Section 7 - Airframe and Systems Description
Last updated
Last updated
P2012âs airframe can be divided in the following main groups:
Fuselage
Wings
Empennages
Flight Controls
Landing Gear
Brakes and Wheels
The fuselage is composed by aluminum formed sheet metal frames, beams, stringers and panels coupled by solid rivets or, where necessary, by bolts. Some major items, as undercarriage beams, wing and tail surface attachments, are machined. The cabin can accommodate up to eleven occupants: nine passengers and two pilots (one pilot is considered as minimum crew). Suitable allowance for luggage and optional equipment is also provided both in nose compartment and in the rear of the fuselage.
The cantilever wing structure is composed of three main parts of about equal span. The central part, having rectangular plan form, constant chord, without dihedral angle, is joined to fuselage through four attachments, at its extremities two tapered wing parts are connected. In the central section, nearby its ends, the two engines in their nacelles are installed, extending throughout the wing chord. In the taper wing parts, the fuel tanks are located.
The empennages, like the wing, are made of aluminum traditional construction except for the leading edges made of composite material. The vertical surface consists of a conventional vertical stabilizer and rudder.
The horizontal surface, consists of a stabilizer and an elevator. Rudder and elevator tips are extended so to provide aerodynamic and static balancing by means of weights. Wide trim-tabs are provided on all control surfaces.
The primary flight controls are of conventional type operated by control wheels and rudder pedals. Elevator is actuated by push-pull rods and cables. Ailerons are actuated by push-pull rods on wing and cable in fuselage. The rudder is operated by a cable line. Trimming devices for primary controls are provided by push/pull rod-type system controlled by electrical actuators. Trim controls are provided by means of knob/switches as reported in the following:
Pitch trim switch on the control wheel;
Rudder and Aileron trim on the central pedestal.
Trim disconnection switches are provided for each trimming device.
The flap control system is actuated by means of a linear electrical actuator connected to rods and cables transmitting the movement to the flap surfaces. Flap control is provided by means of a knob located on central pedestal.
Trims and flap position indications are provided by the avionics system. The position indicators are always visible in Garmin display by means of dedicated box in EIS full page (a), or in a more compact way when EIS information is displayed in Strip page (b).
The main landing gear, is composed of two fixed legs containing the oleo- pneumatic shock absorbers. These legs are installed on a box beam passing across the fuselage and faired by means of small streamlined pods.
P2012 is provided with an independent hydraulically actuated brake system for each main wheel. A master cylinder is attached to each pilotâs rudder pedal. Hydraulic pressure, applied via the master cylinders, enters the brake via lines connected to an inlet fitting on the caliper.
A parking brake valve, mounted under the cabin floor and operated by a knob, intercepts the hydraulic lines, once pressurized by toe brakes, to hold the brake assembly linings tight around the main wheels brake discs. Brakes can be operated from either pilotâs and co-pilotâs pedals: a single vented oil reservoir feeds the pilot side master cylinders which are connected, via hoses, with the co-pilotâs side ones.
The main landing gear is provided with two 6.50-10 type wheel; each wheel is equipped, on the inboard side, with a hydraulically actuated single disc brake.
The nose gear, steerable type, is fixed and consists of an oleo-pneumatic shock-absorber leg, attached to the first fuselage bulkhead and it is provided with a 6.00-6 type wheel.
The cockpit control panels (and their relevant markings/placards) of the P2012 are installed on the following locations:
Overhead Panel
Instrument Panel
Control Wheels
Central Pedestal
With regards to the controls of instruments that provide information about avionics and engine monitoring systems, which are integral to the instrument itself, the controls are appropriately marked by the manufacturer, accordingly to the method of operations.
