On each propeller.40 P1 PROPELLER ANTI ICING 500 JUN 97 AA 40. Two modes are available and automatically selected depending on the temperature. : PW127F . The system is supplied with 115 ACW.1 DESCRIPTION Propeller anti icing is performed by resistors installed near the surface of the inboard sections of the blade leading edges. ICE AND RAIN PROTECTION 1.13.40 P2 PROPELLER ANTI ICING 550 JUN 97 AA 40.2 CONTROLS PROPELLER ANTI ICING PANEL 1 PROP pb Controls the respective propeller heating elements. ON 2 (pb pressed in), the heating units are supplied. The ON light illuminates blue. pb released The heating elements are not supplied. FAULT The light illuminates amber to indicate that at least one blade is not electrically supplied. ANTI-ICING MODE SEL pb Controls the duration of propeller anti icing cycles when MAN is selected on MODE SEL AUTO pb 3 . pb released LOW POWER cycle is selected. ON (pb pressed in) HIGH POWER cycle is selected. The ON lt illuminates blue. Note : • LOW POWER has to be selected when temperature is between 0°C (32°F) and - 10°C (14°F). • HIGH POWER has to be selected when temperature is between - 10°C (14°F) and - 30°C (- 22°F). • Below - 30°C (- 22°F) icing problems should be non existant (no supercooled water). Mod. : 4050 Eng. PW127F ICE AND RAIN PROTECTION 1.13.40 P 2A PROPELLER ANTI ICING 080 JUL 01 AA 3 R R R R MODE SEL AUTO pb (same pb as ENGINE AND WING PROTECTION) pb released Normal operation (automatic operating mode) The ANTI-ICING MODE SEL pb 2 is inoperative. The cycle selection is provided by ADC 1, MFC 1B, ADC 2 and MFC 2B FAULT Illuminates amber and the CCAS is activated when MFC (1B or 2B) and/or ADC failure occurs (see schematic 1.13.30 p 4A). The ANTI-ICING MODE SEL pb 2 is inoperative. In this case, the HIGH POWER CYCLE is automatically activated. MAN (pb pressed in) The ANTI-ICING MODE SEL pb 2 is operative and allows the crew to select the appropriate timing cycle depending on SAT. MAN illuminates white. Mod. : 4050 ICE AND RAIN PROTECTION 1.13.40 P3 080 PROPELLER ANTI ICING JUL 01 AA 40.3 ELECTRICAL SUPPLY/MFC LOGIC/SYSTEM MONITORING ELECTRICAL SUPPLY DC BUS SUPPLY (C/B) AC BUS SUPPLY (C/B) Propeller 1 anti-icing PWR - Nil - AC wild BUS 1 (on lateral panel PROP1 ANTI ICING PWR SPLY Propeller 2 anti-icing PWR - Nil - AC wild BUS 2 (on lateral panel PROP2 ANTI ICING PWR SPLY EQUIPMENT Prop anti-icing CTL and Ind DC EMER BUS (on lateral panel PROP CTL and IND) Note : Propeller anti icing is inhibited when Np is below 63%. MFC LOGIC See chapter 1.01. R SYSTEM MONITORING The following conditions are monitored by visual and aural alerts : - One or more blade heating unit (s) inoperative. • See PROP ANTI-ICING FAULT procedure in chapter 2.05.09. - MFC 1B or 2B and/or ADC failure, discrepancy between outputs. • See MODE SEL AUTO FAULT procedure in chapter 2.05.09. Mod. : 4050 70. 70.3 ELECTRICAL SUPPLY EQUIPMENT Captain wiper F/O wiper DC BUS SUPPLY (C/B) DC ESS BUS (on lateral panel CAPT) DC BUS 2 (on lateral panel F/O) . FAST wiper operates at 130 cycles/mn.1 DESCRIPTION Rain removal from front windshields is provided by two wipers : each wiper is driven by a two speed electric motor. They are controlled by two WIPER selectors on the overhead panel : one for the Captain.2 CONTROLS R R R R R R R R R WIPER rotary selector Controls the windshield wiper on the associated side. Maximum speed to operate the wipers is 160 kt. and one for the F/O. OFF wiper operation stops at the end-of-travel (Park) position. SLOW wiper operates at 80 cycles/mn.ICE AND RAIN PROTECTION 1.70 P1 001 RAIN PROTECTION JUL 98 AA 70.13. . . . . Note 2 As soon as the gear is locked in the selected position. The door is therefore operated by the gear during retraction and extension.Gear retraction hydraulic line. Unlocking is hydraulically achieved. This valve supplies hydraulic pressure (green system) to : . . Uplocking is mechanically achieved. The doors are actuated mechanically by the gear itself. Nose landing gear is assisted by a mechanical device. The two forward doors will be closed after gear extension while the two aft will remain open. Each main gear incorporates a door mechanism linked to it. Mod. The system is controlled from the flight compartment by means of a push/pull handle which permits landing gear mechanical unlocking. The extension line is then connected to tank return for retraction.20 P3 LANDING GEAR 070 SEP 05 AA GEAR NORMAL OPERATION Landing gear extension and retraction is performed by a control lever located on the center instrument panel.Gear extension hydraulic line. Locking springs act as secondary alignment and ensure locking independently of hydraulic pressure availability.14. R Landing gear can not be retracted as long as at least one gear shock absorber senses weight on wheels. GEAR EMERGENCY EXTENSION In the event of normal system failure the landing gear can be extended mechanically. The landing gear extends due to gravity and aerodynamic forces. hydraulic pressure is released from the connecting line.LANDING GEAR 1. Unlocking is hydraulically achieved. Down locking is achieved by means of a dual aligment folding side brace. Main landing gear extension is assisted by a gas actuator. Note 1 The main gear wheels are automatically braked as soon as the lever is selected up. The retraction line is then connected to tank return for extension. : 3986 . The MFC electrically controls the landing gear selector valve located in the LH main landing gear fairing. Four doors close off the nose gear well and restore the fuselage profile. . . . FLT The systems are forced to the in"FLIGHT" position. Mod.20 P7 LATERAL MAINTENANCE PANEL 080 JUN 97 AA 20.W. : 4457 . a selector enables the weight on wheels systems to be selected to the in Flight" position when on the ground. for maintenance purposes.LANDING GEAR 1.14. SELECTOR On the RH maintenance panel.O. FAILURES READOUT DISPLAY The right side maintenance panel includes a readout display for failures of systems linked to the MFC. Landing gear malfunctions are indicated when the rotary selector is selected on the WOW/LDG position.4 LATERAL MAINTENANCE PANEL W. Mode Selector Controls the overriding of weight on wheels system. MAINT PNL" light illuminates amber on the CAP. NORM The system works normally. . . . . . . . . . . . 15.3 60.15.15.15.4 GLOBAL NAVIGATION SATELLITE SYSTEM (GNSS) DESCRIPTION CONTROLS ELECTRICAL SUPPLY SCHEMATIC 1.1 30.1 10.60 60.30 30.3 VOR/ILS/MKR/DME SYSTEM DESCRIPTION CONTROLS ELECTRICAL SUPPLY 1.4 GROUND PROXIMITY WARNING SYSTEM DESCRIPTION CONTROLS ELECTRICAL SUPPLY/SYSTEM MONITORING SCHEMATIC 1.2 20.3 40.2 60.20 20.15.2 40.NAVIGATION SYSTEM 1.15.15.00 CONTENTS 1.1 20.15.1 60.15.2 50.2 10.2 RADIO ALTIMETER DESCRIPTION ELECTRICAL SUPPLY 1.3 WEATHER RADAR DESCRIPTION CONTROLS ELECTRICAL SUPPLY 1.1 40.65 OMEGA (if installed) R Mod : 4654 or 4885 or 5020 090 JUL 99 .40 40.00 P1 CONTENTS AA 1.1 50.3 ADF SYSTEM DESCRIPTION CONTROLS ELECTRICAL SUPPLY 1.10 10.50 50. . . . . . . . . . . 20.) 10 Light sensor Automatically adjusts the display brightness. the knobs directly act on the active frequency.. 9 Annunciators MEM (Memory) : illuminates when a preset frequency is being displayed in the lower window. frequency select knobs(2) change the preset frequency display. A second push on the button enables return to normal operation. RMT (Remote) : illuminates when the ADF control box is being remotely controlled by an other system (FMS.NAVIGATION SYSTEM 1.3 ELECTRICAL SUPPLY R Mod : 0043 EQUIPMENT DC BUS SUPPLY (C/B) ADF 1 DC STBY BUS (on overhead panel ADF 1) ADF 2 (if installed) DC BUS 2 (on overhead panel ADF 2) .15.. NCS. window displays dashes. After a 2 second push on the ACT button the bottom. etc.20 P3 ADF SYSTEM 010 JUL 98 AA 7 ACT button In normal operation. 8 TEST button Is used to initiate the radio self test diagnostic routine. . . below glideslope .excessive descent rate . Mode 5 . Radio Altimeter.Enhanced modes : .40 NAVIGATION SYSTEM P 1 GPWS 080 SEP 02 AA 40. Terrain Awareness & Display (TAD) The system includes : .altitude loss after take-off . Mode 3 . the system requires data supply from ADC1.two DSP SEL" pushbuttons for display selection on EFIS To operate.unsafe terrain clearance .one EGPWS computer .15. Mode 2 . flaps position transmitter and gear lever position transmitter. Mode 6 . Mode 4 .two GPWS/GS" lights illuminated when alert is activated .Basic GPWS modes . Terrain Clearance Floor (TCF) .altitude callouts . AHRS1.one TERR" pushbutton dedicated to enhanced mode .excessive terrain closure rate . The system provides SGU1 & 2 with terrain data to perform display on EFIS. Mod : 5313 . GNSS.1 DESCRIPTION (See schematic p 9/10) The Enhanced Ground Proximity Warning System (EGPWS) provides visual and aural alerts in case of dangerous flight path conditions which would result in inadvertent ground contact if maintained. Mode 1 :. ILS2. The EGPWS performs the following alert modes : .one FAULT/OFF" light and a selector dedicated to basic GPWS part .1. WX Radar. the PULL UP" voice alert is generated and the red GPWS" warning lights illuminate. If the aircraft penetrates the inner envelope. the SINK RATE" voice alert is generated and the red GPWS" warning lights illuminate. R Mod : 5313 or 5467 . This mode does not depend on the aircraft configuration.40 NAVIGATION SYSTEM P 2 GPWS 080 SEP 05 AA ALERT MODES MODE 1 .15.1.EXCESSIVE DESCENT RATE If the aircraft penetrates the outer envelope. the PULL UP" voice alert is now generated with the red GPWS" lights always illuminated.1. the TERRAIN . the TERRAIN .15. Note that the upper altitude limit is reduced to 1250 feet if Geometric Altitude is valid.TERRAIN" aural message is heard until the aircraft pressure altitude has increased by 300 feet of altitude or 45 seconds has elapsed. If the aircraft continues to penetrate the envelope. R Mod : 5313 or 5467 .40 NAVIGATION SYSTEM P 2A GPWS 080 SEP 05 AA MODE 2 .TERRAIN" voice alert is generated and the red GPWS" warning lights illuminate. When the warning conditions no longer exist.EXCESSIVE TERRAIN CLOSURE RATE S FLAPS NOT IN LANDING CONFIGURATION When the aircraft penetrates the envelope. ALTITUDE LOSS AFTER TAKE-OFF When the aircraft penetrates the envelope. the TERRAIN -TERRAIN" voice alert is generated and the red GPWS" warning lights illuminate. R Mod : 5313 or 5467 .1. MODE 3 .40 NAVIGATION SYSTEM P 3 GPWS AA 080 SEP 05 S FLAPS IN LANDING CONFIGURATION When the aircraft penetrates the envelope. the DON'T SINK" voice alert is generated and the red GPWS" warning lights illuminate.15. When the aircraft penetrates the envelope at a speed higher than 190 kts with gear down and locked. the TOO LOW GEAR" voice alert is generated and the red GPWS" warning lights illuminate.15. R Mod : 5313 or 5467 . Note that the TOO LOW TERRAIN" warning area upper limit is reduced to 500 feet if Geometric Altitude is valid.1. If penetration is performed at a speed lower than 190 kts with gear not down. the TOO LOW TERRAIN" voice alert is generated and the red GPWS" warning lights illuminate.40 NAVIGATION SYSTEM P 4 GPWS 080 SEP 05 AA MODE 4 .UNSAFE TERRAIN CLEARANCE S GEAR UP This mode is active during cruise and approach with gear not in landing configuration. the TOO LOW FLAPS" voice alert is generated and the red GPWS" warning lights illuminate. the TOO LOW TERRAIN" voice alert is generated and the red GPWS" warning lights illuminate. The GPWS selector enables a landing with flaps not in landing configuration without incuring a warning.15. Note that the TOO LOW TERRAIN" warning area upper limit is reduced to 500 feet if Geometric Altitude is valid. R Mod : 5313 or 5467 . When the aircraft penetrates the envelope at a speed higher than 159 kts.1. if penetration is performed at a speed lower than 159 kts.40 NAVIGATION SYSTEM P 4A GPWS 080 SEP 05 AA S FLAPS UP This mode is active during cruise and approach with gear down and flaps not in landing configuration. the GLIDE SLOPE" voice alert is generated softly.15. R Mod : 5313 or 5467 . The mode automatically rearms by ascent above 2000 feet AGL or landing or selection of a non-ILS frequency.3 dots below the beam and penetrates the outer envelope . In both cases the amber G/S" caution lights illuminate. When the aircraft is more than 1. the TOO LOW TERRAIN" voice alert is generated and the red GPWS" warning lights illuminate. MODE 5 . the same voice alert repeats faster with an higher volume.40 NAVIGATION SYSTEM P 5 GPWS AA 080 SEP 05 S TAKE OFF This mode is active during take off with either gear and flaps not in landing configuration. If the inner envelope is penetrated. When the aircraft penetrates the envelope.BELOW GLIDESLOPE This mode is active during front course ILS approaches. These alerts can be inhibited below 2000 feet AGL if ILS is tuned by depressing one of the GPWS/GS" switches.1. A FIVE HUNDRED" foot callout is available during non-precision approach or if the aircraft is outside ± 2 dots glideslope deviation.1. Mod : 5313 .ALTITUDE CALLOUTS .40 NAVIGATION SYSTEM P 6 GPWS 080 SEP 02 AA MODE 6 . .MINIMUM" audio callout is generated when the aircraft passes with gear down through the selected decision height.15.A MINIMUM . A Runway Field Clearance Floor (RFCF) alert is also provided for runways that are located on top of a hill.15. This alert is similar to the TCF alert