SECTION 7 AIRPLANE AND SYSTEMS DESCRIPTION

FOUND SECTION 7 SECTION 7 AIRPLANE AND SYSTEMS DESCRIPTION TABLE OF CONTENTS Page Introduction... 7-5 Airframe... 7-5 Flight Controls... 7-7 Aileron Control System... 7-7 Elevator Control System... 7-7 Rudder Control System... 7-8 Horizontal Stabilizer Control System (Pitch Trim) 7-8 Pitch Trim System... 7-9 Rudder Trim System... 7-10 Fowler Flaps... 7-11 Instrument Panel... 7-12 Ignition Switch Panel... 7-13 Flight Instrument Panel... 7-14 Overhead Switch Panel... 7-16 Engine Controls Panel... 7-17 Pitch Trim Panel... 7-18 Fuel Tank Selector Console... 7-19 Circuit Breakers Panel... 7-20 Right Hand Panel... 7-21 Garmin G500 Primary Flight / Multi-function Display System... 7-22 G500 System Power Sources... 7-23 G500 Navigation Sources... 7-24 G500 Synthetic Vision... 7-24 Autopilot Interface... 7-26 Audio Panel... 7-26 Traffic and Weather Systems... 7-26 ISSUE 2 7-1

SECTION 7 TABLE OF CONTENTS (Continued) FOUND Page Landing Gear System... 7-27 Main Gear... 7-27 Nose Gear... 7-27 Brake System... 7-28 Break Wear... 7-28 Parking Brake... 7-29 Engine and Systems... 7-30 New Engine Break-In and Operation... 7-30 Ignition System... 7-30 Oil System... 7-31 Cooling System & Cowl Flaps... 7-32 Induction System & Alternate Air Control... 7-33 Throttle Control... 7-34 Propeller Pitch Control... 7-34 Mixture Control... 7-34 Engine Monitoring System... 7-35 Main Engine Screen... 7-37 External Master Warning Light... 7-40 Voice Alarm... 7-40 Flight Data Screens... 7-41 Display Dimming... 7-44 Cleaning the Screen... 7-44 Propeller and Governor... 7-45 Fuel System... 7-46 Fuel System Schematic... 7-47 Fuel Distribution... 7-48 Fuel Venting... 7-48 Fuel Drain Valves... 7-49 Fuel Quantity Gauges... 7-49 Fuel Flow Rate and Fuel Remaining... 7-50 Fuel Tank Selector Valve... 7-51 Fuel Dipstick... 7-52 Auxiliary Fuel Pump Switch... 7-52 7-2 ISSUE 2

FOUND SECTION 7 TABLE OF CONTENTS (Continued) Page Electrical System and Instruments... 7-53 Electrical System Schematics... 7-54 Circuit Breakers... 7-55 Over- and Under-Voltage Annunciator Lights... 7-56 Master Switch... 7-57 Alternator Enable Switch... 7-57 Avionics Master Switch... 7-58 Backup Power Switch... 7-58 12V Accessory Outlet... 7-58 Ammeter/Voltmeter Gauge... 7-58 Turn Coordinator... 7-59 Pitot-Static System and Instruments... 7-60 Pitot Heat... 7-60 Standby Airspeed Indicator... 7-60 Standby Altimeter... 7-61 Alternate Static Source Switch... 7-61 Vacuum System and Instruments... 7-62 Vacuum (Suction) Gauge... 7-63 Standby Attitude Indicator... 7-63 Stall Warning System... 7-64 Stall Warning Horn and Light... 7-64 Lighting Systems... 7-65 Exterior Lighting... 7-65 Interior Lighting... 7-66 Instrument Panel Lighting... 7-67 Remote Emergency Locator Transmitter (ELT) Switch... 7-67 Magnetic Compass and Deviation Card... 7-68 Outside Air Temperature (OAT) Gauge... 7-68 Seats... 7-69 Pilot and Co-Pilot Seats (Fixed Height)... 7-69 Pilot and Co-Pilot Seats (Height Adjustable)... 7-70 Aft Passenger Seats... 7-71 ISSUE 2 REV 1 March 14, 2013 7-3

SECTION 7 FOUND TABLE OF CONTENTS (Continued) Page Doors... 7-72 Baggage Compartment... 7-73 Heating, Defrost, and Ventilation System... 7-74 Ventilation Controls... 7-74 Additional Ventilation Air... 7-76 Schematic... 7-77 Safety Equipment... 7-78 Pilot and Co-Pilot Seat Belt and Shoulder Harness... 7-78 Aft Passenger Seat Belts and Harnesses... 7-79 Fire Extinguisher... 7-80 Emergency First Aid Kit... 7-81 Control Lock... 7-82 24V Ground Power Plug (Optional)... 7-84 7-4 ISSUE 2

FOUND SECTION 7 INTRODUCTION Section 7 provides detailed descriptions of the airplane s systems, its system components, and their operation. The avionics equipment is unique to individual airplanes and is not described in this section. Refer to Section 9, Supplements, for details of the avionics equipment. AIRFRAME The airplane is a five-seat, high-wing, single-engine airplane equipped with tricycle landing gear. It is designed and built as a robust general aviation airplane. The forward section of the fuselage is a welded tubular steel space frame. This space frame is the backbone of the airplane. It has numerous reinforced weldments for attaching the wing, landing gear, aft fuselage, seats, engine mount and float struts. It is faired with composite skins and doors. The aft fuselage is a semi-monocoque design constructed of frames, longerons, stringers, and external aluminum skins. The one-piece cantilever wing has a straight leading edge and consists of a constant chord center section with outboard sections on either side with forwardswept trailing edges. The all-metal wing is constructed of formed aluminum ribs, spars, stringers, and is skinned with various thicknesses of aluminum sheeting. Three spars form the main structural members of the wing: a forward spar at 3% chord, a mid spar at 40% chord, and an aft spar at 74% chord. Each wing has a 50 U.S. gallon (189.3 litre) wet tank, located between the front and mid spars. The mid (main) spar carries through the cabin, while the other spars terminate at the fuselage sides. The wing is bolted to the steel fuselage at the front and mid spars. ISSUE 2 7-5

SECTION 7 FOUND The aileron skeleton consists of a forward spar and formed aluminum ribs. The skeleton is skinned with aluminum sheet riveted together with an internal trailing edge stiffener. The outer end of the aileron contains a lead-balance counterweight. The flaps are similarly constructed and are attached to the wing via stainless steel flap tracks and rollers. The empennage consists of a conventional vertical stabilizer with rudder, and a trimable horizontal stabilizer with elevator. The vertical stabilizer has two spars, formed aluminum ribs, stringers, aluminum skin panels, a formed leading edge, and a composite dorsal fairing. The rudder has a single spar, formed ribs, flat skins, and formed leading edge skin. The rudder attaches to the vertical stabilizer at three hinge locations. An electric trim tab is located at the base of the trailing edge of the rudder. A forward-pointing shielded horn counter-balance weight is incorporated into the top of the rudder. The horizontal stabilizer is a trimable surface with similar construction as the vertical stabilizer; having a forward and aft spar, ribs, stiffeners, skin panels, and formed leading edge skins. The elevators consist of formed leading edge skins, a single spar, formed ribs, with a trailing edge stiffener. The elevator outboard tips have an unshielded horn with a D-shaped counter-balance weight. Five hinges connect the elevator to the horizontal stabilizer. 7-6 ISSUE 2

FOUND SECTION 7 FLIGHT CONTROLS The airplane s flight control systems (refer to Figure 7-1) consist of conventional aileron, rudder, and elevator control surfaces. The control surfaces are manually operated via bellcranks, cables, and push rods. A control yoke actuates the ailerons and elevator. The rudder is controlled with rudder pedals. AILERON CONTROL SYSTEM ELEVATOR CONTROL SYSTEM Figure 7-1 Flight Control and Trim Systems (Sheet 1 of 2) ISSUE 2 7-7

SECTION 7 FOUND RUDDER CONTROL SYSTEM HORIZONTAL STABILIZER CONTROL SYSTEM (PITCH TRIM) Figure 7-1 Flight Control and Trim Systems (Sheet 2 of 2) 7-8 ISSUE 2