On the overhead panel, in addition to the breakers panel, are provided controls for the following equipment/systems:
Electrical system and external lights
Fuel selectors
Engines and map lights
PFD LH
FIRE DETECTOR PANEL
AIRRAFT IDENTIFICATION PLACARD
EECS PANEL
ALTERNATE AIR CONTROL PANEL
N/A
N/A
ELECTRICAL SYSTEM (EMERGENCY) CONTROL PANEL
TKS ICE PROTECTION SYSTEM
PITOT HEATERS SWITCHES
STALL HEATER SWITCHES
FASTEN/UNFASTEN SEAT BELT LIGHT SWITCH
DISPLAY BACK-UP REVERSIONARY BUTTON
AP MASTER
TRIM DISCONNECT SWITCH
FLAP CONTROL SWITCH
FMS KEYBOARD
AUDIO PANEL
COCKPIT AIR CONTROL PANEL
CABIN HEATER CONTROL PANEL
N/A
AIR CONDITIONING CONTROL PANEL
DISPLAY BACK-UP REVERSIONARY BUTTON
PFD RH
MFD
AUTOPILOT CONTROL PANEL
MD302 STAND-BY INSTRUMENT
ELT
INTERNAL LIGHTS CONTROL PANEL
WoW SWITCH
AOA/SW PFT BUTTON
SW ICE MODE SWITCH
BATTERY HEATING LIGHT
HOUR-METER
On the control wheel (pilot side), are provided controls for the following equipment/systems:
Pitch trim
Autopilot :
Autopilot disconnect
CWS
PTT - Push to talk
The pitch trim control is used to command manual electric trim. This composite switch is split into left and right sides. The left switch is the ARM contact and the right switch controls the DN (forward) and UP (rearward) contacts. The switch can be used to disengage the autopilot and to acknowledge an autopilot disconnect alert and mute the associated aural tone. Manual trim commands are generated only when both sides of the switch are operated simultaneously.
On the central pedestal, controls are provided for the following equipment/systems:
Throttles and Propeller Controls
Aileron trim
Rudder trim
Parking brake
Powerplant levers friction
Alternate static source
P2012 avionics system is mainly based on the Working Title G1000Nxi. Only two avionics devices, ADF and DME, are not part of Garmin system but they interface with G1000Nxi.
An independent stand-by instrument Mid-Continent MD302, along with a magnetic compass, provides the pilot with a backup of primary flight information, namely airspeed, altitude and attitude.
An ELT device is installed and its control switch is located in the lower central area of the cockpit, below the stand-by instrument.
Below in this paragraph, all of the avionics equipment installed on the P2012 will be individually described.
GDU 1050/1250A units display flight and engine parameters and moving map information and act as the user interface for P2012 avionics suite. GDU 1050 units placed on instrument panel LH and RH sides are identified as Primary Flight Display (PFD), while GDU 1250A placed between two PFD displays are identified as Multifunction Display (MFD).
A dedicated calibration curve for automatic brightness adjustment of both displays and keys is included in the system in order to meet the visibility requirements in each foreseen operational condition. This feature provides the pilot with an automatic dimming of both display and keys in accordance with the lighting conditions sensed by the two external light sensors positioned in the upper right and lower left corners of each display. Furthermore, the pilot can easily access a manual dimming mode for both displays and keys.
In the event of a single display failure, the system will automatically switch the critical information including flight and engine parameters on the remaining display presenting them in a compact view. In the event of a failure of the automatic switch logic, the pilot can force the reversionary mode by pressing the dedicated reversionary mode button close to PFD thus getting both flight and engine parameters information, necessary for continued safe flight, on the remaining display.
GDU 1050 provides the interface for NAV and COM information, displaying the functions listed further:
Flight instruments functions
Display of attitude (pitch and roll) rate of turn, slip/skid, directional, airspeed, altitude and vertical speed (PFD function); display of Outside Air Temperature, annunciator panel, navigation functions transponder management;
Display of engine and airframe instrumentation (MFD or reversionary modes only).
Navigation instruments functions
Display of position and ground speed;
HSI, Selected Heading and Selected Course (PFD);
Display of ADF and DME information;
Area Navigation functions
Baro-altitude Vertical Navigation.