but is based on height above runway. TCF is activated during takeoff. The alert is based on current aircraft location. nearest runway center point position and radio altitude. At the same time GPWS" red alert lamp are illuminated and remain on until the alert envelope is exited. R Mod : 5313 or 5467 . cruise and final approach and complement existing Mode 4 by providing an alert based on insufficient terrain clearance even when in landing configuration.40 NAVIGATION SYSTEM P 6A GPWS 080 SEP 05 AA TERRAIN CLEARANCE FLOOR The Terrain Clearance Floor (TCF) mode creates an increasing terrain clearance envelope around the airport runway directly related to the distance from the runway.1. The aural message Too Low Terrain" will occur once at the initial TCF envelope penetration and one time thereafter for each 20% degradation in radio altitude. and to provide aural alert and graphic displays of the conflicting terrain. which conflict with the Caution criteria. PULL UP" is generated with the red GPWS" lights illuminated on each instrument panel. it replaces the Weather Radar display and can be available to the flight crew at any time. S Terrain Display The terrain data can be displayed on the Electronic Flight Instrument System (EFIS). an aural message TERRAIN AHEAD. A discrete pop-up signal provided by EGPWS is used to automatically display on EFIS the detected threatening terrain with an auto-range of 10Nm whatever is the previous information displayed. S Terrain Alerting Caution and Warning envelopes below and ahead of the aircraft path are computed as a function of airspeed and flight path angle. aircraft altitude and a worldwide terrain database to predict potential conflicts between the aircraft flight path and the terrain. It the aircraft penetrates the Caution envelope boundary. The density and color being a function of how close the terrain is relative to aircraft altitude.40 NAVIGATION SYSTEM P 6B GPWS 080 SEP 02 AA TERRAIN AWARENESS This function sues aircraft geographic position provided by an aircraft GPS or an optional internal GPS card. The local terrain forward of the aircraft is depicted as variable density dot patterns in green. are shown in solid red on the Terrain Display.1. When the Terrain Display is present. are shown in solid yellow on the Terrain Display.15. an aural message TERRAIN AHEAD. Simultaneously. terrain areas. terrain areas. If the aircraft penetrates the Warning envelope boundary. which conflict with the Caution criteria. Simultaneously. yellow or red. TERRAIN AHEAD" is generated with the red GPWS" lights illuminated on each instrument panel. Mod : 5313 . Terrain Alerts are depicted by painting the threatening terrain as solid yellow or red. the terrain is displayed independently of aircraft altitude.40 NAVIGATION SYSTEM P 6C GPWS 080 SEP 05 AA An optional Peaks display adds additional density patterns and level threshold to the standard mode display.15. and solid red and yellow colors are unchanged with regard to the standard display. The terrain identified as water (0 feet MSL) is displayed as cyan dots. A single elevation number (high altitude) is displayed when the screen is all black or blue as a result of flying over water or relative flat terrain where there is no appreciable difference in terrain elevations. Two elevation numbers (in hundreds of feet above MSL) with the highest terrain on top and the lowest terrain under it are displayed with the corresponding colors to indicate the highest and lowest terrain currently being displayed. The Peaks display includes a solid green level to indicate the highest non-threatening terrain. At altitudes safely above all terrain for the display range chosen. The red and yellow dot patterns. Peaks terrain background display R Mod : 5313 or 5457 .1. both GPWS and TERR FAULT lights on left panel turn off .TERRAIN AHEAD PULL UP aural alert is broadcoast .2 . G/S illuminates amber as long as a mode 5 alert is activated.TCF .After 12 sec. GPWS illuminates red as long as any mode 1 . GPWS and G/S indications are integrated into the pbs.both GPWS and TERR FAULT lights illuminate on left panel . The illumination is accompanied by the voice alert for the particular mode.G/S amber lights on both pilot panel turn on .GLIDESLOPE aural alert is broadcast . Terrain Display test pattern disappears on each EFIS . The illumination is accompanied hy the voice alert for this particular mode.40 NAVIGATION SYSTEM P 7 GPWS 080 SEP 05 AA 40.GPWS red lights turn off . will inhibit the mode 5 alert (aural and visual) R Mod : 5313 or 5467 . On ground will perform the system test .GPWS red lights on both pilot panels turn on .GPWS FAULT light on CAP turns off .GPWS FAULT light on CAP turns on .GPWS red lights turn off .G/S FB The pbs on CAPT and F/O panels are identical and connected in parallel.TAD alert is activated.2 CONTROLS GPWS .TERR blue lights on DSP SEL pbs on both pilot panels turn on .4 .Terrain Display test pattern is displayed on each EFIS .15. below 2000 feet.PULL UP aural alert is broadcast .G/S amber lights turn off .1. Pressed .GPWS red lights on both pilot panels turn on .3 . OFF all basic GPWS modes inhibited.1. R FLAP Mode 4 alert caused by flap extension. Mod : 5313 or 5467 .15. at less than landing configuration is inhibited to avoid nuisance warnings in case of landing R OVRD R with reduced flap setting.40 NAVIGATION SYSTEM P 7A GPWS 080 SEP 05 AA GPWS SELECTOR The selector is guarded and wirelocked on NORM position. NORM all basic GPWS modes alerts are operative. TERR Guarded pb Pressed all Enhanced (TAD & TCF) modes are operative.15.1. R Mod : 5313 or 5467 . illuminates white when GPWS selector on OFF position. FAULT illuminates amber when some or all Enhanced modes are lost.40 NAVIGATION SYSTEM P 7B 080 GPWS SEP 05 AA GPWS Light FAULT OFF illuminates amber when some or all GPWS basic modes are lost. DSP SEL pb RDR TERR illuminates blue when Weather Radar display is selected on EFIS. OFF illiminated white when TERR pb is released to inhibit Enhanced modes. illuminates blue when Terrain display is selected on EFIS. S GPWS FAULT amber light illuminates on CAP .GNSS or AHRS1 loss S GPWS FAULT amber light illuminates on CAP S TERR FAULT amber light illuminates on the 2VU left panel R Mod : 5313 or 5467 .S GPWS FAULT amber light amber light illuminates on CAP S both GPWS and TERR FAULT amber lights illuminate on left panel.40 NAVIGATION SYSTEM P 8 080 GPWS SEP 05 AA 40.3 ELECTRICAL SUPPLY/SYSTEM MONITORING ELECTRICAL SUPPLY EQUIPMENT R EGPWS computer GPWS-G/S alert GPWS light TERR pb DC BUS SUPPLY (C/B) DC BUS 1 (on overhead panel EGPWS COMPTR & WARN) DSP SEL pb GPWS FAULT indication on CAP DC BUS 2 (on overhead panel) SYSTEM MONITORING The conditions are monitored by visual alerts : .S both GPWS and TERR FAULT amber lights illuminate on left panel.ADC1 loss . . .1.15.Radio altimeter loss S GPWS FAULT amber light illuminates on CAP and on the 2VU left panel .EGPWS computer internal failure or power supply loss. . 40 P 9/10 GPWS AA 40.15.NAVIGATION SYSTEM 1.4 SCHEMATIC R Mod : 5313 or 5467 080 SEP 05 . ANTENNA System is controlled with the weather radar control panel.50 P1 010 WEATHER RADAR JUL 00 AA 50. Mod : 5016 .NAVIGATION SYSTEM 1.15. Modes selected on this panel are displayed on the EHSI. Weather radar can also be used in MAP mode to display ground obstacles.1 DESCRIPTION Weather radar system is designed for weather impediments detection. up to 300 NM in a 45 degree sector on each side of aircraft path. in 4 colors according to the intensity of rain detected. Weather impediments are displayed on the EHSI(S) in ARC mode. it is possible to restore the active WX mode by pushing the STAB button four times in three seconds. displaying five different levels. GMAP position : (GMAP displayed green on the EFIS) Selects the ground mapping mode using four different levels. . WX position : (WX displayed green on the EFIS) Selects the weather detection mode.NAVIGATION SYSTEM 1. If needed. (TEST displayed on the EFIS) displays a test pattern to verify the system. OFF position : The radar system is turned off (WX displayed amber on the EFIS). Level 0 : Black No return Level 1 : Cyan Least reflective return Level 2 : Yellow Moderate return Level 3 : Magenta Strong return FP position : TST position : Mod : 5016 (FPLN displayed green on the EFIS). the system is automatically forced in SBY position for safety. The target alert mode (TGT) can be used in FP mode.15. SBY position : (STBY displayed green on the EFIS) Places the radar in a ready state with the antenna scan stopped and the transmitter inhibited. Level 0 : Black No detectable cloud Level 1 : Green Moderate storm Level 2 : Yellow Less severe storm Level 3 : Red Strong storm Level 4 : Magenta Intense storm On the ground. Clears the screen of radar data and put it in the flight plan mode.50 P2 WEATHER RADAR 010 JUL 00 AA 50.2 CONTROLS WEATHER RADAR CONTROL PANEL 1 Mode selector Enables the selection of the operating mode. TGT is displayed green on the EFIS and the system monitors beyond the selected range and 7. 5 RCT push button Activates or deactivates the REACT mode which compensates for attenuation of the radar signal as it passes through rain fall. When activated. The cyan field indicates areas where further compensation is not possible. calibrated gain mode. 6 STAB push button Turns the pitch and roll stability ON and OFF.50 P3 WEATHER RADAR 010 JUL 00 AA 2 TILT control Is used to adjust the antenna pitch from 15o down to 15o up. If a characteristic return is detected in the monitored area. the TGT legend on the EFIS changes from green to amber.NAVIGATION SYSTEM 1. additional ranges of 500 and 1000 miles are available. the system enters the preset. the rotary control does nothing. the system enters the variable gain mode.5o on each side of the aircraft heading. When the switch is pulled. TGT alert can only be selected in the WX and FP modes. In the FP mode. 3 RANGE push buttons Select the different operating ranges from 5 to 300 NM. 4 GAIN rotary control and push/pull switch When the switch is pushed. Mod : 5016 . in this mode. Any target detected in these areas will be displayed in magenta and should be considered dangerous.15. 7 TGT push button Activates and deactivates the radar target alert mode. 8 SECT push button Is used to select either the normal 12 looks/mn 120o scan or the faster update 24 looks/mn 60o sector scan. adjustable by the rotary control (VAR is displayed amber on the EFIS). • G MAP (green) : enables to display the ground obstacles • RCT (green) : Rain Echo Attenuation Compensation Technique mode is engaged.50 P4 WEATHER RADAR 010 JUL 00 AA EHSI DISPLAY 1 wx 1 Radar Mode Annunciator Appears as soon as the weather radar is switched on. • WAIT (green) : corresponds to the radar unit warning up time (is illuminated during 90 seconds) • STBY (green) : radar is in STBY mode. 2 Distance indication (white) Appears as soon as ARC mode is selected on EHSI.illuminates green when radar is operative . . • TEST (green) : illuminates green when a test is initiated • WX : . 3 Weather indication Is displayed in four colors.illuminates amber to indicate that radar is not working when Ąit is either : . or selected ON on both ECP and Radar Control box but antenna is not scanning.NAVIGATION SYSTEM 1. • TX (magenta) : illuminates when radar operates with the screen dimmed. Mod : 5016 .15. selected ON on EFIS Control Panel (ECP) but OFF on Radar Control box. 15.50 P5 WEATHER RADAR AA LEFT INTENTIONALLY BLANK Mod : 5016 010 JUL 00 .NAVIGATION SYSTEM 1. 15.3 ELECTRICAL SUPPLY Mod : 5016 EQUIPMENT DC BUS SUPPLY (C/B) AC BUS SUPPLY (C/B) Weather radar DC BUS 1 26 VAC STBY BUS (on overhead panel 28 VDC) (on overhead panel 26 VAC) .50 P6 010 WEATHER RADAR SEP 03 AA 50.NAVIGATION SYSTEM 1. the dead reckoning mode (DR) is used like a back-up utilizing true airspeed..) .Non precision approach To know all the functions available.Flight plan navigation . longitude. GNSS is an automatic tridimensional (latitude."Direct To" navigation . NDB.. It mainly allows to perform: . The data base is stored in the NPU and is updated every 28 days on the ground using a specific data loader. FUNCTIONS HT 1000 is capable of performing all the functions associated with the great circle navigation.15.1 DESCRIPTION (See schematic P.Navigation to nearest airport (or nearest VOR.Vertical navigation (non-coupled to auto-pilot) .NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM GNSS 1.60 P1 130 JUL 01 AA 60. heading and the last computed wind data. refer to the table of contents of the HT 1000 pilot's guide. In the case where the GPS position becomes unavailable. altitude) location and navigation means. 