FOUND SECTION 7 PITCH TRIM SYSTEM A manually-operated trim system, used to control the aircraft s pitch, is provided with this airplane. This system uses the airplane s horizontal stabilizer to adjust the pitch trim by changing the stabilizer angle. Trimming is accomplished by rotating the vertically mounted trim wheel on the center console. Upward (forward) rotation of the trim wheel will trim the airplane s nose down. Conversely, downward (aft) rotation will trim the airplane s nose up. The amount of trim is measured on the scale to the left of the trim wheel from 6 up to 4 down. Recommended takeoff trim settings are marked by a green band on the trim indicator. The left edge of the band is for the forward center of gravity at maximum gross weight loading. The right boundary is for the aft center of gravity at maximum gross weight loading. ISSUE 2 7-9

SECTION 7 FOUND RUDDER TRIM SYSTEM The airplane is equipped with an electrically operated trim tab on the rudder. The pilot can change the rudder (yaw) trim by actuating a rocker-type switch. Pressing the left side of the switch initiates a Nose Left trim while pressing the right side initiates a Nose Right trim. An LED bar-type indicator located below the switch shows the position of the trim tab. The switch and indicator are located on the flight panel next to the flap position indicator. A Day/Night switch located next to the indicator allows the pilot to dim the indicator for night operation. 7-10 ISSUE 2

FOUND SECTION 7 FOWLER FLAPS Figure 7-2 Fowler Flaps The flaps are electro-mechanically extended by holding the flaps switch up or down to the desired angle. The angle can be finely adjusted and is indicated on the flap angle indicator at the top of the flight instruments panel. The indicator is marked at 10, 20, and 30 positions. Maximum flaps extension is 30. Any flap location from 0 to 30 may be selected. ISSUE 2 7-11

SECTION 7 FOUND INSTRUMENT PANEL The Instrument Panel (refer to Figure 7-3) is installed in modular sections to permit easy access to instrument clusters and switches without the need to remove the entire instrument panel. 1. Ignition Switch Panel 6. Right Panel 2. Flight Instrument Panel 7. Pitch Trim Panel 3. Engine Controls Panel 8. Overhead Switch Panel 4. Circuit Breaker Panel 9. Center Panel 5. Compass 10. Fuel Selector Console Figure 7-3 Instrument Panel 7-12 ISSUE 2

FOUND SECTION 7 IGNITION SWITCH PANEL For descriptions of the ignition switch, master switch and alternator enable switch, refer to Electrical System and Instruments in this section. 1. Master Switch 3. Alternator Enable Switch 2. Post Light 4. Ignition Switch Figure 7-4 Ignition Switch Panel ISSUE 2 7-13

SECTION 7 FOUND FLIGHT INSTRUMENT PANEL 1. Standby Airspeed Indicator 10. Flap Position Switch 2. Stall Warning Indicator 11. Autopilot Master Switch 3. Autopilot/Flight Director Annunciator 12. G500 Multi-function Display 4. Check Engine Light 13. G500 Primary Display 5. Voice Alarm Switch 14. MVP-50 Engine Monitor 6. Day/Night Dimmer Switch 15. Parking Brake Handle 7. Rudder Trim Switch 16. Vent 8. Rudder Trim Position Indicator 17. Standby Altimeter 9. Flap Position Indicator 18. Standby Attitude Indicator Figure 7-5 Flight Instrument Panel 7-14 ISSUE 2

FOUND SECTION 7 For a description of the standby attitude indicator, refer to Vacuum System. For a description of the stall warning horn and light, refer Stall Warning System. For descriptions of the standby airspeed indicator and standby altimeter, refer to Pitot-Static System and Instruments. ISSUE 2 7-15

SECTION 7 FOUND OVERHEAD SWITCH PANEL 1. Dimmer Switches 3. Cowl Flap Position Indicator 2. Upper Deck Function Switches 4. Lower Deck Function Switches Figure 7-6 Overhead Switch Panel The Switch Panel is located on the ceiling on the centreline of the airplane. It contains two levels. The upper level contains the following switches and dimmer control knobs, from left to right: DIMMER CONTROLS: The dimmer control knobs govern the intensity of the lighting for the control panels. PITOT HEAT BACKUP POWER COWL FLAP The lower level of the panel contains the following switches, from left to right: FUEL PUMP (AUXILIARY) AVIONICS MASTER 7-16 ISSUE 2

FOUND SECTION 7 NAVIGATION LIGHTS STROBE LIGHTS LANDING LIGHT AFT CABIN LIGHT CREW LIGHT ENGINE CONTROLS PANEL 1. Throttle Control 4. Post Light 2. Propeller Pitch Control 5. Ventilation Levers 3. Mixture Control 6. Alternate Air Induction Control Figure 7-7 Engine Controls Panel For a description of the engine instruments and controls, refer to Engine and Systems section. For a description of the ventilation controls (heating, cooling & defrost), refer to Heat, Defrost, and Ventilation section. ISSUE 2 7-17

SECTION 7 FOUND PITCH TRIM PANEL 1 1. Alternate Static Source Switch 3. Pitch Trim Control 2. Remote ELT Switch 4. Pitch Trim Position Indicator Figure 7-8 Pitch Trim Panel For descriptions of the pitch trim wheel and trim scale, refer to Flight Controls section. For a description of the alternate static source switch, refer to Alternate Static Source. 7-18 ISSUE 2

FOUND SECTION 7 FUEL TANK SELECTOR CONSOLE The Fuel Tank Selector Console (refer to Figure 7-9) is installed on the floor between the pilot and co-pilot seats. For a description, refer to Fuel System Management. Figure 7-9 Fuel Tank Selector Console ISSUE 2 7-19

SECTION 7 FOUND CIRCUIT BREAKER PANEL 2 1. Post Light 5. Circuit Breakers (Third Row) 2. Circuit Breakers (First Row) 6. Circuit Breakers (Fourth Row) 3. Circuit Breakers (Second Row) 7. Under Volt Indicator 4. 12V Plug 8. Over Volt Indicator Figure 7-10 Avionics and Circuit Breaker Panel For a description of the circuit breakers, refer to Electrical System and Instruments. 7-20 ISSUE 2

FOUND SECTION 7 RIGHT HAND PANEL 2 1. Avionics Bay 3. Access Panel 2. Vent Figure 7-11 Right Hand Panel Avionics for the are custom ordered for each airplane. For complete operating instructions on this airplane s avionics systems, refer to the original equipment manufacturer s manuals and documentation provided with this airplane. Also note that the Center Panel, located directly to the left of the Right Hand Panel s avionics bay (Not shown here), contains space for additional avionics equipment. ISSUE 2 7-21

SECTION 7 FOUND GARMIN G500 PRIMARY FLIGHT / MULTI-FUNCTION DISPLAY SYSTEM The G500 PFD/MFD System consists of a Primary Flight Display (PFD) and Multi-Function Display (MFD) housed in a single Garmin Display Unit (GDU), plus an Air Data Computer (ADC) and Attitude and Heading Reference System (AHRS). The G500 interfaces with other installed systems in the aircraft, including Garmin GNS and GTN series GPS/WAAS navigators, Garmin GDL 69 data link radios, and various audio panels, traffic systems and ADF navigators. The primary function of the PFD is to provide attitude, heading, air data and navigation information (from GNS units) to the pilot. The primary function of the MFD is to provide mapping, terrain, and flight plan information. The standby instruments (altimeter, airspeed, attitude, and magnetic compass) are completely independent from the PFD and will continue to operate in the event the PFD is not usable. These standby instruments should be included in the pilot s normal instrument scan and may be referenced if the PFD data is in question. Garmin Display Unit (GDU) 7-22 ISSUE 2

FOUND SECTION 7 G500 SYSTEM POWER SOURCES The G500 system depends on electrical power to maintain proper operation. The Garmin Display Unit (GDU), Attitude and Heading Reference System (AHRS), and Air Data Computer (ADC) are directly tied to the aircraft s main bus and energized when the aircraft master switch is turned on. Other systems, like the navigation equipment, weather datalink, and autopilot are located on the avionics bus and are not operable during engine start. The major components of the G500 are circuit breaker protected with reset-able type breaker available to the pilot. These breakers are located on the circuit breaker panel and labelled as follows: 1. PFD - Garmin Display Unit (PFD/MFD), GDU 620 2. AHRS - Attitude and Heading Reference System, GRS 77 3. ADC - Air Data Computer, GDC 74A G500 NAVIGATION SOURCES The G500 requires at least one Garmin GPS/WAAS navigation unit to ensure the integrity of the Attitude and Heading Reference System. The AHRS will still operate in a reversionary mode if the GPS fails, and the PFD attitude display will still be presented, see Chapter 2 for limitations as a result of a GPS failure. The G500 HSI can be selected to display Course deviation information from up to four independent sources: two GPS, and two VHF NAV. In addition, the HSI can display two simultaneous bearing pointers sourced from GPS, VHF NAV, or ADF. ISSUE 2 7-23