Interface functions
Interface with GIA 64W Integrated Avionics Unit (IAU) and with second GDU 1050 unit;
Control and operations of two independent communication systems that operate on frequencies from 118,00MHz to 136,975MHz through increments of 8,33kHz or 25kHz;
Control and operations of two independent navigation systems VOR/ILS that operate on frequencies from 108,00MHz to 117,95MHz through increments of 50kHz or 25kHz;
Control and display of GTX 345R transponder;
Control of remote DME;
Control of remote ADF.
GMA 350C unit is a horizontally installed audio panel located in the lower center side of GDU 1250A unit, made with backlit LEDs. This LRU features Bluetooth interface. It is connected to two speakers, in order to allow pilot and co-pilot to communicate with passengers.
MD302 is a Standâby Attitude Module installed in order to provide the pilot with flight information in case of failure of the PFDs and MFD. It is a digital instrument featuring airspeed, altitude, attitude and slip information.
The display has automatic dimming adjustment in order to have proper visibility in all operative conditions.
The design is built around a solid-state electronic sensor array for high reliability and contains an integral and rechargeable NanophosphateÂŽ lithium- ion battery that can power the unit for up to one hour in the event of complete loss of main 14 VDC battery bus power. The dual, high-resolution LCD display uses a configurable lighting response curve to ensure visibility in all conditions. The user interface of the MD302 allows operation using a single push-and-turn control knob that easily navigates through the user options and menu screens. The airspeed indication markings on the MD302 do not automatically compensate for changes in VNE and/or VNO at altitudes above 15000 feet.
The Tecnam P2012 cabin can accommodate up to eleven occupants: nine passengers and two crewmembers (one crew is considered as minimum crew).
The P2012 general layout of doors is shown below and the aircraft is equipped with:
a front baggage compartment door (left side);
two front crew doors;
a main door including passengers and cargo door (upper side of passenger door, functions also as emergency exit);
one emergency rear door (right side).
The cargo doors are positioned one on the left front and another one on the left rear part of the passenger compartment.
All doors (excluding right emergency and front cargo doors) embody a solenoid, which is activated by engine oil pressure, following engine starting, resulting in placing the locking lever up. The locking lever blocks the opening system of the door avoiding inadvertent opening during the flight.
Warning messages are provided in case of open and/or unlocked doors.
P2012 is equipped with two 375 horse power electronically controlled Lycoming piston engines.
TEO-540-C1A engine, is a direct-drive six-cylinder, horizontally opposed, turbocharged, air-cooled engine. It features electronic fuel injection, electronic ignition, down exhaust, and induction air coolers. This engine has an automotive type starter, a 28V alternator (140A) and a propeller governor. Engine specifications are given in the following table.
The EECS (Electronic Engine Control System) is an electronic, microprocessor controlled system that continuously monitors and adjusts ignition timing, fuel injection timing, and fuel mixture based on operating conditions.
The EECS eliminates the need for magnetos and manual fuel/air mixture. The EECS connects engine hardware with electronic controls to replace mechanical control systems and enables single lever engine control.
Monitoring of engine parameters is performed by means of sensors. The EIS (Engine Interface System) transmits data to the avionics system to be shown on the relevant displays.
The engine features an Electronic Engine Control System (EECS), a microprocessor continuously monitoring and automatically adjusting operating conditions such as ignition timing, fuel injection timing, and fuel mixture.
In the EECS there are several over-limit protection schemes, the most aggressive of which is a cyclic deactivation of cylinders ignition to avoid extensive over-speed or over-boost of the engine, active until any perceived limit exceedance is negated. This misfiring of the engine translates itself to the pilot as a hesitation of the engine, even if power is preserved.
Under some circumstances (such as rapid PWR levers movements or RPM increases with PWR levers full, especially with cold oil) it is possible that the engine can be over boosted until the activation of the over-limit protection scheme. This would most likely be experienced during transitions to full power. Therefore, it is still necessary that the pilot observe and be prepared to control the manifold pressure.