11/12) Using information provided by a constellation of 24 satellites (the HT 1000 is able to track up to 12 satellites at a time). The navigation is normally performed using the GPS sensor (GPS mode). Mod : 5176 + (4839 or 4656) . The effective date periods are displayed on the MCDU IDENT page. It also uses data recorded in a data base. NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM GNSS 1.2 CONTROLS HT 1000 CONTROLS The MCDU is the pilot interface for operation and data entry of the HT 1000 and also displays routes and advisory data on a color 5.5" liquid crystal display. The display has 14 lines of data with 24 characters per line.60 P2 090 JUL 00 AA 60. R Mod : 4654 or 4885 or 5020 or 5176 .15. The MCDU keyboards provides for data input and display selection and control. RNV pb Action on this pushbutton causes selection of the GNSS mode. Refer to 1. In MAP selection. waypoints of the flight plan are displayed in white except the active waypoint which is magenta.15.10. V/L pb Action on this pushbutton causes selection of the VOR/LOC mode.NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM GNSS 1. R Mod : 4654 or 4885 or 5020 or 5176 .60 P3 090 JUL 00 AA EFIS CONTROL PANEL 1 2 3 4 Map pb Repetitive action on this pb selects alternately MAP display and ARC display on EHSI.30. Waypoints G next waypoint (magenta) G other waypoint (white) f indicates an airport indicates a VOR Mod : 5176 .60 GLOBAL NAVIGATION SATELLITE SYSTEM P4 110 GNSS JUL 00 AA EHSI .15.MAP DISPLAY V < 1 2 3 15 NAV source annunciation Identifies the source which supplies the EHSI.NAVIGATION SYSTEM 1. Distance counter Indicates the distance computed by GNSS to the next waypoint. It becomes amber when both Pilot and F/O use GNSS as navigation source. This information is blue when only one crew member uses GNSS. VOR/RNV 1 is displayed in blue.15. TO/FROM indicator (magenta) 10 Drift angle indicator (magenta) 11 Radar's range selector may be used to select the distance scale. WPT/DGR alerting WPT illuminates amber when approaching a waypoint DGR illuminates amber when the UNABLE RNP" message is displayed on the MCDU.NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM 1. OFS (cyan) Indicates that a parallel offset has been activated. RADAR status 13 14 15 Vertical deviation Scale and Index Mod : 5176 . Wind indicator An arrow and a figure indicates the direction and the velocity (in kt) of the wind. VOR/RNV 2 is displayed in green.60 P5 GNSS 110 JUL 00 AA 4 5 6 7 8 9 Ground speed indicator Indicates the ground speed calculated by the GNSS. 12 DME 1/2 Indicates the distance given by DME 1/2. VOR/RNV symbols G RNV VOR Only VOR/DME or RNV waypoints will be presented on the MAP display. Track deviation Indicates in NM and tenths of NM the track deviation to the left (L) or to the right (R) of the intended track. TO/FROM annunciator 7 8 RADAR status 9 APP/OFS APP illuminates cyan when in approach phase.60 P6 GNSS 110 JUL 00 AA EHSI . Desired track indication 5 Lateral deviation to the track 6 WPT/DGR alerting WPT illuminates amber when approaching a waypoint DGR illuminates amber when the UNABLE RNP" message is displayed on the MCDU. It becomes amber when both Capt and F/O use GPS as navigation source. OFS illuminates cyan when an offset has been activated. Ground speed indicator Indicates the ground speed calculated by the GPS.ARC DISPLAY V 1 2 3 4 NAV source annunciation Identifies the source which supplies the EHSI. This information is blue when only one crew member uses GPS. Mod : 5176 .NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM 1.15. Distance counter Indicates the distance computed by GPS to the next waypoint. It becomes amber when both Capt and F/O use GPS as navigation source. Mod : 5176 . TO/FROM annunciator 7 8 RADAR status 9 APP/OFS APP illuminates cyan when in approach phase. This information is blue when only one crew member uses GPS.NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM 1. Distance counter Indicates the distance computed by GPS to the next waypoint.FULL MODE V 1 2 3 4 NAV source annunciation Identifies the source which supplies the EHSI. Desired track indication 5 Lateral deviation to the track 6 WPT/DGR alerting WPT illuminates amber when approaching a waypoint DGR illuminates amber when the UNABLE RNP" message is displayed on the MCDU. Ground speed indicator Indicates the ground speed calculated by the GPS.60 P7 GNSS 110 JUL 00 AA EHSI . OFS illuminates cyan when an offset has been activated.15. and RAIM prediction at FAF/MAP fails.DEAD RECKONING: message annunciated when NAV source becomes dead reckoning (GPS and DME modes are lost).15. VERTICAL TRACK CHANGE ALERT This message and annunciator is displayed prior to reaching a vertical track change. (1) RNV MSG is triggered by one of the following messages displayed on the MCDU: . or navigation source is not GPS. Mod : 5176 .VERIFY RNP-POS REF: message annunciated when flight phase changes and current pilot entered RNP is greater than the default RNP for new flight mode.VERIFY RNP ENTRY: message annunciated when the pilot entered RNP is greater than default current RNP.60 P8 110 GNSS JUL 00 AA EADI Alerting messages and displays related to the HT 1000 operation. . END OF DESCENT This message appears whenever the aircraft reaches the last altitude constraint on the descent path. .UNABLE RNP: message annunciated when ANP exceeds RNP or integrity is greater than twice the selected RNP (Once the value for approach). The conditions for displaying this message are flight phase dependent as follows: PHASE OF FLIGHT DEFAULT RNP (NM) Unable RNP Time to Alarm (s) OCEANIC 12 80 EN ROUTE 2 80 TERMINAL 1 60 APPROACH * 0.NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM 1. .3 10 * Not applicable .UNABLE APPROACH: message annunciated when within 2 NM from the FAF. RAIM LIMIT EXCEEDS XX NM: message annunciated when the GPS RAIM protection Limit exceeds TSO-C129 requirement for current flight phase.60 GLOBAL NAVIGATION SATELLITE SYSTEM P9 GNSS 270 JUL 00 AA In addition there are some advisory messages such as: . .3 ELECTRICAL SUPPLY Mod : 1603 + 5176 EQUIPMENT DC BUS SUPPLY GNSS DC STBY BUS (on overhead panel GPS) . RNV MSG will extinguish when the associated MCDU message is cancelled.NAVIGATION SYSTEM 1. 60.15.CHECK DEST RAIM-POS REF: message annunciated within 30 NM of destination airport if active route contains approach and approach RAIM predicted to be non available for some period of time within 15 minutes of destination ETA. 60 P 10 090 JUL 00 .15.NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM GNSS AA LEFT INTENTIONALLY BLANK R Mod : 4654 or 4885 or 5020 or 5176 1. 60 P 11/12 110 JUL 01 .4 SCHEMATIC Mod : (4654 or 4885 or 5020 or 5176) + (4839 or 4656) 1.15.NAVIGATION SYSTEM GLOBAL NAVIGATION SATELLITE SYSTEM GNSS AA 60. . one driven by the HP spool. Propeller regulation is electronically controlled. take-off rating will be 2475 SHP with an automatic power increase to 2750 SHP (reserve take-off rating RTO) in case of other engine failure.10 P1 GENERAL 550 JUN 97 AA The engine is a Pratt & Whitney of Canada PW 127 F certified for a 2750 SHP max take-off rating . Power setting is characterized by constant power lever and condition lever positions.POWER PLANT 1.16. : PW127F .9 ft) : clockwise (looking forward) : 1200 RPM : 180 kg The engine accessories are mounted on two accessory gear boxes. Mod : 4457 Eng. The propeller is an Hamilton Standard 568 F - Diameter Rotation 100 % Np Weight: : 3.93 m (12. in normal operation. However. and one by the propeller reduction gear box. The power adapted to the flight phase is selected by the pilot through a power management selector. The engine comprises two spool gas generators driving a six blade propeller via a free turbine/concentric shaft/reduction gear box assembly. . Control of the secondary airflow is achieved by automatic oil cooler flaps positioning 4 . The other components are actuated through the reduction gear box. . : PW127F . and a secondary flow directed to the oil cooler 3 . The curvature 2 is intended to provide inertial separation and protection in the event of foreign object ingestion. The speed reduction is obtained in two stages. its activation by the CL is inhibited. AIR INLET As presented on figure.20 P2 SYSTEMS DESCRIPTION 550 JUN 97 AA 9 Propeller Reduction Gear Box The power turbine shaft is connected to the propeller reduction gear box by a coupling driveshaft flexible diagram connections at each end. On ground. .The fuel cooled oil cooler (FCOC) Note : Auxiliary feather pump is driven electrically.The ACW generator .16.POWER PLANT 1.The (HP) pump and overspeed governor . On the reduction gear box are installed.The propeller Valve Module (PVM) controlled by Propeller Electronic Control (PEC). It is also used to divide airflow in a primary flow directed to the engine.The auxiliary feather pump .The propeller brake (on RH engine only) . Mod : 4457 Eng. The gear box is mounted offset of the centerline of the turbo machine. the engine air intake 1 is offset and is a shallow "S" bend designed to provide uniform inlet flow to the compressor. 2 FUEL SYSTEM The fuel supplied from the A/C fuel tank flows through : . A fuel heater outlet temperature indication is provided. .A Fuel Cooled Oil Cooler (FCOC) 7 .16. and the fuel temperature is thermostatically controlled.POWER PLANT 1.a screen 1 with a by-pass capability.A fuel heater which includes : . a clogging indicator is provided on pilot's panel. . the excess being returned to HP pump inlet.A HP pump 3 with a filter.A flow divider 8 to the fuel nozzles. . . .The Hydro Mechanical Unit (HMU) which has two functions : • to meter the fuel flow delivery to the engine by a metering valve assembly 4 . that provides cooling of the lubricating system by using fuel system as cooling source. The source of heat is engine oil.20 P3 SYSTEMS DESCRIPTION 001 JUL 98 AA R 20. . • to provide the HP motive flow required by the fuel tank jet pump through an engine valve 5 .A fuel flowmeter 6 . .a fuel heater element 2 . the reduction gear box and the propeller pitch change system. Quantity indication is checked by sight glass (or by a dipstick) on the side of the tank. 5 as well as the FCOC 6 .20 P4 SYSTEMS DESCRIPTION 550 JUL 00 AA 20.In flight starting using exciters A+B regardless of start selection.4 IGNITION SYSTEM R R Each engine is equipped with a high energy ignition system. exciter B. It consists of two engine mounted ignition exciters (A and B) powered by the DC ESS BUS and two spark igniters.16. 20. Note : Using exciter A or exciter B may allow to detect an hidden failure. Note : When EEC is deselected.POWER PLANT 1. RGB oil is also cooled in a fuel heater. Oil pressure is controlled by a regulating valve 7 . : PW127F . .EEC is deselected.3 LUBRICATION SYSTEM (See schematic P9/10) Synthetic oil specification MIL-L-23699. A pressure transducer 9 and a low pressure switch 10 are installed. for each engine.4 L tank 1 . Ignition cycle includes two phases. or exciters A+B (according to ENG START rotary selector position). A filter cap is provided on the tank. Mod : 4457 Eng. During 25 s.Oil Tank Oil is contained in a 14.Pressure system A gear pump 2 driven by the accessory gear box supplies oil through an air cooler 3 mounted in the nacelle and a filter 4 both fitted with by-passes in case of clogging. or . or . An oil temperature sensor is provided.CL is set on feather or fuel S/O position. . A single oil system serves the turbo machinery.On the failed engine in case of ATPCS sequence. This action is inhibited if : . A low temp. The engine ignition system provides ignition for : . the intensity is 5 to 6 sparks/s and then.Scavenge system Scavenging is blown down or gravity drained except for N° 6 and 7 bearing cavity and the reduction gear box. in case of NH drop below 60 % exciters A+B are automatically activated.On ground starting using exciter A. . on which gear pumps are used. or .NH drops below 30 %. the intensity becomes 1 spark/s. one for each ignition exciter. using the MAN IGN guarded push-button. . valve 8 is provided to eliminate damaging pressures surges on cold starts. excites A+B can manually be activated. In addition. Interface between flight deck and PEC is ensured by a Propeller Interface Unit (PIU) installed in the electronic rack.Low pitch protection -14° (Reverse) < βref < 78. .An Electro Hydraulic Valve (EHV) which meters the pitch change oil to the pitch change actuator and allows a normal feathering of the propeller.Overspeed. PVM The PVM is installed on the reduction gear box and allows : .Reversing . Pitch (B) change is hydromechanically controlled by a Propeller Valve Module (PVM). low pitch and back-up feathering functions.The basic speed set . The PVM is controlled by a Propeller Electronic Control (PEC) installed in each engine nacelle which provides the synchrophasing between the two propellers. Mod : 4457 Eng.A protection valve which is a part of overspeed. The system is protected against : .Feathering . : PW127F .5 PROPELLER (See schematic P11/12) The propeller is driven by a free power turbine by means of a reduction gear box. with the overspeed governor. the PWR MGT rotary selector and the power lever to activate the pitch change mechanism through the governors and associated equipments. to contain propeller overspeed. The propeller control system uses the condition lever.Low pitch angle in flight.A feather solenoid (EHV back-up).Beta scheduling . The RH Propeller Reduction Gear Box is provided with a brake to be used on the ground for Hotel mode operation.16. .5° (Feather) Additionally it is used.Synchrophasing .A Rotary Variable Differential Transducer (RVDT) which adjust and confirm PLA position. . .POWER PLANT 1.Hydraulic pressure loss. .20 P5 SYSTEMS DESCRIPTION 550 JUL 98 AA R 20. The PVM comprises : . Note : After a propeller braking or releasing sequence. The PEC detects. when gust lock is released and propeller brake is still engaged.The HP spool drives the DC generator. PROPELLER BRAKE The propeller brake is fitted on a countershaft on the RH engine reduction gearbox in order to stop the propeller (and the power turbine). will not imply propeller brake disengagement. PIU The PIU (one per PEC) is an electronic box located in the electronic rack that realizes the interface between the PEC and the cockpit for propeller speed selection. isolates and accomodates systems faults. : 4457 + (4237 or 4238) Eng.20 P6 SYSTEMS DESCRIPTION 570 JUN 97 AA PEC The PEC is a dual channel electronic box which provides closed loop control over the propeller pitch change system. prior to any propeller brake activation. However.16. Mod. associated with CRC. In the event of a failure of the primary channel. PROP BRK light will illuminate red on CAP. ENGAGEMENT LOGIC READY light must be illuminated. Propeller speed is calculated by the PEC through EEC (altitude and airspeed data) and Np sensors. and PEC fault signalisation logics. control of the propeller system will automatically be transfered to the back-up channel. .POWER PLANT 1. When the engine is running in Hotel mode : . : PW127F . READY light may remain illuminated for about 15 s.Bleed pressure is available downstream the HP compressor and supplies both packs. Loss of one of the above mentionned required conditions for engagement. 