SECTION 7 FOUND G500 SYNTHETIC VISION TECHNOLOGY (IF INSTALLED) SVT uses an internal terrain database and GPS location to present the pilot with a synthetic view of the terrain in front of the aircraft. The purpose of the SVT system is to assist the pilot in maintaining situational awareness with regard to the terrain and traffic surrounding the aircraft. A typical SVT display is shown below: SVT provides additional features on the G500 primary flight display (PFD) which display the following information: Synthetic Terrain; an artificial, database derived, three dimensional view of the terrain ahead of the aircraft within a field of view of approximately 25 degrees left and 25 degrees right of the aircraft heading. Obstacles; obstacles such as towers, including buildings over 200 AGL that are within the depicted synthetic terrain field of view. Flight Path Marker (FPM); an indication of the current lateral and vertical path of the aircraft. The FPM is always displayed when synthetic terrain is selected for display. Traffic; a display on the PFD indicating the position of other aircraft detected by a traffic system interfaced to the G500 system. Horizon Line; a white line indicating the true horizon is always displayed on the SVT display. Horizon Heading; a pilot selectable display of heading marks displayed just above the horizon line on the PFD. Airport Signs; pilot selectable signposts displayed on the synthetic terrain display indicating the position of nearby airports that are in the G500 database. Runway Highlight; a highlighted presentation of the location and orientation of the runway(s) at the destination airport. 7-24 ISSUE 2

FOUND SECTION 7 The synthetic terrain depiction displays an area approximating the view from the pilot s eye position when looking directly ahead out the windshield in front of the pilot. Terrain features outside this field of view are not shown on the display. The synthetic terrain display is intended to aid the pilot awareness of the terrain and obstacles in front of the airplane. It may not provide either the accuracy or ISSUE 2 7-25

SECTION 7 FOUND fidelity, or both, on which to solely base decisions and plan maneuvers to avoid terrain or obstacles. The synthetic vision elements are not intended to be used for primary aircraft control in place of the primary flight instruments. AUTOPILOT INTERFACE (IF INSTALLED) The G500 may be interfaced to an optional autopilot. The G500 typically provides course and heading datum to the autopilot based on the data selected for display on the HSI. For multiple GPS/NAV systems, the G500 acts as a selection hub for the autopilot s NAV mode, and the G500 may also provide GPS Steering data. The autopilot may provide Flight Director capabilities which can be displayed on the G500 Attitude Indicator as a Single Cue Flight Director. AUDIO PANEL The G500 PFD/MFD system is interfaced to the aircraft audio panel to provide aural alerting generated by the G500. TRAFFIC AND WEATHER SYSTEMS (IF INSTALLED) The G500 PFD/MFD system supports TIS traffic via the Garmin GTX Series Mode-S Transponders. The system also supports TAS/TCAS/TIS traffic from various active traffic awareness systems. The information from these systems is available and controllable on the MFD. The G500 PFD/MFD system supports XM datalink weather via the Garmin GDL69 and GDL69A receivers. If an optional XM datalink receiver is installed, the pilot will be able to access graphical and text weather products on the MFD and control the audio entertainment data from the MFD while listening via an appropriately installed audio panel. 7-26 ISSUE 2

FOUND SECTION 7 LANDING GEAR SYSTEM The airplane s landing gear is tricycle arrangement. MAIN GEAR Figure 7-12 Landing Gear The main gear legs are tapered steel tubes. The legs are attached to machined aluminum trunnions that are secured to bearing housings bolted to the fuselage. A steel tube running across the bottom of the fuselage limits the amount rotation of the trunnions and legs. The flexing of the cross tube and the tapered legs provide the shock absorption for the main gear. The main gear wheels have 6.00-6 tires with inner tubes. The wheels and tires are enclosed in wheel pants that are easily removable for maintenance. Plugs located on the wheel pants provide access to the valve stem for tire inflation and pressure checking. The tire pressure should be 42 to 44 psi. Each main gear wheel is equipped with a hydraulically-actuated, disc brake on the inboard side of each wheel. ISSUE 2 7-27

SECTION 7 FOUND NOSE GEAR The nose gear strut is a tapered steel tube. It is bolted to a welded steel tube A- frame that is pinned to the engine mount. The strut and frame are limited in their rotation by a stack of elastomer pucks that are attached to the engine mount. The nose wheel is free castering and can turn through an arc of 216 degrees (108 per side). Steering is accomplished by differential application of the main gear brakes. The nose wheel tire is a 5.00-5 tube type tire and it is enclosed in a wheel pant as on the main gear. A small plug on the side of the wheel pant allows access to inflate the tire and check the pressure. The tire pressure should be 52 to 54 psi. BRAKE SYSTEM Each main wheel has a single, hydraulic disc brake. A hydraulic line from each brake is connected to a master brake cylinder attached to each of the pilot s rudder pedals. The brakes are operated independently by applying pressure to the brake pedals. Pressure applied evenly to both brake pedals will provide even braking on both wheels. Differential braking can be used to help steer the airplane by applying pressure to one brake only. BRAKE WEAR Some indicators of poor brake condition and/or impending brake failure are: Gradual decrease in braking action (fading) after brake application Noisy or dragging brakes Soft or spongy pedals Excessive brake pedal travel plus weak braking action If any number of these symptoms begins to appear, the brake system requires immediate attention. If, during taxi or landing roll, the brakes fade, let up on the pedals and then re-apply the brakes with heavy pressure. If the brakes become spongy or pedal travel increases, pump the pedals to build brake pressure. If one brake becomes weak or fails, use the other brake sparingly while using opposite rudder, as required, to offset the good brake and maintain directional control. Have the brake system repaired immediately. 7-28 ISSUE 2

FOUND SECTION 7 PARKING BRAKE PARKING BRAKE OFF PARKING BRAKE ON Figure 7-13 Parking Brake Off (Left) and On (Right) When the airplane is parked, both main wheel brakes may be set using the parking brake handle located below the lower edge of the Flight Instrument Panel. To apply the parking brake, depress both brake pedals, and pull out the parking brake handle (while depressing the button in the center of the handle). To release the brakes, depress the brake pedals and push the parking brake handle forward (again while depressing the button in the center of the handle). ISSUE 2 7-29

SECTION 7 FOUND ENGINE AND SYSTEMS The airplane is powered by a six-cylinder, direct-drive, horizontally opposed, aircooled, fuel-injected engine manufactured by Textron Lycoming. This 583 cubic inch displacement engine is a Lycoming IO-580-B1A model and is rated at 315 horsepower at 2700 RPM. Major accessories include a starter, a beltdriven alternator mounted on the front of the engine, dual magnetos, single vacuum pump, and a full-flow oil filter mounted on the aft of the engine accessory case. For locations of engine controls and instruments, refer to Engine Instruments and Controls Panel. NEW ENGINE BREAK-IN AND OPERATION New engines have been carefully broken-in by the manufacturer and, therefore, no further break-in is necessary. New or newly-overhauled engines should be operated using only lubricating oils recommended in the latest Textron Lycoming Service Instruction No. 1014. To ensure proper seating of the piston rings, cruise power should be limited to 65% to 75% until a total of 50 hours has been accumulated or the oil consumption has stabilized. IGNITION SYSTEM The ignition switch is located on the Ignition Panel and is a rotary switch requiring a key. This switch is labelled clockwise: OFF, R, L, BOTH, and START. The engine should be operated on both magnetos at all times except during magneto checks. With the MASTER switch ON, the engine starter can be engaged by rotating the ignition switch to the spring loaded START position. In this position, the starter contactor is energized and the starter motor will crank the engine. When the switch is released, it will automatically spring back to the BOTH position. 7-30 ISSUE 2