The engine air intake system for P2012 airplane consists of two openings installed on wing leading edge (one for each side), they are both located on the right side of the engine cowling and feed the necessary air required for proper operation of engine turbocharging system.
In case of loss of flow from air filter due to ice build-up, snow or other causes that prevent air passage through the air filter, a lateral door of the valve can be opened manually by the pilot through a push pull alternate air control knob allowing the air to enter in the valve directly from the wing leading edge.
The alternate air control knobs are located on the left side of instrument panel, easily visible and accessible to the flight crew and placarded according to the method of operation. In the normal operation the alternate air control knobs are set to OFF position (all in).
Fire detectors (two for each engine) are installed inside engine compartments. In case of fire in engine compartment, the fire detector advises pilot through a warning on primary flight display. Check of proper functioning of the system is carried out by the pilot during normal pre-flight procedures through a dedicated push-to test button located on the instrument panel.
P2012 A/C is equipped with two MTV-14-B-C-F/CF195-30 four blades variable pitch and constant speed propellers. Blades are made of laminated wood composite structure, epoxy-fiber glass cover, with leading edge and erosion protection; propeller hub is made of aluminum alloy.
The P2012 fuel system provides the following functions: fuel storage, fuel distribution and fuel indications, and is therefore composed by sub-systems.
The RH system is exactly symmetric with respect to the LH.
Fuel system consists of two fuel tanks integrated in the wing box and having a capacity of 375 lt (99 USG) for a total capacity of 750 lt (198 USG).
Each engine is fed by means of an engine-driven mechanical pump (Main Pump) provided by Lycoming and integrated in the housing of the engine coupled with an electric fuel pump (connected to the essential bus) provided by Weldon and activated by means of ignition switch. In addition, there is an electric fuel boost pump (connected to the main bus), always provided by Weldon, and activated by means of switches located in the Engine Control Panel.
Two fuel selectors, one per engine, are located at the top of the cabin and have three operating positions:
normal feeding is attained with each fuel selector set on the fuel tank corresponding to the same engine;
engine crossfeed operation is attained with the fuel selector set on the opposite fuel tank;
fuel shutoff is attained setting the fuel selector to the shutoff position. In order to enter in shut-off position or exit from, it is necessary to unlock the mechanical tab on the lever itself.
Two resistive type fuel quantity senders are installed in each tank and provide the fuel indication on the A/C cockpit. Each tank is provided with a low fuel quantity sender.
P2012 electrical system is based on 28Vdc voltage. The primary electrical power source is provided by LH and RH engine driven alternators rated of 140 Amps each and three batteries, listed below:
Main battery (17Ah)
LH battery (17Ah)
RH battery (17Ah)
The batteries, namely True Blue Power TB17 Lithium Ion type, can operate individually or in parallel and are used to start the engines and to provide power to the units when the A/C is on the ground while the engines are not running and therefore the alternators are switched off. They also act as a reserve of power in emergency conditions (dual alternator failure). When engines are running the alternators also supply the power required in order to recharge the batteries.
Each battery is provided of an internal heating system that, in case of cold weather conditions, allows to be operated also at temperatures down to -40°C (-40°F). An indication light is provided to monitor heating system functioning. On ground, in extremely cold temperatures, this light ON informs of the momentary unavailability of MAIN Battery electrical power due to its self-heating and, therefore, the impossibility to power up the Aircraft.
The system features four distribution buses, namely: Main bus, Essential bus, LH bus, RH bus.
Dedicated switches allow the selection of active/inactive buses. Under normal operation, the LH DC generator will power the LH bus, and the RH gen the RH bus. In order to avoid the need to balance the generator outputs, a TIE bus switch is added.
The electrical loads are connected to the buses through dedicated circuit breakers. Switches are installed in order to allow the pilot the control of loads, where required. Essential bus is fed from 4 points and includes electrical loads required for continued safe flight and landing.