20 P 7/8 SYSTEMS DESCRIPTION AA 550 JUN 97 CROSS SECTION 20.16.POWER PLANT 1. : 4457 Eng. : PW127F .6 SCHEMATICS EXTERNAL VIEW (RIGHT) EXTERNAL VIEW (LEFT) Mod. 20 P 9/10 SYSTEMS DESCRIPTIONS 001 DEC 97 AA LUBRICATION SYSTEM R R R R R R R R R R R R R R R R R R R R R R R R R .16.POWER PLANT 1. : PW127F .16.POWER PLANT 1. : 4457 Eng.20 P 11/12 SYSTEMS DESCRIPTION 550 DEC 97 AA PROPELLER REGULATION SCHEMATIC R R R R R R R R R R Mod. 16. An engine Electronic Control (EEC) provides control of fuel flow in the HydroMechanical Unit (HMU).1 GENERAL The power control parameter is the torque : TQ = P (engine power) NP The maximum torque for each flight condition. Power Lever (PL) PWR MGT selector BLEED position EEC Power Control + Np min POWER LEVER HMU NH PVM Blade angle governing + reverse PVM Np max + Feathering HMU Engine shutdown (HP Fuel shutdown valve) PWR MGT Selector POWER LEVER (PL) CONDITION LEVER (CL) Mod : 4457 PEC Eng. is computed by the FDAU independently of the engine governing and displayed on the TQ indicator (Automatic BUG).POWER PLANT 1. .The power lever (PLA). . The HMU delivers a fuel flow which generates the NH compressor rotation speed. through a stepper motor in such a way as to control the torque in accordance with outside conditions and positions of : .30 P1 550 SYSTEMS OPERATING JUN 97 30. at the selected rating.The bleed valves. : PW127F .The power management selector (PWR MGT). ENGINE ELECTRONIC CONTROL (EEC) . : 4457 Eng.Delivers a motive flow to the fuel tank jet pump. in accordance with commands transmitted by the EEC.Commands a rotor speed in accordance with 2 laws (NH = f (PLA)) : .Includes a stepper motor which adjusts the flow controlled by the hydromechanical channel. 1 st law (called top) used when EEC is ON to protect NH overspeeds. . by controlling the stepper motor. . to obtain a predicted torque as a function of :  the power lever position  the PWR MGT selector position  flight conditions  the position status of the bleed air valves R Note : Operating line with EEC ON may be placed above or below the HMU BASE LAW depending on weather conditions Mod. 2 nd law (called base) used when EEC is OFF. .Ensures engine shutdown (HP fuel S/O).30 P2 SYSTEMS OPERATING 550 JUL 98 AA 30. .Regulates a given power. : PW127F .Performs fuel metering in steady state operation and protects the system in case of transients.2 ENGINE GOVERNING MAIN UNITS (HMU-EEC-PVM-PWR MGT) HYDROMECHANICAL UNIT (HMU) . .POWER PLANT 1.16. 16. according to the PWR MGT selection. on ground and at low power (see propeller governing). automatic uptrimmed take-off power to the valid engine (ATPCS) by responding to the signal generated by the Auto-Feather Unit (AFU) of the failed engine. Mod : 4457 Eng. . controls the propeller maximum speed Np. . so as to ensure correct LP compressor operation. : PW127F .Controls the modulated opening of the Handling Bleed Valve (HBV). in case of engine failure at take-off. .At high power. PROPELLER VALVE MODULE (PVM) .Delivers. PWR MGT SELECTOR LINE A : One engine out operation LINE B : Normal TO or MCT LINE C : CLB LINE D : CRZ Note : Sensible sector designed to allow fix throttle engine control.Ensures minimum propeller speed control.POWER PLANT 1.Controls propeller pitch at low power and when using reverse.Ensures low pitch through a solenoid (when PLA are below FI position).30 P3 SYSTEMS OPERATING 550 JUL 98 AA R . . In case of engine torque control failure. with the PL at a set point. Mod. .At low power (authority of engine torque control is gradually reduced to be cancelled out at FI). Thus. is controlled (with PLA constant) in all ambient conditions.In Hotel mode. According to the rating selected on the PWR MGT rotary selector. ensuring a constant power. .POWER PLANT 1.16. : 3973 or 4371 or 4457 .30 P4 SYSTEMS OPERATING 070 JUL 98 AA ENGINE POWER CONTROL LAWS R 1 TOP LAW (EEC ON) This is a TQ (PLA) control law. It is backed-up by an NH (PLA) law which becomes active : . the torque which is the engine control parameter. the EEC commands a determined engine power and therefore a torque value (for a given propeller speed). authorized for the rating considered (thermo dynamic limit). Example for : sea level.16. static conditions. bleed off. Mod : 4457 Eng. power is automatically reduced in such a way as to maintain the torque at the maximum value.POWER PLANT 1.30 P5 SYSTEMS OPERATING 550 JUN 97 AA When necessary. : PW127F . in normal T. ADC 1 immediately takes over from ADC 2 . The engine torque must match with the torque calculated by the FDAU. two events may occur : . This value must be crosschecked with temperature and altitude information. When TO is selected at the PWR MGT selector.10 % TQ position.the torque calculated by the FDAU corresponds to RTO . the engine torque will coincide with the FDAU torque (RTO).16. ADC SW FAULT light illuminates.POWER PLANT 1. In the event of engine failure and automatic uptrim. engine sensors immediately take over from ADC 1.30 P6 SYSTEMS OPERATING 550 JUN 97 AA PLA quadrant has TWO CLEARLY IDENTIFIED POSITIONS Position 1 WHITE MARK At this position marked by a notch the control system delivers max rated power corresponding to the mode selected.Engine regulation uses pitot and static data coming from EEC. : PW127F .If ADC 2 was selected. with the ATPCS armed : . Note : If the selected ADC electrical supply fails. . Manual BUG must be set at RTO . ADC SW FAULT light does not illuminate.If ADC 1 was selected. EEC data are elaborated either from the selected ADC (normal configuration) or from engine sensors and imposed data (emergency configuration). TO : P = 2475 SHP MCT : P = 2500 SHP CLB : P = 2192 SHP CRZ : P = 2132 SHP ENGINE REGULATION . except when TO is selected at the PWR MGT selector. Mod : 4457 Eng.O. configuration PL NOTCH PWR MGT TO ATPCS ARMED The FDAU BUG is positioned at RTO. . . according to the operating point position prior to EEC failure with respect to HMU base law).Handing bleed valve (HBV) is still monitored by the EEC deselected with a law function of NH instead of PLA. B NH changes to NH1 value (at that time a power increase or decrease can be noted. : PW127F .30 P8 SYSTEMS OPERATING 500 JUN 97 AA 3 BASE LAW (EEC OFF) The NH (function of PLA) base law is used when the EEC is deselected.Loss of torque regulation at constant power lever position (changes in ambient conditions will call for PLA adjustments to maintain maximum engine torque). Note :  Loss of the EEC has no effect on the two torque indications (digital and analogic) displayed . This mode of operation (REVERSION) features : .Loss of propeller underspeed control at low power (FUEL GOVERNING). .POWER PLANT 1.16. (REVERSION MODE). C NH follows the NH (PLA) schedule of the HMU base law. Refer to the schematic p 4 EEC Deselection Sequence A at time of EEC failure PLA = Plo (NOTCH) NH = NHo NH remains fixed at NHo value until either PL travel reaches 52° or NH reaches its overspeed protection. Eng. . This governing is available whether EEC is ON or OFF. : 4457 Eng.3 PROPELLER SPEED GOVERNING BLADE ANGLE GOVERNING This is the normal in flight governing mode. .8 %) CL is set in AUTO position.8 % and NP selected.POWER PLANT 1. : PW127F . depicting evolution of the propeller speed NP function of PLA (example given in MCT mode).when EEC is OFF. at a given NP) FUEL GOVERNING This is the ground governing mode at low speed and low power.30 P 10 SYSTEMS OPERATING 550 JUN 97 AA 30.when the propeller is in FEATHER position. PWR MGT selector commands NP propeller speed (through the PEC) PL commands power (and therefore TQ. The PVM adjusts the propeller pitch according to the power setting in such a way as to maintain a constant propeller speed NP. or in flight at low power and low speed. Mod. Note : This control mode is cancelled : . The EEC automatically increases the fuel flow so as to maintain a minimum propeller speed (N P = 70. .Control operation may be summarized through the graph below. TRANSITION MODE This is the intermediate mode between the two previous ones.16. The NP speed is comprised between 70. It only applies on ground. * The feather solenoid mounted on the PVM. : PW127F . This system enables to reduce the power normally used for take-off by an amount of about 10% below the power certified by the engine manufacturer. * The PL position (sw set to 49°).to the cockpit indicators (needles only).40 P1 CONTROLS 550 JUN 97 AA 40. * A test selector located on the pedestal.to the MFC which includes the autofeathering/uptrim logic functions. * The EEC which transmits a signal enabling the power to increase from TO to RTO (or a
NH signal during ATPCS test at ground idle). . Full ATPCS (i. ARMING CONDITIONS PWR MGT selector TO ATPCS pb ON Both PL above 49° Both torques above 46% Aircraft on ground uptrim and auto feathering functions armed PWR MGT selector TO ATPCS pb ON Both PL above 49° Both torques above 46% Aircraft in flight Mod : 4457 auto featherĆ ing function armed Eng.to the FDAU. and delivers the corresponding control signals to the feather solenoid. .e. COMPONENTS The ATPCS operates with the following components on each engine : * The Auto Feathering Unit (AFU) which is the main system element. uptrim and autofeather) is only available for take-off (see arming conditions below). * The feathering electric pump installed on the reduction gear box. to the feathering electrical pump and to the opposite EEC.POWER PLANT 1. It conditions the torque signal coming from the engine and provides the torque indication : .16.1 ATPCS GENERAL The propulsion unit includes an ATPCS (automatic take-off power control system) which provides in case of an engine failure during take-off the uptrimmed take-off power on the remaining engine combined with an automatic feathering of the failed engine. In the Cockpit : * The ATPCS pb on the cockpit center panel. This is favorable to engine/propeller life without affecting the take-off performance in case of an engine failure. inhibition of autofeather on the remaining engine . IT WILL BE NECESSARY TO SELECT PWR MGT to MCT position after relight in order to be able to UNFEATHER the propeller.feather solenoid activated .POWER PLANT 1. or .15 seconds period automatically disarms the mode. Eng. or .feathering electric pump energized .ATPCS Pb set to OFF. its cancellation can only result from either : .PWR MGT other than TO. Once the mode has been triggered. This feature enables to perform an acceleration stop without having autofeather in order to benefit from some reversing action on the failed engine.both PL retarded.15 s B B uptrim is triggered and bleed autofeather is activated valve is shut off on the on the affected engine remaining engine Y . In this case.16. : PW127F . but uptrim is energized on the remaining engine.15 seconds. CAUTION : If the engine is restarted.ARM light extinguishes t " B ARMED IN FLIGHT autofeather is activated on the affected engine Note : Nothing happens on the affected engine for 2. the throttle reduction occuring within 2.40 P2 CONTROLS 500 DEC 97 AA TRIGGERING CONDITIONS One torque below 18% SEQUENCE AFTER TRIGGER Time ATPCS R ARMED R ON GROUND Trigger 2. Eng.16. : PW127F .POWER PLANT 1.40 P3 CONTROLS 500 JUN 97 AA ATPCS SEQUENCE armed on ground A/FEATH FUNCTION Disarming conditions Note : During a normal flight (without engine failure) uptrim/autofeather will be disarmed after take-off when leaving the TO position on PWR MGT. the pilot can increase the power (if necessary) by pushing the PL up the RAMP (after GO AROUND position) to the FWD stop. the gust lock.2 POWER LEVERS (PL) PL is mechanically connected to the HMU and to the PVM through cables and rods. Note : On the ground. For take off acceleration the pilot will push PLs from GI to the TO position which is identified by a notch. Reverse sector is protected" by a spring rod : a force must be exercised by the pilot to position the PL into reverse sector. When the PL are on the MAX RATED TQ position.POWER PLANT 1. R At landing. CAUTION :in case of engine failure.40 P4 CONTROLS 550 SEP 05 AA 40. when engaged. POWER LEVER SWITCHES Mod :4457 Eng : PW127F . Then after flight idle gate automatic unlocking. prevents excessive PL in the forward traction sector angle. and eventually to reverse. Releasing this pull force will bring PL back to around GI.16. This lever controls the power plant thrust from Max rated TQ to max reverse. he will act on the triggers to reduce down to GI. the pilot will reduce PLs to FI. and therefore associated propeller drag as long as propeller is not feathered. the PL remains active controlling the pitch angle. . 