Twin Otter GenFam.pdf

DeHavilland Twin Otter ENGINEERING TRAINING DEPARTMENT Gen Fam

For traIning purposes only - Issue 2 Rev 1 Page 1

DeHavilland TWIN OTTER

DHC-6 SERIES 310

GENERAL FAMILIARIZATION

WARNING This study guide is intended for training purposes only. The information it contains is as accurate as possible at the time of issue, and it is not subject to ammendment action. Where the information contained in this study guide is at variance with official documents, the latter must be taken as the overiding authority. The contents in this study guide shall not be reproduced in any form without the expressed permission of MAA Engineering Training Department.



Contents Subject Page General 4Dimensions 6Powerplant 10Engine Oil 12Propellers 14Fuel System 16Doors 20Baggage Compartment Doors 26Loading Jury Strut 26Cargo Compartments 28Landing Gears 30Wheel and Brakes 32

Subject Page Hydraulic System 38 Hydraulic Power Package 40 Flight Controls 42 Gust Locks 42 Pitot Static System 44 Electrical Power 46 Exterior Lighting 50 Fire Detecting System 52 Fire Extinguishing System 54 Hand Operated Fire Extinguishers 56 Emergency Equipment 56



Servicing diagram



GENERAL The de Havilland DHC-6, Series 310 Twin Otter aircraft is an all-metal, high wing monoplane, powered by two wing- mounted, turbo-prop engines, each driving a three-bladed, reversible pitch, fully feathering propeller. The aircraft carries a pilot, copilot, and up to 20 passengers (MH 19 Pax), depending upon the seating configuration. Alternatively the Twin Otter can be equipped for cargo transportation, ambulance duties, supply dropping, aerial survey operations and fire-fighting. The aircraft can be adapted for operations on wheel-skis, spring skis, floats (which necessitates the installation of a short nose), and intermediate flotation gear (for "soft field" operation) Optional installations available include wing fuel tanks, anti-icing and deicing systems, crew and passenger oxygen systems, autopilot, propeller synchronizer, air conditioning, and various radio, navigation and communication systems.

NOTES



Layout of the cockpit



AIRCRAFT DIMENSIONS AND GROUND CLEARANCES With the aircraft supported on its landing gear, the principal dimensions and ground clearances are as follows: Note: Dimensions to Ground Line are approximate only and vary depending on aircraft configuration and loading conditions. OVERALL DIMENSIONS: Span 65 ft - 0 in. Length 51 ft - 9 in. Height (tires normal) 19 ft - 6 in. Propeller ground clearance (static) 5 ft - 0 in. Design gross weight 12,500 pounds WINGS: Type High Chord 6 ft - 6 in. Incidence 2 - 1/2o Dihedral 3o Aspect ratio 10.0 Airfoil Section DeHavilland High Lift Section

HORIZONTAL STABILIZER AND ELEVATORS: Span 20 ft - 8 in. Chord 4 ft - 9 in. Incidence 0o Dihedral 0o Aspect ratio 4.35 Airfoil Section 63 A213 Inverted and Modified VERTICAL STABILIZER: Height 13 ft - 6 in. Root Chord 10 ft - 0 in. (Basic) Tip Chord 5 ft - 0.6 in. Aspect ratio 1.84 FUSELAGE: Width (maximum) 5 ft - 9.2 in. Height (top of center fuselage) 9 ft - 8 in. AREAS: Wing 420 sq ft Horizontal stabilizer and elevators 98 sq ft LANDING GEAR: Main Wheels track 12 ft - 2 in.



Dimensions of the Twin Otter



JACKING THE AIRCRAFT Five jacking points are provided on the aircraft and are located on the underside of the fuselage and main gear. Each jacking pad is coded with a letter, the related insert shows the jacking application. Tail lifting or slinging is accomplished by a bar inserted in the rear fuselage lifting tube and supported by two jacks, one at each end. PARKING PROCEDURES Park the aircraft by heading into the wind, applying the parking brakes, locking the control surfaces, fitting protective covers and propeller restrainers as necessary, and closing all doors and access provisions. Before leaving the flight compartment, insure that controls are set as follows: Propeller levers to full INCREASE Power levers to IDLE. Fuel levers to OFF BATTERY/EXTERNAL switch to OFF PARKING BRAKE A push-pull type parking brake handle is located on a pedestal between the pilot's rudder pedals. To apply the parking brake, depress both brake pedals fully, pull out the parking brake handle and release pedals. To release parking brakes, apply pressure to both brake pedals, push brake handle fully in and release pedals. MOORING PROCEDURES Mooring is accomplished by parking the aircraft, chocking all wheels fore and aft, and tying aircraft down with mooring ropes using aircraft tie-down points.

