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Section 6 - Airframe and Systems Description
Each wing consists of a central light alloy torque box which carries all the wing bending, shear and torque loads; an aluminium leading edge is attached to the front spar while flap and aileron are hinged to the rear spar.
The torque box houses an integrated fuel tank and supports the engine mount.
Flap and aileron, respectively located inboard and outboard of wing and made up of light alloy, are constructed with a central spar to which front and rear ribs are jointed. Wrapped-around aluminium stressed skin panels cover all the structures. Steel alloy attachments connect left and right wing to each other.
Following figure shows the left wing fitted with the engine nacelle, fuel tank and composite winglet. Steel alloy attachments link left and right wing to each other.
The fuselage is constituted by a light-alloy semi-monocoque structure wrappedaround by stressed skin panels. Radome and stern fairing are of composite material. Cabin and baggage compartment floor is a warping of beams and keelsons supporting the seats guides and other components.
Two spar frames support on the top the wings attachments and on the bottom the sponson beans sustaining the main landing gear. The forward frame, to which radome is connected, supports a steel trestle to which the nose landing gear is connected.
The front and rear seats access occur by means of two doors located in the opposite sides of the fuselage; a ditching emergency exit is available on the top of the cabin. In tail cone, two spar frames support the horizontal and vertical empennages attachments.
The vertical tail is entirely metallic: vertical fin is made up of a twin spar with aluminium alloy stressed skin. Rudder, providing directional control of the airplane, is made up of aluminium alloy.
The rudder is connected to the vertical tail at two hinge points. A trim tab system increases directional stability of the airplane.
The horizontal empennage is an all-moving type (stabilator); its structure consists of a twin spar to which front and rear ribs are jointed and it is covered by stressed aluminium alloy skin. The trim tab completes the assy.
The main flight control system controls the airplane in three axes. All primary controls (ailerons, rudder and stabilator) are manually operated by a conventional control column and rudder pedals, pulleys, cables, bellcranks and rods.
The secondary flight controls consist of a two-axis trim system and a flaps system. Complete dual controls are provided for pilot and co-pilot.
Longitudinal control acts through a system of push-pull rods connected to the control column and moving the stabilator whose anti-tab winglet works also as trim tab. Autopilot pitch servo is connected to the push-pull rods system through driving cables.
Longitudinal trim is performed by a small tab positioned on the stabilator and manually operated via a control wheel positioned between the two crew seats. As optional, it is available an electrically operated longitudinal trim which it is also controlled by the autopilot system, when installed.
Trim position is monitored by an indicator on the instrument panel. A trim disconnect toggle switch is provided.
Ailerons control is of mixed type with push-rods and cables; a cable control circuit is confined within the cabin and it is connected to a pair of push-pull rod systems positioned in each main wing which control ailerons differentially.
The U-shaped control wheels, hinged on the top of the control column, control the ailerons. Control wheel motion is transferred to the ailerons through a cable loop, up to the interconnecting rod linking the two push-pull rod systems which finally transmit the motion to the ailerons. When either aileron control wheel is rotated, the crossover cable rotates the other control wheel.
The left aileron has a trim tab adjustable on ground: its deflection allows for lateral trimming of the airplane.
Both flaps are extended via a single electric actuator controlled by a switch on the instrument panel. Flaps act in continuous mode; the analogue indicator displays three markings related to 0°, takeoff (T/O) and landing (FULL) positions.
Rudder is operated through a cable system. A rudder trim tab allows aircraft directional trimming, especially in case of OEI operation: it is electrically operated via a switch located on the central console placed between crew seats. Its position is monitored by an indicator on the instrument panel. A trim disconnect toggle switch is provided.
P2006T is equipped with two four-cylinder four-stroke Rotax 912S engines of 98hp (73kW) each, both rotating clockwise. These are partially liquid cooled and they feature an integrated reduction gear driving constant speed propellers with pitch feathering devices.
Cooling system is designed for liquid cooling of the cylinders heads and ram-air cooling of the cylinders. The liquid system is a closed circuit with an overflow bottle and an expansion tank.
The engine is provided with a dry sump forced lubrication system with an oil pump with integrated pressure regulator. A thermostatic valve regulates the oil flow to the heat exchanger (oil radiator) on the basis of oil temperature: this allows the engine starting in cold conditions. Please refer to the quick start guide regarding a proper engine startup procedure in realistic operation mode.
Carburettor heat control knobs are located between throttle and propellers levers; when the knobs are fully pulled backwards, carburettors receive maximum hot air. During normal operations, the knobs are fully forward set (carburettors heating set to OFF).