100 % OVRD position sets manually Np MAX. . HP fuel shut off valves and propellers speed (NP). .from FTR to FUEL SO (and return).POWER PLANT 1.40 P6 CONTROLS 550 JUN 97 AA 40.4 CONDITION LEVERS (CL) They operate feathering control.AUTO position controls propeller speed through PWR MGT selector position. CONDITION LEVER SWITCHES Mod : 4457 Eng.16. It is necessary to act on a trigger located on the lever side to travel .from AUTO to FTR (and return). controlled by PVM when in blade angle governing propulsion mode. A red lt incorporated in the lever will illuminate if a fire is detected on the associated engine provided CL is not in FUEL SO position. : PW127F . . ) Note : In case of FDAU target failure associated with a pointer misfunctioning. scale to identify 115%. One of them sends a signal to the AFU which supplies the analogic torque ind. position where reserve T. : PW127F . If "000" is displayed. a wrong EEC is installed.O. 2 Pointer Actual torque is displayed. torque sensor is failed. an AFU failure may be suspected (see page 1). displays a manually selected torque target. 6 Test pb Allows to test the ind.40 P7 CONTROLS 500 JUN 97 AA 40. Note : A blue dot is provided on the ind.16.3%-Red dot : 120% 3 FDAU target Displays the maximum torque value computed by the FDAU depending on the PWR MGT selection (except on the T. 1 Digital counter Actual torque is displayed. Green sector : 0-100% Red mark : 100% Amber sector : 100-106% Red dashed radial : 106. The other one sends a signal to the EEC which supplies the electronic torque ind. EEC cannot control the HBV which is then closed. If LAB" is displayed. torque is displayed. 4 Manual target Controlled by the knob 5. During test. both counter and pointer will display 115%. Eng. 5 Knob Enables setting of target bug 4. (pointer). If "ĆĆĆ" is displayed.POWER PLANT 1. (digital counter).5 INDICATORS & CONTROL PANELS TORQUE IND Two sensing torque probes are located on the reduction gear box.O. 6-65% Green sector : 70. During test. 2 Pointer Actual NP is displayed.POWER PLANT 1. Note : A blue dot is provided on the ind.16. Eng. : PW127F .8-100% Red mark : 100% Red dot : 120% 3 Test pb Allows to test the ind. Amber sector : 41.40 P8 CONTROLS 500 JUN 97 NP IND 1 Digital counter Actual NP is displayed. both counter and pointer will display 115%. scale to identify 115%. Pointer Actual ITT is displayed. scale to identify 1150°C. Note : A blue dot is provided on the ind.POWER PLANT 1.40 P9 CONTROLS 001 JUL 98 AA ITT IND 1 R 2 3 4 Digital counter Actual ITT (T6) is displayed. Green sector : 300-765°C Red point + H : 715°C (Hotel mode) Amber sector : 765-800°C Red mark : 765°C (Temperature limit during normal take-off to be checked in chapter 2. Test pb Allows to test the ind. . During test.01) White/red mark : 800°C (Temperature limit in uptrim conditions) Red point : 840°C (Temperature limit for 20 sec) Red point + S : 950°C (Temperature limit for 5 sec for start) Alert lt Illuminates amber and the CCAS is activated when ITT > 800°C or 715°C in hotel mode. both counter and pointer will display 1150°C.16. Note : A blue dot is provided on the ind. OIL TEMP indication Actual oil temperature is displayed.7% Red mark : 102.40 P 10 CONTROLS 500 JUN 97 NH/NL IND Digital counter Actual NH is displayed. A separate pressure switch activates the CCAS at 40 PSI.16. Green sector : 55-65 PSI Amber sector : 40-55 PSI Red mark : 40 PSI Dashed white/red radial at 55 PSI OIL LOW PRESS It Illuminates red when OIL PRESS indication drops below 40 PSI. During test. 2 Pointer Actual NH is displayed.POWER PLANT 1.2% Red mark : 104. OIL IND 1 1 2 3 OIL PRESS indication Actual oil pressure is displayed.7% 3 Pointer Actual NL is displayed. both counter and pointer will display 115%. : PW127F . scale to identify 115%.2% 4 Test pb Allows to test the ind. Green sector : 45-125_C Amber sector : 125-140_C and below 0_C Red mark : 140_C Eng. Green sector : 62-104. Green sector : 62-102. . only the autofeather function is preselected. uptrim and autofeather function are preselected . Power is locked at its pre-failure value.40 P 12 CONTROLS 001 JUL 98 AA ENG 1/ENG 2 CONTROL PANEL 1 EEC pb Controls the EEC of the associated engine ON : (pb pressed in) EEC adjusts HMU action.16.If pressed in on ground. 2 ATPCS pb pb pressed in : .if pressed in in flight. ARM : Illuminates green when arming conditions are met (see P1) 3 UP TRIM light Illuminates green when the uptrim signal is sent to the associated engine at the beginning of ATPCS sequence. OFF lt illuminates white. OFF : (Pb released) Uptrim and autofeather functions are deselected. by controlling the stepper motor which lowers fuel flow ordered by HMU. R . Reversion to HMU base law is achieved by deselection of failed EEC. (See 1. The CCAS is activated in flight only. FAULT : Illuminates amber and the CCAS is activated when an EEC failure is detected.POWER PLANT 1. This light is illuminated during all ground operation below FI. 4 LO PITCH light Illuminates amber when the actual blade angle is lower than the normal FI blade angle. OFF : (pb released) The HMU controls only NH as a function of PL angle.30).16. the other engine start is accomplished as a "cross start" : initiated on Main Bat supply only. In case of starting. Mod : 4111 Eng. it will extinguish automatically when NH reaches 45% which "identifies" sequence end.starter remains engaged after 45% .40 P 13 CONTROLS 400 JUL 01 AA ENG START PANEL 1 ENG START rotary selector Selects the ignition mode and/or start sequences. Ignition is selected when fuel shut-off valve is open (controlled by CL) . The ON It illuminates white. the start is assisted by the opposite DC GEN from 10% NH (on ground only).on RH engine when the propeller brake is ON but the gust lock is not engaged. but the cross start does not operate normally. OFF START ABORT Ignition circuit is deenergized. START A Only ignition exciter A is supplied on ground. Note : There are three START positions.POWER PLANT 1. START B Only ignition exciter B is supplied on ground. FAULT Illuminates amber and the CCAS is activated if : .GCU fails during starting . Selects a start sequence. ON (pb pressed in) Initiates a sequence.16. If the DC GEN is connected to the network. Starting sequence is disarmed or interrupted. START pb Initiates the starting (or cranking) sequence of the related engine provided the ENG START selector is in one of the START positions (or CRANK). START A and B Both ignition exciters are supplied. Ignition is inhibited. : PW127F . Note : As soon as one engine is running and the associated DC GEN is connected to the main DC electrical network. the amber "X START FAULT" light illuminates on the main electricaL panel. CRANK START 2 R Enables engine cranking. starter and ignition are automatically deactivated when NH reaches 45%. . The system will automatically be transfered to the other channel. : PW127F . Therefore. 2 PEC "SGL CH" lts SGL CH lt illuminates amber when one channel of propeller electronic control is lost. Note : On ground. at each propeller unfeathering.40 P 15 CONTROLS 550 JUN 97 AA PWR MGT PANEL 1 PWR MGT rotary selector Made up of two independent parts (front and back). PIU and EEC with basic power requirements corresponding to the selected position. Mod : 4457 Eng. Provides FDAU. 3 PEC FAULT pbs FAULT Illuminates amber and CCAS is activated when the two channels are lost.16.For right engine with the front part of the selector. the correct working of back-up channel is confirmed. The failure will be indicated on the FDEP. LO PITCH protection is tested by the PEC and the back-up channel is used during 2 sec.For left engine with the back part of the selector.POWER PLANT 1. SGL CH illuminates during unfeathering then extinguishes. . . OFF (pb released) PEC is deactivated and Np is blocked at 102% whenever power is sufficient. . (Selected by EEC/PEC SEL switch). Engine TRIM test and LRU switches 3 Switch with two stable positions used to : . Mod : 4457 Eng. : PW127F . for maintenance purpose only. All buttons on this panel are to be used on ground only. several tests and control device are provided.Perform EEC or PEC trim to ensure that power delivered corresponds to PL position (can be performed with engine not running and PL in the notch) or LRU code failures. BUS ARINC switches Bus arinc function is tested. R R R R R R R R R R R R 1 EEC/PEC SEL switch 2 Used to select EEC or PEC depending on appropriate maintenance test on Engine TRIM switches or LRU (line replaceable unit) code failures.16.50 P1 LATERAL MAINTENANCE PANEL 550 DEC 97 AA On RH Maintenance panel. .POWER PLANT 1. Maintenance data appear on the FDEP.Test EEC or PEC channel. 50 P2 550 LATERAL MAINTENANCE PANEL JUN 97 AA 4 Prop Overspeed test switches Used to test hydraulic part of overspeed governor. on the affected engine. LOW PITCH light illuminates Mod : 4457 Eng. the PL low pitch protection switch and feather solenoid are tested.16. 5 Propeller Feather Pump test switch This switch with two stables positions enables to test the feathering pump. secondary low pitch solenoid is tested.With the test switch on PLA > FI position. : PW127F .POWER PLANT 1.With the test switch on PLA < FI position. A + B . 6 Propeller LOW PITCH test switches . . 2 nd OVSP threshold is tested at 106 % Np. Note : In both cases. For safety reasons. A First OVSPD threshold is tested at 102 % Np. this test is impossible in flight. . . 16. DC BUS 1 (on overhead panel SOL) .60 ELECTRICAL SUPPLY/ P3 550 MFC LOGIC/SYSTEM MONITORING DC BUS SUPPLY (C/B) EQUIPMENT Propeller brake Feather pump test DC ESS BUS (on lateral panel PWR SUPPLY CTL IND) DC SVCE BUS (on lateral panel FEATH PUMP TEST) PEC 1 2 (NORMAL) associated PVM and PIU DC EMER BUS (on overhead panel) PEC 1 DC ESS BUS (on overhead panel) 2 (BACK-UP) Idle gate .01. Mod : 4457 JUN 97 Eng.POWER PLANT 1. DC BUS 2 (on overhead panel CAUTION) MFC LOGIC See chapter 1. : PW127F . 160 JUN 97 R Mod : (3973 or 4371 or 4457) + (4237 or 4238) . 06 – ICING CONDITIONS Model : 212 A FEB 99 001 FEB 02 OCT 01 . 04 – POWER PLANT ENGINES PROPELLERS OIL SYSTEM STARTER FUEL SYSTEM 2 . 05 – SYSTEMS AIR .2 –00 LIMITATIONS PAGE : 1 AFM CONTENTS DGAC APPROVED 2 . 01 – GENERAL INTRODUCTION KINDS OF OPERATION MINIMUM FLIGHT CREW PERFORMANCE CONFIGURATION MAXIMUM OPERATING ALTITUDE MANEUVERING LIMIT LOAD FACTORS CONFIGURATION DEVIATION LIST NOISE CHARACTERISTICS MAXIMUM NUMBER OF PASSENGER SEATS 2 . 02 – WEIGHT AND LOADING WEIGHTS CENTER OF GRAVITY ENVELOPE LOADING 2 . 00 – CONTENTS 2 . 03 – AIRSPEED AND OPERATIONAL PARAMETERS AIRSPEEDS OPERATIONAL PARAMETERS 2 .PRESSURIZATION HYDRAULIC SYSTEM LANDING GEAR FLAPS AFCS INSTRUMENTS MARKING CARGO DOOR OPERATION ELECTRICAL MFC COMMUNICATIONS TCAS GPS CABIN LIGHTING EGPWS 2 . 01.01.01.05 2.01 2.01.2.08 2.01.04 2.01.09 R 2.01.01.02 2.01.01.03 2.01.00 2.10 C0NTENTS GENERAL WEIGHT AND LOADING AIRSPEED AND OPERATIONAL PARAMETERS POWER PLANT SYSTEMS TCAS (if installed) GPS (if installed) CABIN LIGHTING EGPWS (if installed) COCKPIT DOOR SECURITY SYSTEM (if installed) 001 SEP 03 .07 2.00 LIMITATIONS P 1 CONTENTS AA 2.01.06 2. . LIMITATIONS 2.01 P2 180 GENERAL JUN 97 AA MANEUVERING LIMIT LOAD FACTORS FLAPS RETRACTED = + 2.01. Model : 102-202-212-212A .1G FLAPS EXTENDED = + 2 TO 0 G GEAR DOWN = + 2 TO 0 G The corresponding positive accelerations limit the bank angle in turns and the severity of pull up maneuvers.5 TO . Other limitations such as that given by the emergency evacuation demonstration must be respected. CARGO DOOR OPERATION Do not operate cargo door with a cross wind component of more than 45 kt. DISPATCHIBILITY For dispatch in the event of equipment failure or missing equipment refer to MEL/CDL. MAXIMUM NUMBER OF PASSENGER SEATS 74 R as limited by emergency exits configuration. 59 inches).02 P1 520 WEIGHT AND LOADING JUL 00 AA DESIGN WEIGHT LIMITATIONS MAXIMUM WEIGHT KG LB TAXI 22 180 48 898 TAKE OFF 22 000 48 501 LANDING 21 850 48 170 ZERO FUEL 20 000 44 092 CENTER OF GRAVITY ENVELOPE The limits of center of gravity are given in percentage of the mean aerodynamic chord (MAC).99 inches) forward of the fuselage nose. R Mod : 4439 or 5150 Model : 212A . The distance from station 0 to reference chord leading edge is 13.01. landing gear extended.362 meters (92. Station O is located 2.303 meters long (90.LIMITATIONS 2. The MAC is 2.604 meters (535.67) inches. . 