TOWING Towing the aircraft is accomplished by attaching a towbar to the nosewheel axle and disconnecting the nose leg torque links. To connect the towbar to the nosewheel axle, insert pawl pins through ends of the towbar into the axle spools, then insert L-shaped locking pins vertically through ends of the towbar and the groove in pawl pins. Station a crew member in flight compartment to operate brakes when aircraft is being towed. CAUTION: INSURE THAT BRAKE ACCUMULATOR IS FULLY CHARGED TO 1500 PSI BEFORE COMMENCING TO TOW AIRCRAFT.



Aircraft Jacking Points



POWER PLANTS The power plants comprise two United Aircraft of Canada Limited PT6A-27 engines, each housed in a wing nacelle. Each engine has two independent turbines, one driving the engine compressor, fuel pump and accessory gearbox, and the other driving the propeller through a reduction gearing. A starter-generator and a gas generator tachometer are mounted on the accessory gearbox. Retractable intake deflectors for ice and snow protection are installed in the engine air inlet ducts. On some aircraft (applicable to MH) a spray ring for engine compressor washing is installed inside each engine nacelle. ENGINE WASH SYSTEM. An engine wash spray ring installed in each engine nacelle provides for engine compressor washing when operating in atmospheric conditions of high salt content or industrial pollution. The installation in each nacelle consists of a pipe assembly (spray ring) attached to the nacelle structure and partially encircling the engine; perforations in the pipe face inward to direct cleansing agent into the engine. A capped connector in the pipe assembly, which protrudes through the inboard side of the nacelle, provides for the connection of a hose and pumping apparatus when a washing operation is to be carried out.

ENGINE / PROPELLER CONTROLS. The engine / propeller controls are mounted in the overhead console in the flight compartment, and comprise power levers, propeller levers, and engine fuel levers. Friction control knobs for the power and propeller levers are also installed. IGNITION SYSTEM. The ignition system for each engine comprises a regulator, two ignitors, an ignition selector switch, two shielded cables, an ignition switch, and an ignition relay. Power is supplied through push-to- reset circuit breakers on the main circuit breaker panel marked IGN L or IGN R. STARTING SYSTEM. The starting system for each engine consists of a starter-generator, a start switch (common to both engines), and two starter relays. When the dc master switch is on and the start switch selected for the required engine, power is supplied to the starter-generator, which rotates the gas generator turbine at sufficient speed to provide engine light-up. When light-up occurs and the engine has accelerated to idle speed and the switch is released, the starter relays deenergize the starter circuit, and when the generator is switched on the starter-generator produces dc power for the aircraft electrical system. Power for the starting systems is supplied through two circuit breakers on the main circuit breaker panel, labeled START L and START R.



The engine with cowls opened



ENGINE OIL SYSTEM Each engine oil system is integral with the engine and incorporates an oil pump, filter, and tank. An oil cooler is installed on the underside of the engine and is enveloped by the engine lower cowling which directs the airflow through the cooler. Oil flow through the core of the cooler is controlled by a regulator which diverts the oil through a bypass until normal operating temperature is reached; it also ensures oil circulation in the event of restricted or blocked flow through the cooler. Heated oil is used as a medium for raising the temperature of the fuel prior to its delivery to the engine. This is accomplished by piping oil through a fuel heater mounted on the engine upper casing. Each engine oil tank has a capacity of 2.3 U.S. gallons of which 1.5 U.S. gallons is usable. The oil tank filler neck is accessible at the engine accessory gear case. For oil pressure indication, a pressure transmitter is installed in the engine accessory gear case and is connected to the oil pressure indicator; a pressure switch is installed also, in a line tapped into the accessory gear case, and which is connected to the low oil pressure caution light. For oil temperature indication a resistance bulb is installed in the accessory gear case and connected to the oil temperature indicator.