FOUND SECTION 7 OIL SYSTEM Engine lubrication is achieved using a wet sump lubrication system. This pressure-operated lubrication system uses aviation grade oil as per the latest Textron Lycoming Service Instruction No. 1014. The engine sump, located at the bottom of the engine, has a maximum capacity of 10.4 litres (11 U.S. quarts). The lubrication system operates by drawing oil from the sump through an oilsuction strainer screen and into the engine-driven oil pump. Oil is routed to a bypass valve, which enables the oil to bypass the oil cooler when operating at cold temperatures and pass directly from the oil pump to the full-flow oil filter. Hot oil passes out of the accessory housing to the oil cooler on the engine's port side of the firewall. Pressurized oil from the oil cooler returns to the accessory housing by passing through the full-flow oil filter. The filtered oil then enters the oil relief valve, which maintains engine oil pressure within specified limits. The oil pressure is regulated by allowing excess oil to return to the oil sump. The remaining oil is circulated at proper pressures throughout the engine. Residual oil within the crankcase is returned to the sump by gravity. The filler cap/dipstick is accessible through an access door on the top port side of the engine cowling. The engine should not be operated with less than 4 litres (4 U.S. quarts) of oil in the sump. For normal flights fill to 9 litres (9 U.S. quarts) (dipstick indication only) and for extended flight, fill to 10.4 litres (11 U.S. quarts). For engine oil grade and specifications, refer to Section 8. ISSUE 2 7-31

SECTION 7 FOUND COOLING SYSTEM AND COWL FLAPS Cooling for the engine cylinders and the oil is provided by ram air that enters through intake openings on the nose of the cowl. Baffles on top of engine direct the cool ram air through the cylinder cooling fins and the resulting warm air exits the engine compartment through openings at the bottom of the cowl. Ram air is also directed into a duct at the back of the engine on the left side. This duct leads to the oil cooler which is mounted on the engine mount. The warm air from the oil cooler exits the engine compartment in the same manner. The exit on the bottom of the cowl consists of the area around the tail pipes and the cowl flaps. The cowl flap system consists of a two flaps each actuated by an electric actuator. The actuators and flaps are designed to function simultaneously. Should one actuator fail the other will continue to operate providing a level of redundancy. The actuators are pilot controlled and may be selected to fully open, fully close or stop at an intermediate position as required. The cowl flap switch is located on the overhead switch panel. An LED bar-type indicator located next to the switch indicates the position of the cowl flaps. The indicator may be dimmed for night operation using the Day/Night dimmer switch located on the flight panel by the rudder trim indicator. CLOSE COWL OPEN COWL FLAP CLOSE OPEN Figure 7-14 Cowl Flap Switch & Indicator on Overhead Switch Panel 7-32 ISSUE 2

FOUND SECTION 7 INDUCTION SYSTEM AND ALTERNATE AIR CONTROL The Induction System receives ram air through a circular intake on the cowl just below the propeller. A conical air filter inside the intake prevents dust and other foreign matter from entering the engine. After passing through the filter, induction air enters a fuel/air control unit under the engine. It is then ducted to the engine cylinders through the intake manifold tubes. In the event of loss of manifold pressure due to icing or blocking of the air filter a pilot controlled alternate air door may be opened. When the alternate air door is open the normal induction air system is bypassed and the engine draws unfiltered air from a warm section of the engine compartment. The alternate air door is operated by a push-pull cable. The control knob is located just below the throttle control. The alternate air control has a square knob to differentiate it from round throttle knob. The alternate air door is open when the knob is in the full aft position. Operation using alternate air should be minimized as the engine is drawing unfiltered air. The cause of the filter blockage should be corrected as soon as practical. THROTTLE CONTROL ALTERANATE AIR CONTROL Figure 7-15 Alternate Air Control ISSUE 2 7-33

SECTION 7 FOUND THROTTLE CONTROL The black throttle control regulates the engine s manifold pressure and engine power. The throttle is open (full power) in the full forward position and closed (to idle) in the full aft position. A round knurled friction lock is located at the base of the throttle and is operated by rotating the lock clockwise to increase friction or counter clockwise to decrease it, locking/unlocking the throttle at any desired power setting. PROPELLER PITCH CONTROL setting. The blue propeller pitch control with center release button increases the propeller rpm (revolutions per minute) when in the forward position and decreases the propeller rpm when in the aft position. For fine adjustments, the Vernier-type pitch control may be rotated to produce the desired rpm MIXTURE CONTROL The red fuel-to-air mixture control with center release button regulates the amount of fuel flow to the engine. The engine rich fuel position is full forward, and the idle (full fuel lean) cut-off position is full aft. The mixture control is operated in the same manner as the Vernier-type pitch control described above. 7-34 ISSUE 2

FOUND SECTION 7 ENGINE MONITORING SYSTEM The airplane is equipped with an Electronics International MVP-50 engine analyzer and system monitor. This system will monitor and display the following parameters: Engine: Manifold Pressure Engine RPM (tachometer) Oil Temperature Oil Pressure Cylinder Head Temperature (CHT for all six cylinders) Exhaust Gas Temperatures (EGT for all six cylinders) Fuel Flow Fuel Level (Left and Right Tanks) Fuel Remaining % Horsepower (advisory) System: Cylinder Shock Cooling Time (Local and Zulu) Volts & Amps Vacuum CAUTION The pilot must understand the operation of this monitor before flying the aircraft. Do not allow anyone to operate the aircraft that does not know the operation of this monitor. It is possible for any instrument (sensor) to fail thereby displaying inaccurate high, low or jumpy readings. Therefore, you must be able to recognize an instrument failure and you must be proficient in operating your aircraft safely in spite of an instrument failure. The position of the throttle, pitch and mixture are to be used as backup for engine condition. There are NO APPROVED checklists, flight plans or general information files for this instrument. The MVP-50 allows the pilot to enter checklists, flight plans and general information through the USB port. This data must be verified by the pilot before it is used. ISSUE 2 7-35

SECTION 7 FOUND CAUTION Do not rely solely on the fuel level displayed on the MVP-50 to determine the fuel levels in the airplane. The use of the MVP-50 does not eliminate or reduce the necessity for the pilot to use good flight planning, pre-flight and in-flight techniques for managing fuel. It is important the pilot adopt the practices listed below. If you are not familiar with these techniques, contact a flight instructor to acquire proper training. 1. Flight Planning - Always calculate the fuel requirement for each leg of a flight, including any alternate plans for bad weather. Keep this information available in the aircraft during the flight. See Chapter 5 for the published fuel flows for various flight/engine conditions of the aircraft. 2. Preflight - Do not rely on the MVP-50 to determine the fuel level in the fuel tanks. The pilot must visually check/measure the fuel levels in the tanks before every takeoff. Crosscheck the measured fuel levels with the displayed levels on the MVP- 50. Also, crosscheck these levels with the fuel requirements for the planned flight. 3. In Flight - Make the MVP-50 part of your normal instrument scan. Crosscheck the fuel levels displayed on the MVP-50 with your flight plan at each leg of the flight or every 30 minutes (if a leg is longer than 30 minutes). Calculate the fuel flows from the MVP-50 displayed fuel levels and compare them with your charts of measured and published fuel flows for the aircraft. If there is a discrepancy, land the aircraft at the nearest airport and verify the fuel levels. Discrepancies should be taken seriously. 7-36 ISSUE 2

FOUND SECTION 7 MAIN ENGINE SCREEN The Main Engine Screen provides the aircraft system and engine instruments the pilot will view most frequently during a flight. The parameters and layout of this screen are factory-set and cannot be changed. Figure 7-16 Engine Monitor Display (Main Engine Screen) Power-up Add Fuel Message: An Add Fuel Message located in the engine analyzer section of the screen will appear when the MVP-50 is powered up. The purpose of this message is to remind the pilot to update the fuel computer when adding fuel to the aircraft. The MVP-50 not only measures and displays the fuel in the fuel tanks but it also displays the fuel onboard the aircraft calculated from the fuel flow. This allows the pilot to cross check fuel readings to insure accuracy. Consult the MVP-50 operator s manual for instructions on updating fuel information. Note As long as the Add Fuel Message is displayed all the annunciators will be lit. This feature provides a test for the annunciators. ISSUE 2 7-37