55 * MAXIMUM DESIGN MANEUVERING SPEED VA Full application of roll and yaw controls as well as maneuvers involving angles of attack near the stall should be confined to speeds below VA. * MAXIMUM FLAPS EXTENDED OPERATING SPEEDS VFE FLAPS 15 185 kt FLAPS 30 150 kt * MAXIMUM LANDING GEAR EXTENDED OPERATING SPEEDS VLE = 185 kt VLO RET= 160 kt VLO LOW= 170 kt * MAXIMUM ROUGH AIR SPEED VRA= 180 kt * MAXIMUM WIPER OPERATING SPEED VWO = 160 kt * MAXIMUM TIRE SPEED : 165 kt (Ground speed).LIMITATIONS 2. This limit must not be intentionnaly exceeded in any flight regime. VMO = 250 kt MMO = 0. . VA = 175 kt R CAUTION : Rapidly alternating large rudder applications in combination with large sideslip angles may result in structural failure at any speed.01.03 P1 AIRSPEED AND OPERATIONAL PARAMETERS 001 SEP 02 AA AIRSPEEDS * MAXIMUM OPERATING SPEED. 01.03 P2 AIRSPEED AND OPERATIONAL PARAMETERS AA ROFA-02-01-03-002-A500AA R STALL SPEEDS Ć VSR Eng.LIMITATIONS 2. : PW127F 500 DEC 97 . 01.LIMITATIONS 2.C) Eng. : PW127F .03 P3 AIRSPEED AND OPERATIONAL PARAMETERS 500 DEC 97 AA V1 LIMITED BY VMCG CAS (KT) V1 LIMITED BY VMCG (FLAPS 15) OUTSIDE AIR TEMPERATURE (DG. 01.03 P4 AIRSPEED AND OPERATIONAL PARAMETERS 500 JUN 97 AA ROFA–02–01–03–004–A500AA VMCA CAS (KT) VMCA (FLAPS 15) VMCL Flaps VMCL (CAS) 30 98 kt 15 98 kt Eng.LIMITATIONS 2. : PW127F . . 2 104.LIMITATIONS 2.54 min 800 106 (*****) 5s 20 s(1) (2) 120 120 840 106. (***) Time beyond 5 mn is linked to actual single engine operations only.3 %. (****) Up to 75 % NH only.04 P1 500 POWER PLANT AA JUL 01 ENGINES ENGINE PARAMETERS Operating limits with no unscheduled maintenance action required. (**) Value linked to 100 % NP.2 % during climb and 73.2 101 55 to 65 0 to 125 (3) TAKE OFF 5 mn 90 (**) (*) 101. (4) : see page 4. Beyond these limits.4 106. refer to 2.4 101 55 to 65 0 to125 (3) MAXIMUM CONTINUOUS NONE (***) 90.04 P 3 for detailed information. (1).2 101 55 to 65 0 to 125 (3) 40 mini (****) .40 to 125 (3) 55 to 65 125 (3) GROUND IDLE 66 mini HOTEL (4) MODE STARTING 715 5s 950 (2) OTHER TRANSIENT . POWER SETTING TIME LIMIT TQ (%) ITT (°C) NH (%) NL (%) NP (%) OIL PRESS (PSI) OIL TEMPERATURE (°C) RESERVE TAKE OFF 10 mn (***) 100 (**) 800 103.01. : PW127F . (2).01. TQ indication may exceed 100 % but not 106.8 108 140 20 mn During RESERVE TAKE OFF. (*****) Permissible for completion of flight provided TQ does not exceed 75.9 (**) 800 103.2 104. (*) ITT limite depends on outside air temperature.9 101. R Note : Flight with an engine running and the propeller feathered is not permitted. (3).13 % during cruise. refer to maintenance manual. Eng. LIMITATIONS 2.04 P2 POWER PLANT AA Note : Operation up to 106.3% torque is time unlimited when NP is below 94% Eng. : PW127F 500 JUN 97 .01. LIMITATIONS 2. : PW127F 500 JUN 97 .01.04 P3 POWER PLANT AA ROFA–02–01–04–003–A500AA ITT LIMITS Eng. Record in engine log book for maintenance. Oil temperature must be maintained above 71°C to ensure fuel anti-icing protection in absence of the low fuel temperature indication. refer to the propeller maintenance manual. R (3) .Determine and correct cause of overtemperature. (4) .Engine ground operations with crosswind between 5 and 20 kt should not exceed 58 % TQ. Eng. Note : Oil temperature must be maintained above 45°C to ensure inlet strut de-icing. GROUND OR FLIGHT If a propeller is involved in an overspeed or in an engine overtorque.Do not use engine 2 in HOTEL MODE without a qualified person (flight crew or maintenance) in the cockpit. IN FLIGHT OPERATION USE OF NP SETTING BELOW 82 % IN ICING CONDITIONS IS PROHIBITED ATR airplanes are protected against a positioning of power levers below the flight idle stops in flight by an IDLE GATE device. .Engine run up must be performed into the wind. . . .Record in engine log book for maintenance. 20 mn are authorized between 125°C and 140°C. (2) .Temperature up to 125°C is authorized without time limitation. Such positioning may lead to loss of airplane control or may result in an engine overspeed condition and consequent loss of engine power.LIMITATIONS 2. It is reminded that any attempt to override this protection is prohibited.01. : PW127F .Determine and correct cause of overtorque. PROPELLERS GROUND OPERATION .04 P4 POWER PLANT 500 JUL 98 AA (1) .Refer to ENG OIL Hl TEMP procedure. Aero Shell Turbine oil 500 .Acceptable fuels (refer to PWC SB 20004 to determine equivalent approved fuels).33 + 57 .26 .Mobil Jet oil 254 .01.Use of JP 4 is prohibited. MINIMUM FUEL TEMP (°C) FUELS FREEZING POINT (°C) Starting Operation MAXIMUM FUEL TEMP (°C) JET A JET A1 JET B .34 .50 .40 .LIMITATIONS 2.Mobil Jet oil II . FOLLOWED BY 4 MN OFF OIL SYSTEM Approved lubricating oils (from PWC SB 20001): .Royco Turbine oil 560 .50 .15 % per volume) : .Castrol 4000 . .38 . FUEL SYSTEM .MIL-I-27 686 D .34 .Royco Turbine oil 500 .Castrol 5000 . Eng : PW127-127F .Aero Shell Turbine oil 560 .48 .Ethylene Glycol Monomethyl Ether as defined in MIL-I-27 686 E.Exxon Turbo oil 2380 Mixing of different brands of oil or viscosities of oil is not recommended.Philips PFA 55 MB .04 P5 100 POWER PLANT SEP 05 AA STARTER 3 STARTS WITH A 1 MN 30 SEC MAXIMUM COMBINED STARTER RUNNING TIME.48 + 57 + 57 + 50 JP 5 .Approved anti icing additives (maximum concentration allowed : 0.34 .46 . 04 P6 POWER PLANT AA REFUELING MAXIMUM PRESSURE 3.ONE ELECTRICAL PUMP AND ASSOCIATED JET PUMP ARE ABLE TO SUPPLY BOTH ENGINES IN THE WHOLE FLIGHT ENVELOPE R . Eng.LIMITATIONS 2.01.ONE JET PUMP IS ABLE TO SUPPLY BOTH ENGINES IN THE WHOLE FLIGHT ENVELOPE. EXCEPT WHEN USING JET B.5 BARS (50 PSI) USABLE FUEL THE TOTAL QUANTITY OF FUEL USABLE IN EACH TANK IS 2500 KG (5510 LBS) NOTE : FUEL REMAINING IN THE TANKS WHEN QUANTITY INDICATORS SHOW ZERO IS NOT USABLE IN FLIGHT UNBALANCE MAXIMUM FUEL UNBALANCE : 730 kg (1609 lb) FEEDING . : PW127F 500 SEP 05 .EACH ELECTRIC PUMP IS ABLE TO SUPPLY ONE ENGINE IN THE WHOLE FLIGHT ENVELOPE . LIMITATIONS 2. flight level must be limited to FL 200 HYDRAULIC SYSTEM SPECIFICATION : HYJET IV OR SKYDROL LD 4 LANDING GEAR R TOWING WITH TOWBARLESS SYSTEM IS PROHIBITED DO NOT PERFORM PIVOTING (SHARP TURNS) ON A LANDING GEAR WITH FULLY BRAKED WHEELS EXCEPT IN CASE OF EMERGENCY MFC TAKE OFF WITH TWO OR MORE FAILED MFC MODULES IS PROHIBITED.PRESSURIZATION MAXIMUM DIFFERENTIAL PRESSURE MAXIMUM NEGATIVE DIFFERENTIAL PRESSURE MAXIMUM DIFFERENTIAL PRESSURE FOR LANDING MAXIMUM DIFFERENTIAL PRESSURE FOR OVBD VALVE FULL OPEN SELECTION MAXIMUM ALTITUDE FOR ONE BLEED OFF OPERATION 6.05 P1 030 SYSTEMS AA SEP 04 AIR .5 PSI 0.01. Mod : 1603 .35 PSI .0.35 PSI 1 PSI 20 000 ft ELECTRICAL SYSTEM SOURCE DC GEN INV ACW GEN TRU MAX LOAD 400 A 600 A 800 A 500 VA 575 VA 750 VA 20 KVA 30 KVA 40 KVA 60 A 90 A TIME LIMIT NONE 2 mn 8s NONE 30 mn 5 mn NONE 5 mn 5s NONE 5 mn SINGLE DC GEN OPERATION In flight : if OAT exceeds ISA + 25. the system protects the flight crew members during 15 min. an emergency descent from 25.000 ft. on ground = Cabin Temperature in flight Minimum bottle pressure required to cover a cabin depressurization at mid-time of the flight. In case of smoke emission.000 ft to 13.Ref.01.unusuable quantity .05 P2 SYSTEMS 100 JUL 01 AA OXYGEN R Reference temperature = Cabin Temperature or OAT whichever is higher.normal system leakage .LIMITATIONS 2.000 ft within less than 4 minutes and a flight continuation at an altitude below 13. Mod : 4411 . A 25 % pax oxygen consumption is assumed. Temp errors. Provision is made to cover : . Note : At dispatch the computed flight time after decompression should be at least 1/2 of estimated flight time to destination or flight time to the longest en route alternate which ever is higher. . 5- Prior to perform RA's climb or increase climb. safe horizontal separation must also be assured by visual means. and limited to the minimum required to comply with the RA. 4- Evasive maneuvring should be made with the autopilot disengaged. manually adjust CLs. 3- Compliance with TCAS resolution advisory is required unless the pilot considers it unsafe to do so. These displays and advisories are intended only for assistance in visually locating the traffic and lack the resolution necessary for use in evasive maneuvering. without visually sighting the traffic. Mod : 3074 or 3113 or 3625 or 8259 or 3832 or 5146 or 5103 . maneuvers which are in the opposite direction of the resolution advisory (TCAS RA) are extremely hazardous and are prohibited unless it is visually determined they are the only means to assure safe separation. Therefore.01. 2- The pilot must not initiate evasive maneuvers using information from the traffic display only or from a traffic advisory (TCAS TA) only. in order to avoid negating the effectiveness of a co-ordinated maneuvre by the intruder. the crew should select the appropriate engine power setting on the power MGT rotary selector and. if necessary. The pilot must promptly return to the previous ATC clearance when the TCAS CLEAR OF CONFLICT" voice message is announced.06 P1 050 TCAS SEP 03 AA TCAS LIMITATIONS The limitations in Part 2. a normal procedure of go-around should be followed including the appropriate power increase and configuration changes.LIMITATIONS 2. 2- However. WARNING : Non compliance with a crossing RA by one airplane may result in reduced vertical separation.01 are applicable with the addition of the following: R 1- TCAS operation is approved for use in VFR meteorological conditions (VMC) and IFR meteorological conditions (IMC). 6- When a climb or increase climb RA occurs with the airplane in the landing configuration or in the go-around phase. CAUTION : Once a non crossing RA has been issued the vertical speed should be accurately adjusted to comply with the RA. Conditions where this may occur include operations with a bank angle (wings level is assumed). and abnormal configurations such as landing gear not retracted or stick pusher/shaker failure. This table does not consider worst turboprop flight conditions especially operations using minimum operating airspeeds as are sometimes required (e. ATC constraints). leaving aircraft in landing configuration during climb RA on approach.06 P2 TCAS 050 SEP 02 AA TCAS (cont'd) 7- Because of the limited number of inputs to TCAS for determination of aircraft performance inhibits. R Mod : 3074 or 3113 or 3625 or 3832 or 5146 or 5103 or 8259 . with the aid of the ATC system. however it is not possible to do so.LIMITATIONS 2.01. In all cases.g. In these cases. obstacle clearance. WARNING : Priority must be granted to increasing airspeed when reaching stall warning even when this requires deviation from an RA command issued by the TCAS. NOTE : TCAS is viewed as a supplement to the pilot who. altitude and temperature combinations outside those noted below. The table below entitled Flight Envelope in which climb resolution advisory can be accomplished without stick pusher/shaker activation" outlines the parameters used in the development of the performance inhibits. weight. engine out operations. stall warning must be given precedence over climb RA commands. TCAS may command maneuvers which may significantly reduce stall margins or result in stall warning. has the primary responsibility for avoiding mid-air collisions. there are instances where inhibiting RAs would be appropriate. LIMITATIONS 2.06 P3 050 TCAS SEP 02 AA TCAS (cont'd) FLIGHT ENVELOPPE IN WHICH CLIMB RESOLUTION ADVISORY CAN BE ACHIEVED WITHOUT STICK PUSHER / SHAKER ACTIVATION FLIGHT REGIME WEIGHT ALTITUDE TEMP.13 VS1g Power for level flight increase to Max Continuous Up Up Long Range Cruise Higher of 1.01.3G to buffet onset Temperature range up to ISA + 27° .13 VS1g 15 UP 1.Enroute .Take off .13 VS1g Spin up to go around power during maneuver from power required for 3° Glide Slope Transition from 30 to 15 DN to Up VAPP +10 1.Approach and landing Wings Level Assumed Altitude range R Mod : 3074 or 3113 or 3625 or 3832 or 5146 or 5103 or 8259 0 0 0 25000 ft 6000 ft 7000 ft . POWER FLAPS GEAR Take off FAR25/JAR25 Take off 15 Spin up to go around power during maneuver from power for level flight AIRSPEED INITIAL MIN.13 VS1g if defined or buffet onset Climb limit Approach FAR25/JAR25 Climb limit Landing Transitioning to go around at RA FAR25/JAR25 Enroute Critical Wt/Alt Climb limit giving 1.51 VS1g 1. Up V2 + 20 1. 01.LIMITATIONS 2. R Mod : 3074 or 3113 or 3625 or 3832 or 5146 or 5103 or 8259 .06 P4 TCAS 050 SEP 02 AA TCAS (cont'd) 8- Inhibition schemes 8.1. Maneuvers based solely on information displayed on the traffic display are not authorized.Non icing conditions CONFIGURATION RA CLIMB RA INCREASE CLIMB FLAPS 0 AUTHORIZED AUTHORIZED FLAPS 15 TO AUTHORIZED INHIBITED FLAPS 15 APPROACH AUTHORIZED AUTHORIZED FLAPS 30 AUTHORIZED INHIBITED CONFIGURATION RA CLIMB RA INCREASE CLIMB FLAPS 0 Z < 18000 ft AUTHORIZED INHIBITED FLAPS 0 Z > 18000 ft INHIBITED INHIBITED FLAPS 15 TO AUTHORIZED INHIBITED FLAPS 15 APPROACH AUTHORIZED INHIBITED FLAPS 30 INHIBITED INHIBITED 8.1.Icing conditions NOTE 1 : NOTE 2 : Pilots are authorized to deviate from their current ATC clearance to the extent necessary to comply with a TCAS resolution advisory. 04 are applicable. EMERGENCY PROCEDURES The emergency procedures in Part 2. Mod : (3074 or 3113 or 3625 or 3832 or 5103 or 5146 or 8259) + 5205 .06 P5 TCAS 080 SEP 02 AA TCAS (cont'd) NORMAL PROCEDURES The normal procedures in Part 2.LIMITATIONS 2.01.03 are applicable.05 are applicable with the addition of the following: The TCAS must be turned TA ONLY in the following cases: -Engine out operations -Stick pusher/shaker failure -Flight with landing gear down The TCAS must be turned STBY in the following cases: -ATC request -LOSS OF RADIO ALTIMETER INFORMATION -Errors or differences between independant air data sources PERFORMANCES The performances in Part 3 are applicable. PROCEDURES FOLLOWING FAILURES The procedures following failures in Part 2. In addition.has been demonstrated to meet the requirements of JAA TGL n°2. . 