CHECKING OIL QUANTITY CAUTION: WAIT AT LEAST 5 MINUTES AFTER ENGINE SHUT DOWN BEFORE OPENING THE OIL FILLER CAP. A quantity dipstick is contained in the filler cap and indicates the quantity of oil required to top up tank to the full mark. To check oil quantity, open panel to gain access to oil filler cap, remove cap and check oil level in normal manner. Install and secure cap on completion of check. Oil level should be checked within 20 minutes of engine shut down. Normal oil consumption is 0.1 lb/hr. More than 0.2 lb/hr is not acceptable. NOTE: Normal oil level is at the one quart mark on the dipstick. Overfilling may cause discharge through breather. If oil change has been performed or engine has remained sationary for a period of more than 12 hours, start engine, run at idle for 2 minutes, feather propeller, shut down engine, check oil and replenish if required with same brand of oil. CAUTION: DO NOT MIX DIFFERENT BRANDS OF OIL WHEN CHANGING OR ADDING OIL BETWEEN OIL CHANGES, SINCE THE CHEMICAL STRUCTURE OF DIFFERENT BRANDS OF SYNTHETIC OIL MAY DIFFER SUFFICIENTLY TO MAKE THEM INCOMPATIBLE.



OIl dipstick attached to the oil cap



PROPELLERS The two propellers are Hartzell HC-B3TN-3D, metal, counterweight, three-bladed, fully feathering, reversible pitch, speed governed units with pitch change mechanisms and hydraulic systems. Each propeller is 8 ft 6 in. in diameter and has a pitch range of -15o (reverse) to +87o (feather), and a pitch setting of +17o. Each is controlled in the constant speed and feather ranges by the propeller lever through a propeller governor on the propeller reduction gearbox. The power lever is connected to the propeller reverse cam mechanism for control of the propeller in its 17o to -15o (beta) range. Each propeller system incorporates a propeller overspeed governor, an automatic feathering system and an electrically-operated beta back-up system; the latter to prevent either propeller form going below IDLE rpm blade angle unless selected. Propeller blade latches (when installed) provide for engine starting and stopping with propeller blades at the zero thrust position. On special order a propeller synchronizing system can be installed. The speed of each propeller is indicated on a propeller tachometer on the engine instrument panel.

NOTES



Keep clear of the propeller arcs when working on the aircraft.



FUEL SYSTEM Fuel is contained in two fuselage fuel tanks, one forward and one aft. The tanks can be gravity refueled through two filler points, one for each tank, located on the left side of the fuselage near the wing strut. The forward tank supplies the right engine and the aft tank the left engine, with the main feed lines to both engines passing through No. 6 cell (aft tank). A crossfeed valve provides a fuel supply to both engines from one tank. Each fuel tank consists of four interconnected fuel cells. Two fuel booster pumps, installed in the collector cell of each tank, deliver fuel under pressure to the engine; this pressure also operates a fuel ejector which transfers fuel from the end cell into the collector cell. From No. 6 cell the main feed line to each engine is routed up the rear side of the wing strut, along the wing and into the appropriate nacelle. Mounted in each nacelle behind the firewall is a fuel strainer, a flow meter transmitter and a fuel emergency shutoff valve. To permit checks for water accumulation in the fuel system, water drain valves are provided in the forward and aft fuselage tank interconnector manifolds and in each fuel strainer. In addition, a fuselage tank fuel drain valve is provided in each sump plate for No. 4 and No. 5 collector cells. A fuel quantity indicating system of the capacitance type is provided for each fuel tank. Fuel low pressure and low level indicating lights are provided on a caution lights panel in the flight compartment.

FUEL TANK CAPACITIES. Usable capacities of the two fuel tanks are as follows:- Weight - lb Imp Gal U.S. Gal Forward tank 1,178 151 182 Aft tank 1,280 164 198

JP4 fuel = 7.8 lb per Imp. gallon. Fuel system arrangement

REFUELING Filler neck refueling (1) Observe refueling safety precautions (2) Insure external power supply to aircraft is disconnected and aircraft supply is off (3) Ground supply nozzle to ground socket, remove filler cap, insert nozzle and fill tank. When tank is full, remove nozzle, replace and secure cap and disconnect ground plug from ground socket.