SECTION 7 FOUND Main Engine Screen Layout The Main Engine Screen is laid out in four areas: RPM and Manifold Pressure: The RPM and Manifold Pressure instruments are located at the top left of the screen. Each of these instruments incorporates a large arc and digital display. Beneath the Manifold Pressure are digital readouts of the % Horsepower and the shock cooling in both Fahrenheit and Celsius. Beneath the RPM is the local and Zulu time. Horizontal Strip and Digital Gauges: Oil Pressure and Oil Temperature, presented as Horizontal Strip Gauges (with digital readouts) and three digital instruments, (VOLTS, AMPS, VACUUM), are located on the right side of the screen. Engine Analyzer: The engine analyzer is located on the lower left portion of the Main Engine Screen. The engine analyzer monitors both EGTs and CHTs. The engine analyzer provides both bar graph and digital formats and incorporates features for leaning, detecting and diagnosing engine problems. Consult the MVP-50 operator s manual for instructions on using these features. Fuel Instruments: The digital instruments on the blue background, located in the middle of the display, are dedicated to fuel instruments. They are as follows: FUEL L Fuel Level in Left Tank FUEL R Fuel Level in Right Tank F.REM Fuel Remaining in Both Tanks F.FLOW Fuel Flow Rate RPM and Manifold Pressure (Details): The RPM and Manifold Pressure instruments incorporate a digital readout and an analog arc. The colour of the digital readout will reflect the current operating level of the instrument (i.e., if the RPM is operating in the red, the digital readout will be red). The digital display is set to blink when the RPM operating level exceeds 2700 RPM, the red operating range. To stop the blinking, push any button, or rotate the SELECT knob. Also, acknowledging a voice warning using the external Voice Alarm Control Panel will stop the blinking of any digital display. The Manifold Pressure gauge has no programmed limits. The MVP-50 s RPM Instrument provides a Mag Out feature in addition to the arc and digital display. The MVP-50 continually monitors both mag signals. If one mag fails, an appropriate L. Mag Out or R. Mag Out warning will be displayed on the appropriate side of the RPM digital display. 7-38 ISSUE 2

FOUND SECTION 7 Horizontal Strip and Digital Gauges (Details): The Horizontal Strip and Digital Gauges provide the following features: A. The coloured operating ranges shown on the Horizontal Strip are programmed for your aircraft and may not be changed. See Chapter 2 of this manual for a summary of the operating ranges and various display colours for any instrument with limitations displayed on the MVP-50 screen. B. Each Horizontal Strip Gauge features an arrow indicating its current operating level. Also, the arrow allows you to interpret rate and trend information and provides field of vision data. C. A digital display is featured with each Horizontal Strip Gauge. D. The digital display will blink when a function s operating level reaches a red operating range. To stop the blinking, push any button, or rotate the Select knob. Also, acknowledging a voice warning using the external Voice Alarm Control Panel will stop the blinking of any digital display ISSUE 2 7-39

SECTION 7 FOUND EXTERNAL MASTER WARNING LIGHT A red external Master Warning Light is mounted on the flight panel in front of the pilot. This warning light provides a heads-up visual warning. The Master Warning Light will blink any time the operating level of any monitored function reaches a red operating limit. Pushing any button, or rotating or pushing the Select knob when the Master Warning Light is blinking will acknowledge the blinking (visual warning) and the blinking will stop. Also, acknowledging a voice warning using the external Voice Alarm Control Panel will stop the blinking of the Master Warning Light. Acknowledging a red blinking display will cause the Master Warning Light to stop blinking and go solid red. If another parameter reaches its red operating limit while the light is solid red it will begin to blink once again. This light will only go out when all the parameters in question return to their normal operating ranges. The Master Warning Light may be dimmed during night operation using the Day/Night dimmer switch located next to the Voice Alarm Control. VOICE ALARM The MVP-50 is equipped with a Voice Alarm System. It is a powerful system that provides an immediate and intelligent audible warning regardless of the pilot s head position or focus. The instant an operating level of any function reaches a red operating limit, a chime will sound in the headset and a pleasant female voice will annunciate a phrase, such as: Check Oil Pressure. Power-up Announcement: When the MVP-50 is powered up and the Voice Alarm Control switch is placed in the ON position, the MVP-50 will announce, Voice Annunciator enabled. Have a nice flight. This announcement will be made only once, at the beginning of each flight. The Voice Alarm Control switch is located on the light panel. Acknowledging and Silencing an Alarm for One Minute: To acknowledge and silence an alarm, push the switch on the Voice Alarm Control Panel momentarily to the ACK (acknowledge) position. A high tone beep will be heard in the headset and all active alarms will be silenced for one minute. This is handy if you don t want to permanently shut off any alarms but you need silence for one minute in order to deal with other pressing matters. After one minute the silenced alarms (if still active) will once again be announced in the headset. During the time one or more alarms are 7-40 ISSUE 2

FOUND SECTION 7 silenced, any newly activated alarm will be announced immediately in spite of the minute of silence. To silence this new alarm, once again push the Control Switch once momentarily to the ACK position, which will silence the new alarm and all active alarms for one minute. Acknowledging and Silencing an Alarm for 10 Minutes: To acknowledge and silence any active alarm for 10 minutes, push the Voice Alarm Control switch to the ACK position three times within three seconds or less. On the third push, a low tone boop will be heard in the headset, indicating that all active alarms will be silenced for 10 minutes. Turning the Voice Warning System OFF : To disable the MVP-50 Voice Warning System, silence all voice alarms in the headset and reset any silence delay times, set the Voice Alarm Control Panel switch to the OFF position. When the Control Panel Switch is once again set to the ON position, the MVP-50 will announce Voice Annunciator enabled. This will be followed by the announcement of any active alarms. Adjusting the Volume of the Voice Warnings: The Voice & Display Controls screen provides a control to adjust the volume level of the voice warnings. To navigate to the Voice & Display Control screen start by viewing the Main Engine Screen, push the Menu button and select the Voice & Display Control screen near the bottom of the page. Task List: If two or more alarms are activated, the alarms are placed on a task list and are announced one at a time with a one-second delay between alarms. After the last alarm on the task list is announced a five-second delay will occur. The alarms then once again announced in order. FLIGHT DATA SCREENS Besides the Main Engine Screen the MVP-50 can also display other screens such as checklists, flight plans, weight & balance, and fuel management. These screens are displayed against a blue background and are displayed on the left side of the MVP-50 screen. The engine instruments are displayed in a digital format on the right side of the screen against a black background as shown in the following figure. ISSUE 2 7-41

SECTION 7 FOUND Figure 7-17 Engine Monitor Display (Flight Data Screen) At the top of the engine instruments is an annunciator, for the Main Screen. If any function on the Main Screen goes into the red, the annunciator will blink, the external warning light will blink and an appropriate voice warning will be played. In this way the pilot is immediately alerted of a potential problem and should view the Main Engine Screen for further information. Consult the MVP-50 operator s manual for instructions on how to program these blue screens. CAUTION These screens are not preset by Found Aircraft Canada. It is the responsibility of the pilot to ensure that the information displayed on these blue screens is correct. 7-42 ISSUE 2

FOUND SECTION 7 SELECT Knob and Button Operation SELECT Knob: The SELECT knob can be rotated or pressed. Depending on the screen and field being viewed, rotating the knob can move an arrow, select a digit, or change a digits value. Pressing the SELECT knob will choose the highlighted item. EXIT Button: Pressing the EXIT button always returns you to the screen just prior to your pressing the SELECT knob. You can always return to the Main Engine Instrument Screen by repeatedly pressing the EXIT button. SCREENS Button: Pressing the SCREENS button sequences the MVP through a list of Flight Data Screens. The Flight Data Screens that will be displayed can be pre-selected on the Screens Button Setup page. The SCREENS button allows you quick access to the screens you use most frequently. To navigate to the Screens Button Setup page start by viewing the Main Engine Screen, push the Menu button and select the Screens Button Setup near the bottom of the page. MENU Button: The MENU button displays the menu for the current screen viewed, if a menu is available for that screen. Some screens do not have a menu available. When the MENU button is pushed while displaying the Main Engine Screen, the Flight Data Screens Menu will be displayed. ISSUE 2 7-43

SECTION 7 FOUND DISPLAY DIMMING The Brightness Control is set for Auto Dimming the display will automatically dim as the ambient light reduces. The light sensor is located next to the USB port on the MVP-50 front panel. The Brightness Control can be found on the Voice & Display Controls screen. To navigate to the Voice & Display Controls screen start by viewing the Main Engine Screen, push the Menu button and select the Voice & Display Controls screens near the bottom of the page. CLEANING THE SCREEN The MVP-50 incorporates a flat panel full colour TFT display. The TFT display should be cleaned using only isopropyl alcohol and a soft cleaning cloth. Individually wrapped lens-cleaning tissue (used to clean glasses or plastic lenses) works best. 7-44 ISSUE 2