3 . : 5176 . .complies with TSO C 129 and TSO C 115A. 2 .LIMITATIONS GLOBAL POSITIONING SYSTEM GPS 2. AC 20-138 and DGAC CRI S-9902.LIMITATIONS Compliance with the above regulations does not constitute an operational approval/authorization to conduct operations. en route.The HT 1000 pilot's guide must be available on board.This equipment is approved for use as : . .is installed in compliance with FAA AC 20-129. Aircraft operators must apply to their Authority for such an approval/ authorization.01.If the event of DGR alarm illumination the flight crew must cross-check the aircraft position using conventional means or must revert to an alternative means of navigation.PROCEDURES . Dispatch must not be made in the event of predicted continuous loss of RAIM of more than 5 minutes for any part of the intended flight. the availability of GPS integrity (RAIM) must be confirmed for the intended flight (route and time). .The system must be used with an updated active data base and the waypoints position must be cross-checked with official charts.The system must operate with HT 1000-060 software version or any later approved version. REV1 and FAA AC 20-138 and FAA Notice N8110-60. Mod. where the coupled DME option is not installed or if the coupled DME is not operative.supplemental navigation means.07 P1 270 JUL 00 AA GPS 1 . Conventional means must be permanently cross-checked during the approach. . (b) Traditional navigation equipment must be selected to available aids so as to allow immediate cross-checking or reversion in the event of loss of GPS navigation capability. in terminal area and for non precision approach operations until the missed approach point with respect of the MDA. NOTE : VDEV function must be permanently monitored. . AC 20-130A. the following procedures apply for B-RNAV operations : (a) during the pre-flight planning phase.primary navigation means for oceanic and remote operations when only one long range navigation system is required. . NOTE : Stand alone GPS approach is not approved. . .GENERAL The Honeywell/Trimble GNSS 1000 : .advisory VNAV means. LIMITATIONS 2.01.08 P1 CABIN LIGHTING 100 SEP 04 AA The general cabin illumination system must be used during not less than 15 minutes before each flight. R Mod : 5040 2.01.09 LIMITATIONS P 1 EGPWS 080 SEP 05 AA 1 EGPWS - Navigation is not to be predicated on the use of the terrain display. Note :The Terrain Display is intended to serve as a situational awareness tool only. It does not have the integrity, accuracy or fidelity on which to solely base decisions for terrain or obstacle avoidance. 2 - To avoid giving nuisance alerts, the predictive TAWS functions must be inhibited when landing at an airport that is not included in the airport database. R Mod : 5313 or 5467 2.01.10 LIMITATIONS P 1 COCKPIT DOOR SECURITY SYSTEM 200 SEP 04 AA COCKPIT DOOR OPERATION : NORMAL PROCEDURES This procedure should be applied, if local authorities require that the cockpit door remain closed throughout the entire flight Before Pushback or engine start COCKPIT DOOR LOCKING SYSTEM SW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON COCKPIT DOOR CLOSED . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK With the cockpit door toggle switch at CLOSE, the cockpit door is closed and locked. After Engine Start H If routine access is requested from the cabin : The buzzer sounds in the cockpit for at least 2 seconds Prior to unlocking the door, the flight crew should identify the person requesting entry H If entry is not authorized by the flight crew : DOOR TOGGLE SW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DENY Emergency access, the buzzer, and the Door Call Panel are inhibited for 3 minutes H If entry is authorized by the flight crew : DOOR TOGGLE SW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPEN The flight crew should pull the switch and maintain it in the OPEN position, until the cabin crew pulls the door open. Note : If the flight crew does not take any action after the routine cabin request, the cabin crew will be able to open the door by using the emergency access procedure. H If the emergency access is initiated from the cabin : The buzzer sounds continuously in the cockpit for 30 seconds, and the OPEN light flashes on the central pedestal's cockpit Door Control Panel. Note : If the flight crew does not take any action, the door will unlock after 30 seconds. H If entry is not authorized by the flight crew : DOOR TOGGLE SW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DENY Emergency access, the buzzer, and the Door Call Panel are inhibited for 3 minutes. When the situation in the cockpit permits, the flight crew should identify the person requesting entry prior to unlocking the door. H If entry is authorized by the flight crew : DOOR TOGGLE SW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPEN The flight crew should pull the switch and maintain it in the OPEN position, until the cabin crew pulls the door open. Before leaving the aircraft COCKPIT DOOR LOCKING SYSTEM SW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF FAULT LIGHT on Door Control Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CHECK R Mod : 5377 + (5434 or 8330 or 8333) 2.01.10 LIMITATIONS P 2 COCKPIT DOOR SECURITY SYSTEM 200 SEP 04 AA COCKPIT DOOR OPERATION : NORMAL PROCEDURES (cont'd) Daily Check Ć COCKPIT DOOR LOCKING SYSTEM SW . . . . . . . . . . . . . . . . . . . . . . . ON On the DOOR CALL PANEL Ć EMER PB . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PUSH Ć Check OPEN LIGHT flashes In the cockpit : Ć Check buzzer Ć Check OPEN LIGHT flashes H If correct : In the cockpit : Ć TOGGLE SW on COCKPIT DOOR CONTROL PANEL . . . . . . . . . . DENY Ć Check buzzer stops Ć Check OPEN LIGHT extinguishes On the DOOR CALL PANEL Ć Check OPEN LIGHT extinguishes and DENIED LIGHT illuminates H If correct : Ć TOGGLE SW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OPEN Ć COCKPIT DOOR LOCKING SYSTEM SW . . . . . . . . . . . . . . . . . . OFF Ć FUNCTIONAL CHECK OF THE MANUAL LOCK BOLT(S) R Mod : 5377 + (5434 or 8330 or 8333) 2.01.10 LIMITATIONS P 3 200 COCKPIT DOOR SECURITY SYSTEM SEP 04 AA COCKPIT DOOR OPERATION : EMERGENCY PROCEDURES Electrical Power Lost Ć Move the manual lock bolts to CLOSE position Ć Note : when the door is locked with the manual bolts, the emergency access to R R R R R R R the cockpit is unavailable. It is recommended that at least two crew members remain in the cockpit during that time. Cockpit Door Jammed S When cockpit exit required - Remove the inspection seals from the two latches of the door lelft panel (facing the door) - Cockpit Door Locking System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF - Actuate the two corresponding latches as per placards - Disconnect the electrical plug from the top of the door - Push and remove corresponding door panel COCKPIT DOOR OPERATION : PROCEDURES FOLLOWING FAILURES Cockpit Door Control Panel FAULT LIGHT illuminates Ć Move the manual lock bolts to CLOSE position Ć Note : when the door is locked with the manual bolts, the emergency access to the cockpit is unavailable. It is recommended that at least two crew members remain in the cockpit during that time. DC ESS BUS Lost Ć Apply Cockpit Door Control Panel FAULT LIGHT illuminates procedure Ć Note : Cockpit Door Control Panel FAULT LIGHT is inoperative. R Mod : 5377 +(5434 or 8330) Model : 102-202-212-212A 2.01.10 LIMITATIONS P 4 COCKPIT DOOR SECURITY SYSTEM 200 SEP 04 AA COCKPIT DOOR OPERATION : OPENING THE COCKPIT DOOR FROM THE CABIN CABIN CREW ROUTINE ACCESS . . . . . . . . . . . . . . . . . REQUEST on the Door Call Panel CABIN CREW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS CALL PB CABIN CREW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . STAND IN COCKPIT DOOR AXIS The cabin crew should stand in the axis of the cockpit door. A buzzer sounds in the cockpit. H If entry is not authorized by the flight crew : The flight crew denies the entry request via the DOOR TOGGLE SW. The Door Call Panel red light comes ON steady, and indicates that the door is locked. Emergency access, the buzzer, and the Door Call Panel are inhibited for 3 minutes. H If entry is authorized by the flight crew : The flight crew unlocks the door via the DOOR TOGGLE SW. The Door Call Panel green light comes ON steady, and indicates that the door is unlocked. CABIN CREW . . . . . . . . . . . . . . . . . . . . . . . PULL the DOOR RIGHT PANEL to OPEN CABIN CREW . . . . . . . . . . . . . . . . . . . UNLOCK the LEFT PANEL and PULL to OPEN The door left panel is unlocked by moving the handle located in the cockpit side on the door left panel H If there is no reaction from the flight crew : CABIN CREW SECOND ACCESS REQUEST . . . . . . REQUEST on the Door Call Panel Repeat the above procedure. H If there is no reaction from the flight crew after a second request : CABIN CREW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CALL THE COCKPIT To establish contact with the flight crew and request access to the cockpit. H If there is no reaction from the flight crew after a cabin crew interphone call : CABIN CREW . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . PRESS THE EMERGENCY PB Rotate the protecting plate and press the EMER PB. A buzzer sounds continuously in the cockpit for 30 seconds, and the green light flashes on the Door Call Panel. After 30 seconds, the green light comes ON steady and the cabin crew can then pull the door right panel to open and the buzzer stops. This indicates that the door is unlocked for 10 seconds. CABIN CREW . . . . . . . . . . . . . . . . . . . . . . . PULL the DOOR RIGHT PANEL to OPEN CABIN CREW . . . . . . . . . . . . . . . . . . . UNLOCK the LEFT PANEL and PULL to OPEN The door left panel is unlocked by moving the handle located in the cockpit side on the door left panel R Mod : 5377 + (5434 or 8330) Model : 102-202-212-212A 2 –01 LIMITATIONS PAGE : 1 AFM GENERAL DGAC APPROVED 001 DEC 96 2 . 01 . 01 – INTRODUCTION Observance of the limitations contained in this chapter is required by law. When operating in accordance with an approved appendix or supplement to this manual, the limitations of this basic Airplane Flight Manual section apply, except as altered by such an appendix or supplement. 2 . 01 . 02 – KINDS OF OPERATION The airplane is certified in the Transport Category, JAR 25 and ICAO annex 16 for day and night operations, in the following conditions when the appropriate equipment and instruments required by the airworthiness and operating regulations are approved, installed and in an operable condition : - VFR and IFR - Flight in icing conditions. - Reverse thrust taxi (single or twin engine) 2 . 01 . 03 – MINIMUM FLIGHT CREW 2 pilots 2 . 01 . 04 – PERFORMANCE CONFIGURATION Refer to 6.01.03 for airplane configuration associated with certified performances. 2 . 01 . 05 – MAXIMUM OPERATING ALTITUDE 25000 ft 2 . 01 . 06 – MANEUVERING LIMIT LOAD FACTORS Gear and flaps retracted : + 2.5 to – 1 g Gear and/or flaps extended : + 2 to 0 g 2 . 01 . 07 – CONFIGURATION DEVIATION LIST Refer to 7.01.02 2 . 01 . 08 – NOISE CHARACTERISTICS The aircraft meets the requirements of ICAO annex 16, Chapter 3 with no weight restriction. Refer to noise characteristics in Appendix 7.01.01 of this manual for noise levels. 2 . 01 . 09 – MAXIMUM NUMBER OF PASSENGER SEATS 74 as limited by emergency exits configuration. Mod : – Eng : – Model : 212 A 2 –02 LIMITATIONS PAGE : 1 WEIGHT AND LOADING AFM DGAC APPROVED 001 DEC 96 2 . 02 . 01 – WEIGHTS STRUCTURAL LIMITATIONS MAXIMUM WEIGHT KG LB RAMP 22180 48898 TAKE–OFF 22000 48501 LANDING 21850 48170 ZERO FUEL 20000 44092 PERFORMANCE LIMITATIONS Maximum take–off weight and maximum landing weight may be reduced by performance requirements related to the following (see chapter 6) : – climb performance (first and second segment, final take–off and en route, approach and landing climb) – available runway length (take–off and landing) – tyre limit speed – brake energy limit, observe BRK TEMP light extinguished for take–off – obstacle clearance (take–off and en route) – en route and landing weight Mod : – Eng : – Model : 212 A LIMITATIONS 2 –02 PAGE : 2 AFM WEIGHT AND LOADING DGAC APPROVED 020 FEB 00 2 . 02 . 02 – CENTER OF GRAVITY ENVELOPE For definition of reference (Mean Aerodynamic Chord (MAC) and moments origin) : see chapter 1 2 . 02 . 03 – LOADING The airplane must be loaded in accordance with the loading instructions given in the WEIGHT AND BALANCE manual. R Mod : 4439 or 5150 Model : 212 A SEP 98 FEB 99 2 –03 LIMITATIONS AFM AIRSPEED AND OPERATIONAL PARAMETERS PAGE : 1 DGAC APPROVED 001 FEB 03 2 . 03 . 01 – AIRSPEEDS • MAXIMUM OPERATING SPEED VMO / MMO This limit must not be intentionally exceeded in any flight regime. VMO = 250 kt CAS MMO = 0.55 • MAXIMUM DESIGN MANEUVERING SPEED VA Full application of roll and yaw controls, as well as maneuvers that involve angles of attack near the stall should be confined to speeds below VA. VA = 175 kt CAS R CAUTION : Rapidly alternating large rudder applications in combination with large sideslip angles may result in structural failure at any speed. • MAXIMUM FLAPS EXTENDED OR OPERATING SPEEDS VFE FLAPS 15 FLAPS 30 185 kt CAS 150 kt CAS • MAXIMUM LANDING GEAR EXTENDED OR OPERATING SPEEDS VLE = 185 kt CAS VLO lowering = 170 kt CAS retracting = 160 kt CAS Model : 212 A DEC 96 AUG 02 05 for fuel temperature limitation. 