Fuel system arrangement



FUEL QUANTITY INDICATORS. The two fuel quantity indicators are located on the pilot's flight instrument panel and identified AFT and FWD. Both indicators denote fuel contents of their respective tanks in pounds and each indicator dial is marked FUEL QUANTITY LBS x 100. The aft indicator has scale markings ranging from 0 to 1425 in increments of 25 lb with even numerals from 0 to 14 at 200 lb intervals. The FWD indicator has scale markings ranging from 0 to 1400 in increments of 25 lb with even numerals at 200 lb intervals from 0 to 12. Each indicator reflects the mean value of the fuel levels detected by four capacitance type probes in its associated fuel tank. The fuel quantity indicators are powered from the 400 cycle, 115-volt ac bus through FUEL QTY FWD and AFT fuses on the fuse panel. On some airplanes the fuel quantity indicators are located on the engine instrument panel. DRAIN VALVES. There are four water drain valves incorporated in the fuel system to provide a check for water accumulation. One is installed on each fuel strainer and one in each fuel manifold. Those on the fuel strainers are located in the engine nacelles, the remaining two are located in the trough on the underside of the fuselage. Aircraft with Mod S.O.O 6048 incorporated have an additional water drain valve (Mod 6/1498) fitted in the aft fuel manifold below No. 8 cell. Aircraft with Mod S.O.O. 6118 incorporated have two additional water drain valves, one fitted in each of the fuel manifolds below No. 1 and 8 cells, respectively (applicable to MH). A fuselage tank fuel drain valve is fitted to the sump plate of each collector cell and is located behind an access panel near the fuel manifold water drain.

FUEL SYSTEM WATER DRAINING Water and/or contaminants must be drained from fuel strainers and fuselage tank interconnector manifolds at regular intervals as prescribed in the Inspection Schedule. To accomplish this, open water drain valves provided, keep valves open until clean fuel only is ejected and close valves. NOTE: Due to the higher locations of fuel strainers in engine nacelle, water and/ or contaminants from strainers may be drawn into fuel tanks when associated drain valves are opened. To prevent contamination of fuel tanks, booster pumps must be operating before fuel strainer drain valves are opened.



Fuel control and indication Fuel strainer drain



DOORS The aircraft entrances and exists consists of two flight compartment doors, two passenger cabin doors, front and rear baggage compartment doors, two emergency escape doors, one on each side of the aircraft, and two servicing access doors. On aircraft with Mod 6/1325 embodied, the front baggage compartment front latch incorporates a lock which can be operated externally with a key. Two sliding doors, located at the cabin front bulkhead, provide interconnection between the cabin and flight compartment. A doors unlocked caution system is installed on aircraft incorporating Mods S.O.O 6061, 6/1239, 6/1268 or 6/1284. At the customer's option, a zippered quilted vinyl/fabric doorway, or a hinged door incorporating a vent and a latch may be installed to permit access to the rear baggage compartment. On aircraft incorporating Mod 6/1613, two hinged doors (upper and lower) are installed on the cabin rear bulkhead. Both doors open to the rear to permit access to the rear baggage compartment.

PASSENGER/CREW DOORS. Two crew entry doors, one each side of the flight compartment, and two passenger doors, one each side of the cabin compartment are provided. FLIGHT COMPARTMENT DOORS. Each flight compartment door has a fixed window and a sliding window with a release handle. The doors are secured to the airframe structure by two hinges at the forward edge of each door, which allows either door to be opened outward and forward, to a position limited by a nylon restraining strap. An interior and exterior door handle is installed, which when turned clockwise, operates a latch mechanism which disengages a mid-door bolt, and an upper latch pin. Each exterior door handle can be locked by key from the outside. A map case is provided in the lower part of each door.



Doors - location



CABIN COMPARTMENT DOORS. The cabin left compartment door is in two parts, the forward part is normally an airstair door for passenger entry, and the rear part to provide a larger entrance space for cargo loading. At the customer's option, the airstair door can be replaced by a forward cabin/cargo door (Mod S.O.O 6107), which when installed is hinged to the forward part of the door frame, to permit the door to be opened outward and forward to a limit determined by a stay assembly. An entrance ladder, which in use engages two slots in the door still, can be stowed on the door. This door has an inspection window displaying door locked indicator marks similar to those described later for the right-hand cabin door.

NOTES



Left Rear Cabin Door



The airstair door as installed to the basic aircraft, opens downward and outward, and is supported in the open position by two covered cable assemblies. One cable assembly is connected to the front edge of the door, and the other to the rear, both cables run parallel with each other and attach to the fuselage structure at the upper door frame. The upper section of each cable assembly (Pre Mod 6/1527) incorporates a strap to assist in closing the door. Aircraft incorporating Mod /1527 have an adjustable eye-fitting with a locknut on the rear cable assembly. The fitting is threaded onto the upper end of the rear cable assembly and connected to the rear door-frame. The fitting can be adjusted to equalize cable tension between the front and rear cable assemblies. On aircraft with Mod S.O.O 6114 or Mod 6/1606, the airstair door is provided with a door restraint (door saver) which prevents the door falling free when being lowered and gives spring assist when being raised. The airstair door can be quickly removed to provide easier access for cargo loading. To assist in passenger entry and exit, the airstair door is provided with handrails and an extension step. Each handrail is attached between its associated covered cable assembly and the appropriate side of the airstair door, and is so designed that as the door is closed it will fold up against the inner side of the door. The airstair door has an internal and external handle which operates a latch mechanism to engage the two door locking pins in the fuselage upper door surround structure. An inspection window provides the means of checking that the two fluorescent orange witness marks (one on the locking pin rod and the other on a fixed bracket), align