FOUND SECTION 7 PROPELLER AND GOVERNOR This airplane is equipped with an aluminum Hartzell HC-C3YR-1RF/F8068 82 inch diameter, three-bladed, swept-tip, constant-speed, non-feathering propeller. The corresponding propeller governor is a Hartzell V5-4 designated governor. The hub shell is made in two halves, bolted together along the plane of rotation. The hub shell carries the pitch change mechanism and blade roots internally. The hydraulic cylinder, provides power for changing the pitch, and is mounted at the front of the hub. Oil pressure from a governor is used to move the blades into high pitch (reduced rpm). The centrifugal, twisting motion of the blades tends to move them into low pitch (high rpm) in the absence of governor oil pressure. The propeller pitch is controlled by the propeller pitch control located on the Engine Controls Panel. ISSUE 2 7-45

SECTION 7 FOUND FUEL SYSTEM The airplane s Fuel System is based on gravity feed to a pump-type fuel flow configuration with a sealed, integral fuel tank in each wing (see Figure 7-18). The fuel system consists of the following major components: Two (2) vented integral fuel tanks (one tank in each wing) Two (2) Float-type fuel quantity sensors (one in each wing tank) Two (2) fuel collector tanks Three-position fuel selector/shutoff valve (OFF, LEFT, and RIGHT) Fuel strainer (also known as a gascolator) Auxiliary electric fuel pump Engine-driven fuel pump (part of the engine) Fuel/air control unit (part of the engine) Fuel distribution valve (part of the engine) Fuel injection nozzles (part of the engine) The total fuel capacity of each tank is: FUEL LEVEL (QUANTITY EACH TANK) TOTAL FUEL CAPACITY TOTAL UNUSABLE FUEL TOTAL USABLE ALL FLIGHT CONDITIONS 50 USG 100 USG 1.7 USG 98.3 USG 189 Litres 378 Litres 6.4 Litres 372 Litres USG = US Gallons WARNING UNUSABLE FUEL LEVELS FOR THIS AIRPLANE WERE DETERMINED IN ACCORDANCE WITH TRANSPORT CANADA STANDARDS. FAILURES TO OPERATE THE AIRPLANE IN COMPLIANCE WITH FUEL LIMITATIONS SPECIFIED IN SECTION 2 MAY FURTHER REDUCE THE AMOUNT OF FUEL AVAILABLE IN FLIGHT. 7-46 ISSUE 2

FOUND SECTION 7 Figure 7-18 Fuel System Schematic ISSUE 2 7-47

SECTION 7 FOUND FUEL DISTRIBUTION There are two main fuel tanks (left and right) within the main wing. Each tank feeds fuel into their respective collector tanks by a combination of gravity and suction from the engine-driven pump. The collector tanks are located in a compartment under the floor beneath the crew seats. Access to this compartment is via a large access panel on the belly of the airplane. Each collector tanks has a capacity of 5 litres (1.32 U.S. gallons). Fuel from both header tanks passes through the three-position selector/shutoff valve, labelled LEFT, RIGHT, and OFF. WARNING THE FUEL IS PICKED UP FROM THE INBOARD END OF THE WING TANKS. ACCORDINGLY, IN PROLONGED SKIDS OR SLIPS, AS WELL AS PROLONGED UNCOORDINATED TURNS, THE FUEL SELECTOR VALVE SHOULD BE POSITIONED TO THE TANK PROVIDING THE MOST SUPPLY, OTHERWISE FUEL STARVATION OF THE ENGINE CAN RESULT. The selector valve determines which collector tank (and corresponding main tank) will feed fuel. Fuel then flows through the fuel strainer, past the auxiliary fuel pump, and into the engine-driven fuel pump. Fuel is delivered to the fuel/air control unit, where it is metered and directed to a fuel distribution valve (manifold), which distributes it to each cylinder. Fuel flow into each cylinder is continuous, and the flow rate is determined by the amount of air passing through the fuel/air control unit. FUEL VENTING Proper Fuel System venting is essential for reliable and safe engine operation. Blockage of the Fuel System results in decreased fuel flow and eventual engine stoppage. Since each main wing fuel tank operates independently of the other, each tank is fitted with overboard vent lines which vent at the outboard end of the fuel tank. The vent line outlets protrude at mid-chord on the lower wing surface at approximately the semi-span of each wing. 7-48 ISSUE 2

FOUND SECTION 7 FUEL DRAIN VALVES The Fuel System is equipped with quick-drain valves to provide an easy means for the inspection of fuel in the fuel tanks. Fuel contamination and grade can be verified by draining and inspecting all fuel drain valves and must be done prior to each flight and after each refueling. Fuel can be drained from two drains for each wing tank, each collector tank, and the drain on the fuel strainer a total of seven fuel drains (Refer to Pre-flight Checklist, Section 4 Normal Procedures). If any evidence of fuel contamination is found, it must be eliminated in accordance with the Pre-Flight Checklist and the procedures in Section 8 of this publication. If takeoff weight limitations for the next flight permit, the fuel tanks should be filled after each flight to prevent internal condensation. FUEL QUANTITY GAUGES The fuel quantity of each main fuel tank is measured by a float-type fuel quantity sensor (one in each tank) and the level is indicated by digital readouts on the MVP-50 engine monitor. When the fuel level reaches 0 in either tank the digital readout on the MVP-50 for the applicable tank will turn red and begin blinking. The voice alarm system on the MVP-50 will also initiate an audible warning. To temporarily silence the audible warning and stop the readout from blinking refer to the instructions in the Engine Monitor System of this section. CAUTION These indicators should not be relied upon for accurate fuel quantity readings at any time especially during skids, slips, or unusual attitudes. The fuel gauges are approximate indications of fuel level and are never substitutes for proper planning. Always verify fuel onboard through a visual inspection. During flight compute the fuel used by monitoring time and established fuel flow rates. ISSUE 2 7-49

SECTION 7 FOUND FUEL FLOW RATE AND FUEL REMAINING The fuel system is equipped with a fuel flow sensor. This sensor is factory-set to measure the fuel rate in gallons per hour (GPH). The sensor is connected to the MVP- 50 engine monitor and the rate is displayed on the monitor. The MVP-50 engine monitor will also calculate and display the fuel remaining using the data from the fuel flow sensor and the fuel added quantities input by the pilot. Refer to the MVP-50 operator s manual for instructions for inputting the Fuel Added and updating the Fuel Remaining. CAUTION Do not rely solely on the Fuel Remaining displayed on the MVP-50 to determine the fuel levels in the airplane. The use of the MVP-50 does not eliminate or reduce the necessity for the pilot to use good flight planning, pre-flight and inflight techniques for managing fuel. 7-50 ISSUE 2

FOUND SECTION 7 FUEL TANK SELECTOR VALVE The fuel selector/shutoff valve is located on the center console between the crew seats. The selector allows the engine to draw fuel from the LEFT or RIGHT tanks but not both at the same time. The selector also has an OFF position that is reached by rotating the selector all the way to the right while pulling up the safety catch on the selector. This safety catch prevents the fuel from being shutoff by mistake. It is important that the pilot monitors fuel consumption and manages the fuel appropriately. The fullest tank should be selected prior to takeoffs or landings. Figure 7-19 Fuel Tank Selector ISSUE 2 7-51

SECTION 7 FOUND FUEL DIPSTICK A wooden fuel dipstick is provided with each airplane and is located in the pocket behind the pilot s seat. One side is labelled WHEELS and is graduated in quarter-tank increments (FULL, ¾, ½, and ¼). The opposite side is labelled FLOATS and is graduated in 5- gallon increments (FULL and 35 through 5). Fuel is measured by opening the fuel cap on either wing and dipping the stick into the tank until it touches the bottom. Remove the stick and read the amount of fuel remaining in each tank. The reading appears as a wet discolouring of the wood. Keep the dipstick clean and free of contamination. WARNING REFRAIN FROM SMOKING AROUND THE AIRCRAFT ESPECIALLY WHEN MEASURING FUEL TANK LEVELS OR WHEN HANDLING A FRESHLY-WET DIPSTICK. AUXILIARY FUEL PUMP SWITCH The electric, auxiliary fuel pump switch is located on the Overhead Switch Panel. The switch is labeled AUX. FUEL and OFF. The auxiliary fuel pump should be activated during takeoffs, landings, and low level flight as a backup in the event that the engine-driven mechanical fuel pump fails. The auxiliary fuel pump is rated for intermittent use only and should be switched off after takeoff. In the event of a failure of the engine driven pump the auxiliary fuel pump must be run continuously to keep the engine running. 7-52 ISSUE 2