Internal lights system is composed by following equipment:
- Cabin light, providing lighting for crew and passengers compartment;
- Instruments lights, which in turn are composed by three sub-systems each one fitted with dimming device:
- Switches built-in lights
- Avionics lights
- Cockpit lights
- Emergency light
The cabin light is a ceiling light, fitted with control switches, located on the overhead panel in correspondence of the crew seats.
About the instrument lights (controlled by a switch on the RH instrument panel), the switches built-in lights concern the instrument panels switches lighting, the avionics lights concern the avionic equipment lighting and the cockpit lights concern two lights located on the over-head panel illuminating LH and RH instrument panels (see Figure below).
All above mentioned lights are supplied by the battery bus apart from the Emergency light which is directly connected to the battery. It is a five-leds light located in the over-head panel (see Figure below) controlled by a switch installed on the LH breakers rack.
External lights system consists of the following equipment (see Figure below):
- NAV Lights: they provide, by means of three position lights, the aircraft flight direction identification.
- Strobe Lights: they provide aircraft identification to prevent collision. They are located, like the above mentioned NAV lights, on the winglets and on the top of the vertical fin.
- Taxi Light: supports taxi maneuvering on the ground at night. It is installed on the left wing leading edge.
- Landing Light: provides ground reference information during final approach, touchdown, ground roll and take off and illuminates any major obstructions in the airplane approach glide path or on runway at night. It is installed on the left wing leading edge.
Fuel system consists of two integrated tanks inside the wing torque boxes and fitted with inspection doors.
Each fuel tank has a capacity of 100 litres and is equipped with a vent valve (its outlet is located on the lower wing skin) and a sump fitted with a drain valve for water/moisture drainage purposes.
An electric fuel pump feeds the pertinent engine in case of engine-driven pump failure. The fuel Gascolator (a sediment-filter bowl) is located beneath the engine nacelle, between the fuel tank and the electrical pump, in correspondence of the fuel system lowest point. It is fitted with a drain valve which allows for the overall fuel line drainage.
Fuel quantity indicators and fuel pressure indicators for each engine are located on the RH instrument panel.
In normal conditions, to supply fuel to engines, each engine pump sucks fuel from the related tank; crossfeed is allowed by fuel valves located on the front spar and controlled by Bowden cables from the fuel selectors located on the cabin overhead panel.
Left fuel selector manages the left engine feeding, allowing fuel supply from the left fuel tank or from the right one (crossfeed).
Right fuel selector manages the right engine feeding, allowing fuel supply from the right fuel tank or from the left one (crossfeed).
Each selector can be set in OFF position only pulling and simultaneously rotating the lever: this avoids an unintentional operation.
Primary DC power is provided by two engine-driven generators which, during normal operations, operate in parallel.
Each generator is rated of 40 Amps and 14 VDC, as the two voltage regulators. An automatic overvoltage device protects the circuits and the electric components from
an excessive voltage caused by generator failures. The power rating of each generator is such that if one generator fails the other one can still supply the airplane equipment to maintain flight safety.
Secondary DC power is provided by a main battery (lead type - 12 V, 23-Ah) and a secondary battery (lead type - 12V, 13 Ah).
An external DC power source can be connected to the aircraft distribution system in order to have it fed without starting the engine.
The ammeter section of the G1000 EIS can indicate the current supplied by either left or right generator switching a dedicated selector.
There are five different buses:
• Battery bus,
• LH Alternator bus,
• RH Alternatorbus,
• LH Avionics bus,
• RH Avionics bus.
The distribution system operates as a single bus with power being supplied by the battery and both generators but it is possible to separate the left busses from the right busses when required by means of the Cross Bus switches.
The switches to enable and disable the alternators and battery are grouped in the master switches group and are located in the centre side of the instrument panel. Only the emergency switch, that allow to put in parallel both batteries is located in left side of the instrument panel.
All electrical loads are divided among the five busses on the basis of their importance and required power: equipment with duplicate functions is connected to separate busses.
The Battery bus, which supplies the most important loads, is energized from three sources: the battery and both generators. This allows the bus for remaining active also in case of two independent faults in the supply paths.
- In addition, Emergency Light is connected directly on the battery.
On the central pedestal (see Figure below) there are seven switches disposed on two rows: on the first row there is the MASTER SWITCH which allows for connecting, through the battery relay, the battery to the battery bus.
LH and RH FIELD switches control the pertinent generator: setting the switch to OFF puts the pertinent generator off-line.
In correspondence of the second row there are 4 switches LH/RH AVIONIC and LH/ RH CROSS BUS.