A/C power is based upon 3 independent energy sources, two alternators (one per side) and a set of 3 batteries, whose features are reported in Table above. The power sources are able to run independently or together; each source has a dedicated switch: the switches to enable and disable alternators and batteries are grouped in the master switches group and are located in the forward center zone of cabin ceiling.
In order to monitor the batteries status some CAS messages are integrated in the G1000Nxi system.
In addition, digital indication of the following parameters is provided by the avionics system:
LH Alternator current
RH Alternator current
Main Battery current
and as regards the voltmeter indications:
Main Bus voltage
Essential Bus voltage
LH Battery voltage
RH Battery voltage
The failure of one of the energy sources will not affect the operation of alternate energy sources. In the event of single alternator failure, the remaining alternator will be able to feed the electrical loads even if load shedding may be required by the pilot.
The master switches group is located in the forward center zone of the cabin ceiling. When the MASTER switch is in ON position enables the master relay and connects the battery to the main bus.
Two switches, named FIELD LH and FIELD RH and located respectively to the left and right side of master switch, control the alternator (LH/RH) field signal which enables the alternator by providing it the field current.
Two switches named LH Engine backup battery and RH Engine backup battery, when in ON position, connect the LH and RH batteries to the LH and RH ECU respectively.
Two switches named CROSS LH and CROSS RH, when in ON position, connect the main bus bar to the BUS bar LH and RH. For safety and for engine starting, P2012 has a bus tie switch which can be manually operated by the pilot.
The master switches are marked with dedicated labels and due to their location are easily accessible to the pilot. The following figure shown the master switches layout.
A dedicated emergency panel is provided to housing guarded switches to be activated in case of emergency.
In detail two switches are identified with the following functions:
Essential bus switch: controls an emergency connection between master relay and the essential bus.
Battery switch: activates the emergency relays in order to connect in parallel the main battery and both the LH/RH ECU battery.
There are two EECS (Electronic Engine Control System) enabled by means of LH Engine backup battery and RH Engine backup battery switches which are located in the center zone of cabin ceiling on the electrical panel.
Main components of EECS are:
Engine Control Unit (ECU): a dual channel unit which contains system processors, input signed conditioning, output actuator drive stages and aircraft communication interfaces.
Power Box Unit (APU) that supplies power to the EECS:
(a) RPM ⤠1000: airplanes batteries supply power to the APU;
(b) RPM > 1000: a dedicated Permanent Magnet Alternator (PMA) supplies power to the APU.
Ignition LH/RH panel is located near the breaker panel.
As per figure below the ignition panel includes ignition switches (one for engine), start push-buttons (one for engine), electric fuel pump switches (one for engine).
The ignition switches are guarded so that they cannot be inadvertently operated by the pilot.
Dedicated PFT (preflight test) buttons are provided on the EECS control panel to check the status of the ECU.
The EECS has a simple set of hardwired warning annunciators to supply critical system status data to the pilot. These annunciators are connected directly to the ECU. The annunciators will still operate if airplane electrical power is lost.
The static system of P2012 consists of two static ports installed on the left and right side of aircraft external surface, behind the pilots doors.
A âTâ joint connects the two static ports to a single line that goes to the alternate static port joint (that is available through a âON/OFFâ valve placed in the cabin) and then to the left (ADC#1) and right (ADC#2) air data computer system and to the MD-302 backup instrument (altimeter and airspeed indicator)
The alternate static port, located in the cabin, on central tunnel, is provided with a valve that, in case of failure of the external primary static ports, must be opened to ensure the back-up operation of flight instruments.
The drainage of the static system is provided by means of a dedicated valve located in the lowest point of static system, so the moisture and the water can be drained out from this point.
The total pressure system of P2012 is measured by dual heated pitot probes.
The Pitot tubes are connected to the WT G1000NXi system and to the stand-by instrument (MD302).
A hole at the bottom of pitot tubes provides drainage of moisture and water.
The Pitot tube equipped on P2012 has the possibility to be heated electrically. A switch, located on the instrument panel activates the Pitot heating system.