FOUND SECTION 7 ELECTRICAL SYSTEM AND INSTRUMENTS The airplane is equipped with a 28V, direct current electrical system (refer to Figure 7-20, Electrical System Schematic). The system is powered by a belt-driven, 70A alternator and a 24V battery, located within the engine compartment on the starboard side. Power is supplied to most general electrical circuits through a single primary bus bar. The primary bus bar also provides power for ignition and alternator control circuits. The avionics bus bar is connected to the main bus bar via a single avionics relay. The primary bus bar is activated by the master switch. The avionics bus bar is activated when both the master switch and the avionics master switch are ON. A power distribution module is located within the engine compartment on the starboard side of the firewall. It houses the master and start relays, the amp meter shunt, the two 80 amp fuse links (one for the alternator, one for the battery). In addition, there are four standard circuit breakers (two 1 amp fuses protect the amp meter leads, one 5 amp fuse related to the regulator and one 5 amp fuse connected directly to the battery to supply back up power in the event the main fuse links are blown. The voltage regulator is located on the chassis mounted to the circuit breaker panel. The avionics master switch is located on the Overhead Switch Panel. ISSUE 2 7-53

SECTION 7 FOUND ELECTRICAL SYSTEM SCHEMATICS Figure 7-20 Electrical System Schematic 7-54 ISSUE 2

FOUND SECTION 7 CIRCUIT BREAKERS All circuit breakers are of the push-to-reset type. If a circuit trips, it may be reset by pushing on the head of the circuit breaker until it snaps into place. If the circuit trips again, do not reset. Have the system checked by licensed aircraft maintenance personnel. The circuit breakers are located on the Circuit Breaker Panel. ISSUE 2 7-55

SECTION 7 FOUND OVER- AND UNDER-VOLTAGE ANNUNCIATOR LIGHTS The over-voltage annunciator illuminates anytime the voltage exceeds 32 volts, while the under-voltage annunciator illuminates anytime the voltage drops below 25 volts. The lights will extinguish automatically once the voltage returns to within normal limits. The voltage annunciator lights cannot be cancelled manually. When an over-voltage condition occurs, the alternator control unit may also trip the alternator circuit breaker, thereby removing alternator field current and shutting down the alternator. If this occurs, the battery will then supply all the power to the electrical system at a discharge rate shown by the ammeter. The under voltage warning annunciator will illuminate when system voltage drops below normal. The alternator control unit may be reactivated by resetting the circuit breaker. If the warning light extinguishes, normal alternator charging has resumed. If the light illuminates, a malfunction has occurred, and the flight should be terminated as soon as practical. NOTE Illumination of the under-voltage light plus ammeter discharge indications may occur during low rpm conditions with an electrical load on the system during a low rpm taxi. Under these conditions, the light will go out at higher rpm. The over and under voltage annunciator lights are located on the Circuit Breaker Panel. 7-56 ISSUE 2

FOUND SECTION 7 MASTER SWITCH The master switch, located on the Ignition Switch Panel, is an environmentally-sealed toggle switch. The master switch is ON in the up position and OFF in the down position. It controls the electrical power to the airplane by connecting the battery to the Electrical System. CAUTION PRIOR TO TURNING THE MASTER SWITCH ON OR OFF, STARTING THE ENGINE OR APPLYING AN EXTERNAL POWER SOURCE, THE AVIONICS POWER SWITCH, LABELED AVIONICS MASTER, SHOULD BE TURNED OFF TO PREVENT ANY HARMFUL TRANSIENT VOLTAGE FROM DAMAGING THE AVIONICS EQUIPMENT. The master switch and alternator switch should be activated or de-activated at the same time. The master switch can be activated separately on the ground to check electrical equipment (using the battery). To check or use avionics equipment or radios on the ground, the avionics power switch must be activated. ALTERNATOR ENABLE SWITCH The alternator enable switch (labelled ALTERNATOR), located on the Ignition Switch Panel, is an environmentally-sealed toggle switch. This switch is ON in the up position and OFF in the down position. It enables the belt-driven alternator to supply power for battery charging and airplane systems. The master switch and alternator switch should be activated at all times during flight. In the event of an alternator failure, all necessary electrical equipment can be operated using only the master switch (see Section 3 Emergency Procedures). A malfunctioning alternator or the alternator switch set to OFF, will place the entire electrical load on the remaining battery charge. ISSUE 2 7-57

SECTION 7 FOUND AVIONICS MASTER SWITCH Electrical power for the avionics bus bar is supplied through the main bus bar. This switch is located on the Overhead Switch Panel and isolates the avionics bus bar and all the attached equipment from the main bus bar. It is important that the avionics master switch, located on the Switch Panel, be off at engine start or stop to protect the avionics equipment from power surges. BACKUP POWER SWITCH In the event of a primary electrical failure on the main bus bar, emergency power can be supplied to the MVP-50 engine monitor by bypassing the master solenoid and connecting directly to the aircraft battery. The backup power circuit also provides power to the overhead Crew Light The back up power switch is located on the Overhead Switch Panel. 12V ACCESSORY OUTLET The aircraft is equipped with a 12-volt (13.75V actual) accessory outlet located on the Right Hand Panel. When the master switch is set to ON, this outlet can be used to provide power to devices requiring 12 volts DC on the ground or in flight. Remember that devices connected to this socket for long periods of time with the engine off, will drain the battery. AMMETER/VOLTMETER GAUGE The airplane is not equipped with a separate ammeter/voltmeter gauge, instead the current and voltage is measured by the MVP-50 engine monitor and the values are displayed in digital format in the lower right corner of the monitor. When the engine is operating and the master switch is set to ON, the ammeter indicates the charging rate applied to the battery. In the event of the alternator not functioning or if the electrical load exceeds the output of the alternator, the ammeter indicates the battery discharge rate. 7-58 ISSUE 2

FOUND SECTION 7 TURN CO-ORDINATOR (INSTALLED WITH OPTIONAL AUTOPILOT) The turn co-ordinator, located on the Flight Instruments Panel, is an electrically-operated gyro instrument that indicates both the rate of roll and the rate of turn in degrees per second. It will indicate the direction of the turn, but will not show the bank angle. As a cross-check to the attitude indicator, the turn co-ordinator can be used to keep the wings level. It can be used for timed turns of three degrees per second, or two minutes for 360 degrees of turn. The co-ordination ball indicates the quality of the turn since it responds to centripetal forces in a slip and to centrifugal forces in a skid. A slip or a skid will affect the radius of a turn but not the rate of turn. ISSUE 2 7-59

SECTION 7 FOUND PITOT-STATIC SYSTEM AND INSTRUMENTS The Pitot-Static System supplies ram air pressure to the G500 Air Data Computer and the standby airspeed indicator. It also supplies static pressure to the G500 Air Data Computer, the standby airspeed indicator, and standby altimeter. The system is composed of a heated L-type Pitot tube mounted on the starboard wing and two static ports located on the left and right sides of the aft fuselage. PITOT HEAT The heated Pitot tube consists of a heating element in the Pitot tube, a 10A circuit breaker and switch (labelled: PITOT HEAT). The switch is located on the Overhead Switch Panel; the circuit breaker is located on the Circuit Breaker Panel. The Pitot heat is on in the up position. With the Pitot heat switch on, the element in the Pitot tube is heated electrically to maintain proper operation in possible icing conditions. WARNING Even though the airplane is equipped with a heated Pitot tube, deliberate flight into icing conditions is prohibited and dangerous. The Pitot-Static System s instruments are depicted and described as follows. They are located on the Flight Instrument Panel. Refer to Figure 7-5 for location. STANBY AIRSPEED INDICATOR The standby airspeed indicator displays the airplane s indicated airspeed. White, green, yellow, and red limitation and range markings are defined in Section 2, Limitations, Figure 2-2. 7-60 ISSUE 2