when the door is fully locked. Contacts, to complete the electrical circuit to the internal door unlocked microswitch, are located on both the door and the fuselage door upper surround. The cabin left rear door is hinged to the rear part of the door frame, and opens outward and rearward. When closed, the door is locked top and bottom by engaging sliding bolts which are then secured with elastic straps. The cabin right door, which opens outward and forward, is hinged at the forward part of the door frame. A strap assembly restricts the door movement to prevent it from contacting and causing damage to the fuselage skin. The door latch mechanism, operated by an internal or external handle, has a latch bolt which engages in the door rear surround structure. An inspection window adjacent to the door lock handle shows, when the door is locked, a part of the door latch mechanism marked with orange dayglow paint. With the door locked this painted portion should be aligned with painted lines on each side of the inspection window. The exterior cabin door handles can be locked by a key from the outside, and each interior handle is covered by a spring-loaded plastic safety guard which must be lifted in order to operate the handle.



Airstair Door (open) and Right Cabin Door



EMERGENCY EXIT. The emergency exit consist of two escape doors, one on each side of the passenger cabin compartment. ESCAPE DOOR. Each escape door has a handle release mechanism, a window, door seal, and rubber pads bonded to the side and bottom structure. Pulling downward on the release handle disengages the release handle from fittings on the cabin structure and on the door. The top of the door is then pushed outwards until clear of the door frame. BAGGAGE COMPARTMENT DOORS REAR BAGGAGE COMPARTMENT EXTERIOR DOOR. The rear baggage compartment exterior door on the left side of the fuselage is hinged at its top edge and opens upward and outward, and can be supported in the open position by a door stay. The door handle has an integral key lock. the rear baggage compartment door is included in the DOORS UNLOCKED caution light circuit. REAR BAGGAGE COMPARTMENT INTERIOR DOOR. A fabric covered door assembly is installed on the cabin rear bulkhead for access to the rear baggage compartment on utility aircraft. On commuter aircraft a removable panel located in the rear bulkhead directly above the center rear seat can be removed to provide access to the rear baggage compartment. .

FORWARD BAGGAGE COMPARTMENT DOOR. The forward baggage compartment door on the left side of the nose is hinged at its top edge and opens upward and outward and can be supported in the open position by a door stay. The door is secured in the closed position by two recessed latch assemblies. The forward latch can be key locked. This baggage compartment door is included in the DOORS unlocked caution light circuit, when installed. LOADING JURY STRUT. A loading jury strut is supplied with the aircraft for use when loading heavy cargo items. The strut is attached to the aircraft by engagement of its upper end with an adapter on the underside of the fuselage below the rear baggage compartment door. It will normally be clear of the ground but during loading it may support the aircraft by contact with the ground. After completion of loading it should again be clear of the ground.



Emergency exits



CARGO COMPARTMENTS Provision is made for carrying cargo in the front, rear and cabin compartments. FRONT BAGGAGE COMPARTMENT. Four permanently attached tie-down rings are provided in the front baggage compartment. REAR BAGGAGE COMPARTMENT. Eight permanently attached tie-down rings are provided, four in the baggage compartment and four in the baggage compartment extension. A baggage restraint webbing is fitted at the semi-bulkhead between the baggage compartment and the extension. CABIN COMPARTMENT. With any passenger seat removed, cargo tie-down rings can be attached to anchor nuts on the floor rail and cabin wall side rail. On aircraft with rail and track assembly, tie-down rings can be fitted at any location on the track, using Aeroquip cargo tie-down rings.