03 . TAKE–OFF AND LANDING – Tail wind limit : 10 kt – Maximum mean runway slope : ± 2 % Mod : – Eng : – Model : 212 A 001 DEC 96 .04.2 –03 LIMITATIONS PAGE : 2 AFM AIRSPEED AND OPERATIONAL PARAMETERS DGAC APPROVED 2 . 02 – OPERATIONAL PARAMETERS ENVIRONMENTAL ENVELOPE Note : Refer to 2. 13 % during cruise.2 104.2 101 55 to 65 0 to 125 GROUND IDLE 66 mini HOTEL MODE TRANSIENT 40 mini 715 20 s 120 840 5s 106. 5) ITT limits depend on outside air temperature.2 % during climb and 73. 04 . 2) Oil temperature must be maintained above 45o C to ensure protection for the engine air inlet against ice accumulation. Model : 212 A OCT 00 .9 101. beyond 5 mn.4 106.04 page 2.8 106(3) (4) – 40 to 125 55 to 65 125 40 to 100 120 20 mn STARTING 5s 140 950 –54 mini 1) Value linked to 100 % NP. 01 – ENGINES PRATT AND WHITNEY OF CANADA PW 127F Operating limits with no unscheduled maintenance action required Beyond these limits refer to maintenance manual OPERATING CONDITION TIME LIMIT TQ (%) ITT (oC) NH (%) NL (%) NP (%) OIL PRESS (PSI) OIL TEMP (oC) (2) RESERVE TAKE OFF 10 mn (6) 100 (1) 800 103. Refer to 2.2 –04 LIMITATIONS DEC 96 001 PAGE : 1 POWER PLANT AFM DGAC APPROVED FEB 01 2 .9 (1) 800 103.2 101 55 to 65 0 to 125 NORMAL TAKE OFF 5 mn 90 (1) (5) 101.4 101 55 to 65 0 to 125 MAXIMUM CONTINUOUS none (6) 90. 6) Single engine operation only. 4) Up to 75 % NH only. R NOTE : Flight with an engine running and the propeller feathered is not permitted.2 104. 3) Permissible for completion of flight provided TQ does not exceed 75. 2 –04 LIMITATIONS PAGE : 2 POWER PLANT AFM DGAC APPROVED TAKE OFF ITT LIMITATION AT NORMAL TAKE OFF RATING Mod : – Eng : – Model : 212 A 001 DEC 96 . H Avoid operation in feather above 66. 04 .6 % TQ. Such positioning may lead to loss of airplane control or may result in an engine overspeed condition and consequent loss of engine power.5 % NP. FLIGHT OPERATION H ATR airplanes are protected against a positioning of power levers below the flight idle stop in flight by an IDLE GATE device. H Avoid static operation between 41. 02 – PROPELLERS TWO HAMILTON STANDARD HS 568 F – 1 GROUND OPERATION H Engine run up must be performed into the wind.2 –04 LIMITATIONS PAGE : 3 POWER PLANT AFM DGAC APPROVED 001 DEC 96 2 . It is reminded that any attempt to override this protection is prohibited.6 % and 62. GROUND OR FLIGHT OPERATION H If a propeller is involved in an overspeed or in an engine overtorque refer to the propeller maintenance manual. Mod : – Eng : – Model : 212 A . 2 . a minimum fuel temperature must be taken into account to ensure adequate relight : --34oC for fuel types JET A and JET A1.OIL SYSTEM SPECIFICATION Refer to specification PWA 521 type II.04 PAGE : 4 EASA APPROVED 001 JUL 06 2 .STARTER 3 starts with a 1 minute 30 seconds maximum combined starter running time followed by 4 minutes OFF. R REFUELING Maximum pressure 3. Model : 212 A . R TEMPERATURE • For flight preparation. 04 . Use of JP4 and Jet B is prohibited. UNBALANCE Maximum fuel unbalance : 730 kg (1609 lbs) R FEEDING • Each electrical pump is able to supply one engine in the whole flight envelope. • One electrical pump and associated jet pump are able to supply both engines in the whole flight envelope. • Maximum temperature : 57oC for fuel types JET A. 04 . 03 -. JET A1 and JP5. • One jet pump is able to supply both engines in the whole flight envelope. 05 -. 04 .FUEL SYSTEM R R Acceptable fuels : Jet A. Jet A1 and JP5. 2 . 04 -.LIMITATIONS AFM POWER PLANT DEC 96 FEB 05 2 -. --26oC for fuel type JP5.5 bars (50 PSI) USABLE FUEL The total quantity of fuel usable in each tank is : 2500 kg (5510 lbs) NOTE : Fuel remaining in the tanks when quantity indicators show zero is not usable in flight. .01. 05 . . . if OAT exceeds ISA + 25. . . . 03 – LANDING GEAR Towing with TOWBARLESS system is prohibited. 04 – FLAPS Holding with any flaps extended is prohibited in icing conditions (except for single engine operations). . . . . . . . . . . . . flight level must be limited to FL200. 02 – HYDRAULIC SYSTEM Specification : Hyjet IV or Skydrol LD 4 2 . . 20000 ft 2 . . . . . 05 . . 2 . 0. . . . . . . . . . . . .03 for CAT II operation 2 . .5 PSI Maximum differential pressure for landing . 05 . . . 09 – MFC Take off with two or more failed MFC modules is prohibited 2 . 2 . . 10 – COMMUNICATIONS Not applicable Model : 212 A . . . . . . 2 . 01 – AIR PRESSURIZATION Maximum differential pressure .35 PSI Maximum differential pressure for OVBD VALVE full open selection . . . . . 05 . . . . . . Do not perform pivoting (sharp turns) upon a landing gear with fully braked wheels except in case of emergency. . . . . . . – 0. 05 . 05 . . . . . 05 . . . 05 . . . . . . 05 – AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) – Minimum height for autopilot engagement on take off : 100 ft – NAV mode for VOR approach. . . . . . . . . . . 05 . . . . . . . . . 1 PSI Maximum altitude for one bleed off operation . 05 . . . . . . . . and – DME HOLD is not selected – Minimum height for use of either autopilot or flight director : – Except during take off or executing an approach : 1000 ft – VS or IAS mode during approach : 160 ft – CAT 1 APP mode : 160 ft Refer to 7. . . . . .35 PSI Maximum negative differential pressure . . . . . . 6. . . . . . 2 . . . 06 – INSTRUMENTS MARKING Red arc or radial line : minimum and maximum limits Yellow arc : caution area Green arc : normal area 2 . 07 – CARGO DOOR OPERATION Do not operate cargo door with a lateral wind component of more than 45 kt. 08 – ELECTRICAL Single DC GEN operation : – In flight. . . . . . . . . . . . . . using either autopilot or flight director is authorized only if : – a co–located DME is available.LIMITATIONS PAGE : 1 AFM R SYSTEMS DEC 96 2 –05 DGAC APPROVED 001 FEB 04 2 . . . 04 Mod : 3832 Model : 212 A 100 FEB 00 . 11 – TCAS Refer to appendix 7. 05 .01.LIMITATIONS 2 –05 PAGE : 2 AFM SYSTEMS DGAC APPROVED 2 . In addition. 12 – GPS 1 – GENERAL The Honeywell/Trimble GNSS 1000 : . AC 20–130A. the following procedures apply for B–RNAV operations : (a) during the pre–flight planning phase.This equipment is approved for use as : .REV1 and FAA AC 20–138 and FAA Notice N8110–60. .complies with TSO C 129 and TSO C115B. . en route.LIMITATIONS 2 –05 JUL 00 360 PAGE : 3 AFM SYSTEMS DGAC APPROVED FEB 01 2 .is installed in compliance with FAA AC 20–129. Conventional means must be permanently cross–checked during the approach. Mod : 5176 Model : 212 A . NOTE : Stand alone GPS approach is not approved.supplemental navigation means. in terminal area and for non precision approach operations until the missed approach point with respect of the MDA.The system must be used with an updated active data base and the waypoints position must be cross–checked with official charts.primary navigation means for oceanic and remote operations when only one long range navigation system is required. . 05 . Aircraft operators must apply to their Authority for such an approval/authorization. (b) Traditional navigation equipment must be selected to available aids so as to allow immediate cross–checking or reversion in the event of loss of GPS navigation capability. AC 20–138 and DGAC CRI S–9902.The HT1000 pilot’s guide must be available on board.advisory VNAV means NOTE : VDEV function must be permanently monitored. 2 – LIMITATIONS Compliance with the above regulations does not constitute an operational approval/authorization to conduct operations. . . the availability of GPS integrity (RAIM) must be confirmed for the intended flight (route and time). . . .has been demonstrated to meet the requirements of JAA TGL no2. 3 – PROCEDURES .The system must operate with HT 1000–060 software version or any later approved version. Dispatch must not be made in the event of predicted continuous loss of RAIM of more than 5 minutes for any part of the intended flight. where the coupled DME option is not installed or if the coupled DME is not operative.In the event of DGR alarm illumination the flight crew must cross–check the aircraft position using conventional means or must revert to an alternative means of navigation. . 13 – CABIN LIGHTING The general cabin illumination system must be used during not less than 15 minutes before each flight. Mod : 5040 Model : 212 A . 05 .LIMITATIONS 2 –05 PAGE : 4 AFM SYSTEMS DGAC APPROVED 100 FEB 99 2 . 14 – EGPWS 1. Navigation is not to be predicated on the use of the terrain display. Note : The Terrain display is intended to serve as a situational awareness tool only.LIMITATIONS AFM SYSTEMS 2 –05 FEB 03 PAGE : 5 100 DGAC APPROVED FEB 05 2. 2.05. It does not have the integrity.To avoid giving nuisance alerts. the predictive TAWS functions must be inhibited when landing at an airport that is not included in the airport database. accuracy or fidelity on which to solely base decisions for terrain or obstacle avoidance. R Mod : 5313 or 5467 Model : 212 A . 01 for associated procedures and to FCOM part 3 and to AFM section 7.LIMITATIONS 2 –06 PAGE : 1 AFM ICING CONDITIONS DGAC APPROVED DEC 96 001 FEB 00 2 . NOTE : This supersedes any relief provided by the Master Minimum Equipment List (MMEL). control surfaces or propellers.06. The ice detector must be operative. 06 . snow. fog with visibility of less than one mile.03 for advisory information on contaminated runways penalties./. R Ć Operation in atmospheric icing conditions : NP setting below 82 % is prohibited.04. snow or ice is adhering to the wings.01 for associated procedures and 6. Refer to 3. • Ground icing conditions exist when – OAT on the ground is at or below 5oC. Ć Operation in ground icing conditions : Refer to 3. sleet and ice crystals). standing water or slush is present on the ramps taxiways and runways. – and visible moisture in any form is present (such as clouds... . All icing detection lights must be operative prior to flight at night. rain. Take–off is prohibited when frost. Model : 212 A FEB 99 ..02 for performance data.. – and surface snow.04. 01 – ICING CONDITIONS • Atmospheric icing conditions exist when – OAT on the ground and for take–off is at or below 5oC or when TAT in flight is at or below 7oC. immediately request priority handling from Air Traffic Control to facilitate a route or an altitude change to exit the icing conditions. Visible rain at temperatures close to 0_C ambient air temperature (SAT) . or when unusual lateral trim requirements or autopilot trim warnings are encountered while the airplane is in icing conditions. Water splashing and streaming on the windshield . Unusually extensive ice accreted on the airframe in areas not normally observed to collect ice. 01 – ICING CONDITIONS (cont’d) Ć Severe icing : WARNING : Severe icing may result from environmental conditions outside of those for which the airplane is certificated. Droplets that splash or splatter on impact at temperatures close to 0_C ambient air temperature (SAT) If one of these phenomena is observed. This ice may not be shed using the ice protection systems. Apply procedure specified in the Emergency Procedures chapter. severe icing conditions that exceed those for which the airplane is certificated shall be determined by the following : Visual cue identifying severe icing is characterized by ice covering all or a substantial part of the unheated portion of either side window. – During flight. use of the autopilot is prohibited when the severe icing defined above exists. or mixed icing conditions (supercooled liquid water and ice crystals) may result in ice build–up on protected surfaces exceeding the capability of the ice protection system.LIMITATIONS 2 –06 PAGE : 2 AFM R ICING CONDITIONS DGAC APPROVED 001 FEB 04 2 . – Since the autopilot may mask tactile cues that indicate adverse changes in handling characteristics. 06 . . Model : 212 A DEC 96 FEB 99 . and may seriously degrade the performance and controllability of the airplane. and / or The following secondary indications : . Accumulation of ice on the propeller spinner farther aft than normally observed. Accumulation of ice on the lower surface of the wing aft of the protected area. Flight in freezing rain. The following weather conditions may be conducive to severe in–flight icing : . or may result in ice forming aft of the protected surfaces. freezing drizzle. and / or Unexpected decrease in speed or rate of climb. . PROCEDURES AND TECHNIQUES 2.02.10 P1 FLIGHT PATTERNS AA R R R R R R R R R R R Mod : 3973 or 4371 or 4457 080 JUL 98 . PROCEDURES AND TECHNIQUES 2.02.10 P2 FLIGHT PATTERNS AA R R R R R R R R 001 JUL 98 . 10 P3 FLIGHT PATTERNS AA R R R R R R R R R R R R R 001 JUL 98 .PROCEDURES AND TECHNIQUES 2.02. PROCEDURES AND TECHNIQUES 2.02. : 1368 + 4457 220 JUL 98 .10 P4 FLIGHT PATTERNS AA R R R R R R R R R R R R R R R R R R R R R Mod. 02.PROCEDURES AND TECHNIQUES 2. : 1368 + 4457 280 JUL 98 .10 P5 FLIGHT PATTERNS AA R R R R R R R R R R R R R R R R R R R R R Mod. 10 P6 FLIGHT PATTERNS AA R R R R R R R R R R R R R R R R R 001 JUL 98 .PROCEDURES AND TECHNIQUES 2.02. . PROCEDURES AND TECHNIQUES 2.10 P8 FLIGHT PATTERNS AA R 001 JUN 97 .02.