The heating system is provided by means of dual internal electro-thermal heating elements (two for each pitot) to provide ice protection to the areas of the probe where surface impingement of icing elements is expected to occur. On ground, only one heating system is active (reduced power). The functioning of total heating elements (full power) is provided only when aircraft is âin airâ condition as detected by weight on wheel installed on main landing gear. To verify on ground the proper functioning of the entire system, a WoW switch that simulates âin airâ condition, is installed on the instrument panel.
Since the aircraft is equipped with two heated pitot systems, a related resistance G1000Nxi system is set up in order to show the operational state of heated pitot system.
CAS messages (refer to Section 3, §1.3) are provided in order to show the status of the system.
An additional caution light (PIT HTRS SET ON) is shown when PITOT HEAT #1/#2 are OFF and OAT is equal or lower than 5°C.
The aircraft is equipped with a vane- type stall warning sensor, installed in the leading edge of the right wing. The vane is connected to a stall warning horn located in the overhead panel above the pilot seat. The vane in the wing senses the change in airflow over the wing occurring near to the stall activating the warning horn. As consequence, a relevant CAS aural/warning message STALL is provided by the avionics system. In addition, if autopilot (if installed) is engaged, it will be disconnected.
The stall warning system is checked during the preflight inspection of the aircraft by actuating the vane in the wing while the master switch is ON. The system is functioning properly if the warning horn sounds as the vane is pushed upward. The stall warning system is connected to the Essential Bus and it is protected by a circuit breaker.
A switch located in the instrument panel activates the stall warning heating system. The heater system, connected to the LH Bus, is protected by a circuit breaker. CAS GREEN message STALL HTR ON is provided in order to show the proper status of the system.
An additional caution light (STALL HTR SET ON) is shown when stall heater is OFF and OAT is equal or lower than 5°C (41°F). Even if the stall warning is provided of a heating system, the aircraft is not approved for flight into know icing (FIKI) conditions.
Additional Components of the stall warning system:
dedicated WoW (weight on wheel) pressure switch, installed on the main landing gear, selects automatically the operating mode (GROUND/AIR). This switch prevents the stall warning horn from sounding while performing ground operations.
Dedicated CAS messages provide information to the pilot of the current operating mode.
When the aircraft is on ground, to test proper functioning of entire system, a WoW switch, installed on the instrument panel, allows to force the system in AIR mode. When the switch is in AUTO position, the operating mode is managed by the pressure switch on the landing gear. The switch on the instrument panel can be used by the pilot also in flight in case of emergency (WoW pressure switch forced in ground position), if necessary.
An indication of angle of attack is provided within Garmin PFD below the airspeed tape. The indicator shows non dimensional units from 0.2 to 1.0, and contains colored arcs. The green arc, located from 0.2 to 0.60, represents normal operation. The yellow arc, located between 0.60 and 0.80, represents the area where the aircraft may be approaching the critical AoA value. The red arc, from 0.80 to 1.0, represents where low speed buffet may occur and if uncorrected, could continue on to a full stall. A CAS warning message SW/AOA FAIL is provided to the pilot if a failure of the system is detected.
The Safe Flight Lift Transducer Stall Warning system is equipped with a Pre-Flight Test button which will test functionality of the system during any phase of operation. If the system is functioning normally and the Pre-Flight Test button is pressed, the Stall Warning Horn will sound. If the Stall Warning Horn is not heard while pressing the Pre-Flight Test button, the Safe Flight Lift Transducer Stall Warning system is not functioning properly and is considered inoperative.
During the same test also the functionality of AOA indication will be tested. When the Pre-Flight Test button is activated the AOA digit will indicate 1.00 with the pointer at full upper scale.
In the following paragraph it is provided a description of the external and internal lights installed on P2012 Aircraft.