NOTES



Cargo tie-down point locations



LANDING GEARS MAIN GEAR. Each main gear consists of a leg assembly, shock absorber, and wheel and brake assembly. The leg assembly is bolted to two lower attachment fittings on the fuselage, and these provide the pivoted point for leg movement. Restriction and damping of leg movement is controlled by a shock absorber of the urethane block type. Each shock absorber is bolted to two fuselage upper attachment fittings and to metal pads on the leg assembly. Installed on the axle is a wheel and a hydraulically-operated brake assembly. A two piece fairing encloses the leg. MAIN WHEELS. A magnesium main wheel is carried on the axle of each main gear leg, and is of the split hub type to facilitate removal and installation of tubeless tires. The main wheel has an 8-ply rating, 11.00 x 12 nylon tubeless tire. When modification S.O.O 6048 is incorporated, each main wheel is fitted with a 10-ply rating 15.00 x 12 tire, which may be tubeless or, when Mod 6/1408 is incorporated, fitted with a tube. On aircraft with S.O.O 6048 and Mod 6/1526, a 36 x 13.00 x 12 tire is fitted to each main wheel. When S.O.O 6124 is incorporated, an aluminum wheel (P/N 5000444) is fitted in place of the magnesium wheel.

NOSE GEAR SHOCK STRUT. The nose gear shock strut consists of an outer cylinder, piston assembly, nosewheel fork and a steering collar. The cylinder forms the main structure of the strut and provides the mounting for the nosewheel steering actuator. The nosewheel fork is secured to the piston and telescopes within the cylinder which is charged with MIL-H-5606 oil and air, damping being achieved by restricting the oil and air flow through orifices in the piston. A valve, through which the strut may be recharged, is located on top of the strut. Servicing Nose Gear Shock Strut The nose gear shock strut must be serviced in accordance with instructions on the nose gear identification plate. CAUTION: Dry compressed nitrogen is preferred to compressed air for charging the shock strut. NOSE WHEEL. The nosewheel is carried on the axle mounted in the fork of the nose gear, and is of the split hub type to facilitate removal and installation of tubeless tires. An O-ring is installed between the two halves of the hub to provide an airtight housing for the tire. The nosewheel has 6-ply rating 8.90 x 12.50. Type III, low pressure tubeless tire. When modification S.O.O 6048 is embodied, a main wheel hub fitted with a 15.00 x 12 tire and tube is installed in the nosewheel position. On aircraft with S.O.O 6048 and Mod 6/1526, a 36 x 13.00 x 12 tire is fitted to the nosewheel.



Main gear and nose gear



WHEEL AND BRAKES WHEELS. The main landing gear wheels are carried on the axle of each main gear leg, and the nosewheel is carried on the axle mounted in the fork of the nose gear. The wheels are of the split hub type to facilitate removal and installation of tubeless tires. WHEEL BRAKES SYSTEM. The wheel brakes system consists of hydraulically operated wheel brake units controlled from the pilot's and copilot's rudder pedals. The fluid pressure is metered to the wheel brake units in proportion to brake pedal depression. When the hydraulic system electric motor-driven pump or handpump is operated, fluid is supplied from the main hydraulic system to the control valves and to charge a wheel brakes accumulator. Return fluid from the control valves is fed back to the hydraulic system common return line. A check valve is installed upstream of the supply line to the accumulator. The accumulator supplies brake pressure when the pumps are not in operation.

BRAKE HYDRAULIC PRESSURE INDICATOR. The brake hydraulic pressure indicator is mounted on the pilot's instrument sub panel. The indicator is direct-reading bourdon tube type with dial marked PRESSURE, PSI, and scales marked from 0 to 2000 psi in increments of 100 psi.



Main wheel instalation

Nose gear Identification Plate



WHEEL AND BRAKE UNIT . The wheel is cast magnesium and conforms to all Rim Association standards for a 11.00 - 12 divided type wheel, suitable for use with all 11.00 - 12 8 ply tube type or tubeless tires. It is made up of two wheel halves with an external brake disc fastened together with high strength bolts, washers and self-locking nuts. Rubber lip seals on the inner and outer wheel halves protect the bearings. The wheel brake unit is bolted to the torque plate of each main wheel axle and is a single caliper, 4 piston external disc design, with sintered metallic lining. It is suitable for use with MIL-H-5606 brake fluid. Hydaulic fluid acting on the pistons forces the inner linings against the disc and moves the brake caliper until both the inner and outer linings contact the disc to effect braking. A bleeder screw assembly is fitted at the top rear end of each brake unit.