FOUND SECTION 7 STANDBY ALTIMETER An aneroid-type, three-pointer standby altimeter indicates the airplane s altitude. To properly indicate the altitude above sea level, the altimeter must be set to the current barometric pressure using the adjustment knob near the lower left corner of the meter. This knob enables the adjustment of the instrument s barometric scale (in. Hg) to the current altimeter setting within the Kollsman window. The airplane s altitude is indicated by three needles. The largest needle indicates hundreds of feet. The middle size needle indicates thousands of feet, and the third needle indicates ten thousands of feet. Clockwise motion indicates climb. Counter-clockwise motion indicates descent. The barometric scale appears in the Kollsman window on the right side of the instrument s face. The black and white hatched area indicates whether the aircraft is above or below 16,000 feet. As the airplane climbs to 16,000 feet, the hatched area gradually disappears until, at 16,000 feet, it is completely invisible. ALTERNATE STATIC SOURCE SWITCH If the aircraft is IFR ready, an alternate static source switch is available. This switch is located on the Pitch Trim Panel above the trim indicator. The alternate static source can be activated by lifting the handle to the UP position. This opens the static line to the aircraft cabin and supplies the cabin static pressure to the static source instruments (airspeed, altitude, vertical speed and transponder, etc.). If erroneous instrument readings are suspected, this alternate static source should be turned on and airspeed calibration and altitude correction charts in Section 5 should be utilized. ISSUE 2 7-61

SECTION 7 FOUND VACUUM SYSTEM AND INSTRUMENTS The Vacuum System provides the suction necessary to operate the standby attitude indicator and the heading indicator. It consists of an engine-driven vacuum pump, a pressure regulator, a vacuum relief-valve, a suction gauge, an inlet air filter, and the vacuum-operated instruments. Suction Gauge is located on the MVP-50 Engine Monitor Display Figure 7-21 Vacuum System 7-62 ISSUE 2

FOUND SECTION 7 The Vacuum System s instruments are depicted and described below. They are located on the Flight Instrument Panel. See Figure 7-5 for location. VACUUM (SUCTION) GAUGE The airplane is not equipped with a separate vacuum gauge, instead the vacuum is measured by the MVP-50 engine monitor and the values are displayed in digital format in the lower right corner of the monitor. The vacuum is calibrated in inches of mercury and indicates the vacuum available for the operation of the attitude indicator and directional gyro. The desired vacuum range is 4.5 to 5.2 inches of mercury. Normally, a suction reading out of this range may indicate a system malfunction or improper adjustment. In such a case, the attitude indicator and directional gyro should not be considered reliable. STANDBY ATTITUDE INDICATOR The standby attitude indicator provides an artificial horizon for the pilot. Bank angle is presented by a pointer at the top of the indicator relative to the bank scale indexed at 10, 20, 30, 60, and 90 either side of the center mark. Pitch and roll attitudes are presented by an airplane superimposed over a symbolic horizon area divided into two sections by a white horizon bar. The upper blue area and the lower brown or black area have arbitrary pitch reference lines useful for pitch attitude control. A knob at the bottom of the instrument is provided for in-flight adjustment of the indicator airplane to the horizon bar for a more accurate pitch attitude indication. ISSUE 2 7-63

SAFE FLIGHT DO NOT PAINT VANE OR SLOT 0 AFT SECTION 7 FOUND STALL WARNING SYSTEM STALL WARNING HORN AND LIGHT Figure 7-22 Stall Warning Sensor The airplane is equipped with an electrically-powered Stall Warning System consisting of a vane-operated switch on the leading edge of the starboard wing, electric horn, and red indicator light. The stall warning indicator is located on the Flight Instrument Panel. As the airplane approaches a stall, the vane switch closes. The closed switch completes the stall warning circuit resulting in a warning horn and illuminated red light at 5 to 10 knots above the stall. The warning indicators will extinguish when the airplane has recovered from a stall. 7-64 ISSUE 2

FOUND SECTION 7 LIGHTING SYSTEMS EXTERIOR LIGHTING The aircraft has two types of exterior lighting: 1. A combo navigation/strobe light on each wing tip. 2. A landing light on the cowl. The exterior lights are operated by toggle switches located on the Overhead Switch Panel. The light switches are labelled NAV, STRB, LAND LTS. LED Landing/Taxi Light WARNING THE STROBE LIGHTS SHOULD NOT BE USED WHEN FLYING THROUGH CLOUDS OR OVERCAST. THE FLASHING LIGHT, REFLECTING FROM WATER DROPLETS OR PARTICLES IN THE ATMOSPHERE, PARTICULARLY AT NIGHT, CAN PRODUCE VERTIGO AND A LOSS OF ORIENTATION. The aircraft may be equipped with a combination LED Landing/Taxi light on the cowl. This light has 15 LEDs (9 landing, 6 taxi) The light is controlled with three switches labelled LAND, TAXI, and PULSE located on the Overhead Switch Panel. The light may be operated in the following modes: Landing Light ON Taxi Light ON Landing and Taxi Lights ON Taxi and Pulse ON Activating the TAXI and PULSE switches will cause the taxi light to pulse (turn off and on). Pulsing the taxi light will increase the visibility of the aircraft to oncoming traffic. ISSUE 2 REV 1 March 14, 2013 7-65

SECTION 7 FOUND DIMMER CONTROLS FLT PNL PANELS ON OFF PITOT HEAT BK-UP PWR CLOSE COWL OPEN COWL FLAP CLOSED OPEN FUEL PUMP AVIONICS MASTER NAV LTS STRB LTS ON OFF LAND LTS TAXI LTS PULSE LTS CAB LTS CREW LTS Figure 7-22A Overhead Switch Panel INTERIOR LIGHTING beam. A removable forward cabin (crew) light is provided. The crew light is a multi-function, adjustable light that is mounted in the cockpit ceiling. The light can be removed by pulling down on it and remounted by firmly pushing the light upwards into its holder. Once removed, the light can be used as a flashlight, spotlight, map light or signal light. The light intensity can be adjusted by rotating the circular switch situated on the aft end of the housing, to provide a tighter or wider A switch on the back of the lamp can be activated to beam red light of varying intensity for darker lighting conditions. The crew light switch is located on the Overhead Switch Panel. 7-66 ISSUE 2 REV 1 March 14, 2013

FOUND SECTION 7 INSTRUMENT PANEL LIGHTING Two dimmer control knobs are located on the Overhead Switch Panel. One dimmer controls the Flight Instrument Panel, while the other dimmer controls the other panels. Interior back-lighting is provided with the flight instruments, compass, and fuel selector. Post lights are located at various locations to light the flap position indicator & switch, master & alternator switch, ignition switch, trim position indicator, etc. REMOTE EMERGENCY LOCATOR TRANSMITTER (ELT) The aircraft is equipped with an Artex ME406 remote Emergency Locator Transmitter. The ELT is designed to emit automatically an emergency radio signal to help rescuers locate the aircraft in the event of an accident. The ELT is located in the aft fuselage behind the baggage bulkhead. It can is accessible via an easily removable access panel on the port side of the baggage bulkhead. The ELT antenna is located on top of the aft fuselage. The pilot can activate the ELT with a switch located on the center console next to the trim wheel. The red ELT ON light indicates that the ELT is operational and broadcasting an emergency signal. The pilot can deactivate the ELT by placing the switch in the ARM position, (should it be activated accidentally). The ELT CANNOT be disarmed or disabled from the cockpit. Cockpit operation is limited to deactivating or manually activating the ELT. The ELT transmit a 406 MHz emergency signal to the Cospas/Sarsat satellites and a local 121.5 homing signal. The ELT automatically activates during a crash and transmits a continuous swept tone of 121.5 MHz. During activation, the 406 MHz ISSUE 2 REV 1 March 14, 2013 7-67

SECTION 7 FOUND transmitter sends an encoded 5-watt signal to the Cospas-Sarsat Satellites every 50 seconds for 440 milliseconds to alert Search and Rescue. MAGNETIC COMPASS AND DEVIATION CARD The compass and deviation card are located directly above the Flight Instrument Panel. The compass is a standard, liquid-filled, magnetic compass that points to magnetic north. The deviation card can be found beside the compass. OUTSIDE AIR TEMPERATURE (OAT) GAUGE The airplane is not equipped with a separate outside air temperature (OAT) gauge. A remote OAT probe on the port wing inboard flap track rib is connected to the G500 system. The OAT is displayed in degrees Fahrenheit ( o F) on G500, although it may be changed to degrees Celsius ( o C) by the pilot. 7-68 ISSUE 2 REV 1 March 14, 2013