External lights system is composed of:
Anti-collision lights (on the wing and rudder)
Navigation/Position lights (on the wing and on the rudder)
Landing lights (on the wing)
Taxi lights (on the wing)
Lights are activated by switches located on the external lights control panel installed on the cabin ceiling as indicated in the following figure:
Internal lights system is composed of:
Switches lights â dimmable
Cockpit lights â dimmable
Map lights â dimmable
Passengers cabin lights
Passengers seats lights (one for each seat)
Emergency lights (cockpit)
Emergency door light (always on)
No smoking sign light (always on)
Fasten seat belts light
The avionics suite and the stand-by instrument are equipped with an automatic dimming system based on ambient light detection photocells, able to automatically dim the display and key brightness in accordance with external light conditions.
Lights are activated by switches/knob located on the internal lighting control panel installed on the instrument panel as indicated in the following figures or by means of the rotary knobs (map lights).
Cabin air system is designed to provide the right comfort to the crew and passengers during the flight and, in addition, defog/defrost of windshield.
In detail, the system provides the following functions:
Ventilation for the pilots and passengers;
Heating (for FWD seats (cabin crew)) and Defog/Defrost (for the windshield).
The main components of the ventilation system are:
an external dynamic inlet (always open);
cabin ventilation hoses;
nozzles (one for each passengers cabin seat).
The main components of the heating system are:
Two fans;
A mechanical controlled check valve;
Two thermal resistors;
Y conveyor for windshield diffusers;
Pilot and co-pilot legs air diffuser.
The system is composed of two different lines for air distribution: the first line provides air to the defrost system located on aircraft dashboard, the second one provides air to pilot legs diffuser located on pilot cabin floor behind the pedals.
For both lines the air is directly taken from the front luggage compartments through two dedicated fans. A thermal resistor (one for each line) warms up the airflow (if heating is activated). Each resistor has two transducers arranged in parallel which allow, in the case of failure of one of two items, to keep the heating system running.
The system is activated through two separate knobs located in the cockpit air side of the environmental control panel. With the knob/s in FAN position, the relevant fan is activated to supply the air to the windshield and to the pilot legs. When the HEAT position is selected, the same air is warmed through the resistor.
Two overheat switches automatically cut-off the current to the heater element through their relays in case of inner temperature raising above 100°C (212 °F). In case of overheat relays activation, a reset button allows the system to start working again. It is suggested to keep the blower running for some time (15 to 20 seconds) before the HEATER push button is activated, in order to cool down the system. If the light continues to be ON, the pilot should switch the system OFF and use the emergency control in case defog function is needed.
In case of emergency (failure of the resistor of defrost line), the system is designed to ensure the continuity of windshield defrosting. The pilot must act on a dedicated knob (DEFROST AIR) located on panel instrument to close a mechanical valve and to supply air to the windshield through the legs air system.
| Annunciator Acronym | Annunciator Name | Definition and Pilot Action |
|---|---|---|
NTO Pri
NTO Sec
No Take-Off
If the NTO annunciator is constantly illuminated before, during or after completing the Pre-Flight Test, while the aircraft is on the ground, do not take-off. Maintenance action is required to identify the fault(s), correct the condition(s) and clear the fault(s).
If the NTO annunciator illuminates during flight and stays illuminated for more than 10 seconds during flight, make minimum power control changes and land as soon as safely possible. Although the engine will continue to operate, operation will be in a degraded mode which requires a prompt, safe landing. Identify the fault(s), correct the condition(s) and clear the fault(s) before further flight.
TLO
Time-Limited Operation
If the TLO annunciator illuminates (light permanently ON), it is still safe for flight if necessary. However, it is recommended the fault(s) is identified, the condition(s) corrected and the fault(s) cleared. If service is not performed within 20 hours of cumulative TLO light ON, the NTO annunciator will illuminate.
PFC
Pre-Flight Check
If this annunciator is illuminated, the pre-flight test is in progress. Do not move the power control until the annunciator has turned off. If the pre-flight test sequence is stopped before it is completed, the PFT annunciator will flash. Momentarily press the PFT button. Complete another pre-flight test.