CHECK BRAKE LINING WEAR. Visually check lining for excessive wear or edge chipping. Linings should be replaced when worn to a thickness of .100 in. The brakes do not need to be applied during this check. Whenever the wear limits are exceeded, replace the applicable worn parts. CAUTION: DO NOT MIX NEW AND USED BRAKE LININGS. Brake conditioning may be accomplished as follows: 1. Perform two (2) full stop braking applications from 30 to 35 knots, allowing the brake discs to cool between each stop. 2. This conditioning procedure will wear off high spots and generate sufficient heat to glaze the lining. Once the lining is glazed, the braking system will provide many hours of maintenance free service. 3. Avoid light use, such as taxiing, which will cause the glaze to be worn rapidly.



Wheel and brake assemblies



WHEEL ASSEMBLY REMOVAL CAUTION: RELEASE BRAKES A. Jack aircraft until tire is clear of ground and deflate tire. B. Remove the 4 back plates from the brake housing by removing 12 bolts. C. Remove snap ring and hubcap from wheel. D. Remove axle nut , spacer, grease seal and outer bearing cone. E. Remove wheel and tire assembly from axle as a unit. WHEEL ASSEMBLY INSTALLATION

A. Mount wheel assembly on axle B. Apply a thin coat of bearing grease on axle nut and

threads. C. Install bearing cone in wheel. D. Install spacer and axle nut on axle. E. Tighten axle nut to 150 to 200 in-lbs of torque while

rotating the wheel to insure proper seating of the bearings.

F. Back off the axle nut to zero torque, then retorque the nut to 40 in-lbs while rotating the wheel. If the locking holes do not align, tighten the nut to the next available key position.

G. Install cotter pin. H. Install grease seal, hubcap and snap ring. I. Install 4 brake lining back plates and torque the 12

bolts to 80-85 in-lbs. and safety wire. J. Remove aircraft from jacks and inflate tire to proper

pressure.

PARKING BRAKE. A push-pull type parking brake handle is located on a pedestal between the pilot's rudder pedals. The handle has restricted movement on the parking brake rod assembly and is spring-loaded to the off position. The rod assembly is connected to a forked lever fitted with two adjustable stops. These stops ride in machined portions of the brake control valves torsion tube levers, and when the brake pedals are fully depressed and the parking brake applied, the forked lever holds the torsion tube levers and brake control valves in brakes on position. To release the brakes, pressure on the brake pedals will allow the spring loading of the parking handle to push the forked lever clear of the torque levers.



Parking brake



HYDRAULIC SYSTEM The hydraulic system operates the wing flaps, nosewheel steering, wheel brakes and skis (when installed). The main components of the system are: an electric motor-driven pump, emergency handpump, reservoir, damping and brake accumulators, flap and nosewheel steering actuators, brake valves, and a flap selector. A filter, pressure switch, relief and thermal relief valves, check valves, and pressure gages are also included in the system. The system is pressurized by the motor-driven pump running on 28 volts DC from a 35 amp CB and pressure supply is retained by the damping and brake accumulators. Pressure is immediately available from the damping accumulator for operation of the flaps, nosewheel steering, and wheel brakes. The brake accumulator supplements the damping accumulator in supplying the brakes, and also maintains a reserve pressure for brake operation in the event of pressure loss form the damping accumulator or failure of the motor-driven pump. A pressure switch senses system pressure and regulates the operation of the motor-driven pump automatically without cockpit switching so that a predetermined working pressure is always maintained by the two accumulators.

The emergency handpump is provided for use in the event of failure of the motor-driven pump. It may be used manually to pressurize the accumulators or to operate the wing flaps and nosewheel steering directly. The two accumulators are precharged with nitrogen or compressed air, which is facilitated by a charging valve and a pressure gage adjacent to each accumulator. The main power components of the hydraulic system (motor-driven pump, reservoir, accumulators and their pressure gages) are combined in a power package mounted on the fuselage structure beneath the flight compartment floor. It is accessible for servicing from both sides of the fuselage through access doors. CHECKING FLUID LEVEL (1) Insure that brakes are off and flaps are up (2) Insure discharge of brake and system damping accumulators by operating the brakes and wing flaps until accumulators are completely discharged (brake and hydraulic pressure inidicators read zero). (3) Unscrew hydraulic reservoir filler cap and check that fluid level is between REFILL and FULL marks on dipstick. Add hydraulic fluid, Specification MIL-H-5606, as necessary.