The first two switches allow, through a relay, to cut off the power supply to the pertinent avionic bus.
The second ones allow, through a relay, for realizing the parallel connection between the pertinent generator bus and the battery bus. Setting these ones to OFF, the pertinent generator bus (and related avionic bus supplied) is separated from the battery bus and from opposite generator bus.
When both generators are correctly operating and all above mentioned switches are in ON position, all the busses are connected to the generators.
The ignition switches, two for each engine and grouped on the over head panel, are instead independent from the airplane electrical system (generation and distribution); they only control and open the engine electrical circuit.
This section contains supplemental information to operate, in a safe and efficient manner, the aircraft when equipped with S-TEC Fifty Five X autopilot device interfacing Garmin G1000 NXI.
The System Fifty Five X is a rate based autopilot. When in control of the roll axis, the autopilot senses turn rate, as well as closure rate to the selected course, along with the non-rate quantities of heading error, course error and course deviation indication.
When in control of the pitch axis, the autopilot senses vertical speed, acceleration, and closure rate to the selected glideslope, along with the non-rate quantities of altitude and glideslope deviation indication.
These sensed data provide feedback to the autopilot, which processes them in order to control the aircraft through the use of mechanisms coupled to the control system.
The “autotrim” function senses when the aircraft needs to be trimmed about the pitch axis, and responds by driving the trim servo in the proper direction to provide trim.
Following operating limitations shall apply when the aircraft is equipped with STEC Fifty Five X autopilot:
- Autopilot is certified for Category I – ILS Approaches [with a decision height not lower than 200 feet AGL (61m)]
- Autopilot operation forbidden with flaps extended more than TO position
- During Autopilot operation, a pilot with seat belt fastened must be seated at the left pilot position
- The use of Autopilot during single engine operation is forbidden
- Autopilot DISC during take-off and landing
- Maximum speed for Autopilot operation is 135 KIAS
- Minimum speed for Autopilot operation is 85 KIAS
- Minimum altitude AGL for Autopilot operation is:
- Cruise and Descent: 1000 ft
- Climb after takeoff and not precision approach: 400 ft
- ILS CAT I precision approach: 200 ft
On the instrument panel, in clear view of the pilot, it is placed the following placard reminding the observance of aircraft operating limitations during Autopilot operation:
The landing gear retraction system is of electro-hydraulic type, powered by a reversible pump which is electrically controlled by the LG control knob located on the LH instrument panel and by the legs position micro switches: these ones allow for detecting landing gear “down-locked” and “up” positions and for alerting the pilot by aural means should the approach and landing configuration be incorrect, in terms of flaps/throttle levers/landing gear position, in order to avoid an unintentional gear-up landing.
The three green lights illuminate only when the respective gear is “down-locked”; the red light indicates the gear is in transit “up” or “down” and the amber caution light GEAR PUMP ON indicates that the pump is electrically supplied.
The red transition light extinguishes only when all the three gear legs are “downlocked” or they are “up” while the amber caution light extinguishes only when the electrical pump is “off”.
The Up/Down limit switches control the LG lights lighting and pump operation on the basis of LG configuration set by the pilot through the LG control knob.
The A/C is provided with an independent hydraulically actuated brake system for each main wheel. A master cylinder is attached to each pilot/co-pilot’s rudder pedal: see schematic below.
Hydraulic pressure, applied via the master cylinders, enters the brake via lines connected to an inlet fitting on the wheel brake caliper.
A parking brake valve, mounted in correspondence of the cabin floor and operated by a knob on the cockpit central pedestal, intercepts the hydraulic lines, once the system is pressurized, to hold the brake assemblies linings tightened round the main wheels brake discs.
Brakes can be operated from both pilot’s and co-pilot’s pedals: a single vented oil reservoir feeds the pilot side master cylinders which are connected, via hoses, with the co-pilot’s side ones.
The cabin heating system utilizes hot air coming from engines heat exchangers: here cold ram-air is warmed by engine exhaust gases and then it is routed to the heating system hoses.
The cabin heat control knobs are positioned on the lower side of the LH instrument panel; when knobs are fully pulled, cabin receives maximum hot air. Left knob controls the warm air from LH engine heat exchanger, right knob controls the warm air from RH engine heat exchanger.
The baggage compartment is located behind the passengers’ seats. The baggage must be uniformly distributed on the floor and the weight cannot overcome 80kg.
The Engine Indication System can be accessed via ENGINE button on the MFD. The MFD then opens a page with quite extensive information regarding the Engine's status. A second push to the ENGINE button will close the page.