The hydraulic system



HYDRAULIC POWER PACKAGE The hydraulic power package supplies hydraulic pressure to operate the wing flaps, wheel brakes, nosewheel steering systems, and for lowering and raising the ski actuators, when fitted. The package consists of a tray containing the reservoir, brake and damping accumulators, indicators and charging valves, electric motor-driven hydraulic pump, system filter, system pressure relief valve, brake pressure relief valve, system pressure switch, and associated hydraulic lines and check valves. The package is mounted on the fuselage structure beneath the flight compartment floor and is accessible form both sides of the fuselage through access doors. HYDRAULIC RESERVOIR. The hydraulic reservoir is a cylindrical tank containing a filler neck, a filter screen, a cap and dipstick assembly, two suction outlets, a return inlet, and a vent. The reservoir is bolted to the power package tray. The filter neck, filter screen, and cap and dipstick assembly are accessible through the left access opening. DAMPING ACCUMULATOR. The damping accumulator is incorporated in the hydraulic system to damp out pressure surges in the hydraulic system and provide a secondary source of operating pressure. The accumulator is secured to the power package tray by two clamps. A combined indicator and air charging valve for the damping accumulator is mounted on a bracket on the left side of the power package.

WHEEL BRAKES ACCUMULATOR. The wheel brakes accumulator provides the wheel brakes with supplemental pressure in an emergency or during limited periods of peak demand, and provides pressure when the hydraulic system is shut down while parked. The accumulator is secured to the power package tray by two clamps. A combined indicator and air charging valve for the wheel brakes accumulator is mounted on a bracket on the left side of the power package. CHARGING THE WHEEL BRAKES ACCUMULATOR The wheel brakes accumulator is mounted on the hydraulic power package located beneath the flight compartment floor. A remote air charging valve and pressure indicator are located at the left of the accumulator. To charge the accumulator, release accumulator hydraulic pressure by operating the brake pedals, then charge the accumulator with nitrogen or dry air to 700 to 800 psi. CHARGING THE DAMPING ACCUMULATOR The damping accumulator is mounted on the hydraulic package alongside the wheel brakes accumulator, and is accessible through the same access door. To charge the accumulator, release accumulator hydraulic pressure by operating the wing flaps, then charge the accumulator with nitrogen or dry air to 700 to 800 psi.



The hydraulic package with the wheel brakes and damping accumulators.



FLIGHT CONTROLS The flight controls are conventionally operated through pulley and cable systems and mechanical linkage by a control column, control wheel rudder pedals. The control column is of a dual Y configuration located on the aircraft centerline with a control wheel pivoted at the upper end of each arm. The ailerons lower with the wing flaps and their degree of movement increases proportionately with flap deflection. The ailerons move differentially at any flap position. The left elevator, rudder and left aileron are equipped with flight adjustable trim tabs, and the right elevator with a trim tab that is interconnected with the flaps. A geared tab is installed on each aileron and on the rudder. CONTROL COLUMN. The dual Y configuration control column mounts pilot's and copilot's control wheels that are interconnected through a chain and sprocket system and linked to the aileron control system. A double acting spring strut attached to the forward side of the copilot's arm of the control column assists the return of the aileron controls to the neutral position. A control arm attached to a transverse torque tube on the base of the control column is connected to the elevator control system. A nosewheel steering lever on the control column pivots about the hub of the pilot's control wheel. A microphone switch is mounted on each control wheel.

GUST LOCKS Gust locks are provided to secure the main flight control surfaces (ailerons, rudder and elevator) while the aircraft is parked. The rudder system is provided with a built-in system controlled by an operating lever to the flight compartment floor in front of the pilot's seat. Both aileron and elevator gust locks are removable struts (part of flyway kit) which engage fittings on the instrument panel structure, control column and floor. On aircraft with Mod 6/1676 incorporated, the elevator gust lock strut locks the control column in a forward position. The rudder gust lock must be engage and the handle held up until the vertical strut is installed. RUDDER GUST LOCK. The rudder gust lock is a permanently installed mechanical linkage beneath the flight compartment floor which secures the rudder in the neutral position. The gust lock linkage consists of a hinged lever, connecting rods, a bellcrank and a spring-loaded lever. The hinged lever is flush with the flight compartment floor and when lifted to select rudder gust lock, operates the connecting rods to engage the spring-loaded lever in a slot in the rudder control quadrant. A spring connected to the spring-loaded lever, ensures that the lock cannot be engaged unless selected.



Gust lock installation



PITOT STATIC/PNEUMATIC SYSTEMS PITOT STATIC SYSTEM The pitot-static system supplies the pilot's and copilot's airspeed, altitude and vertical speed indicators. The system, (Mod S.O.O. 6075), consists of two independent pitot systems and two independent balanced static systems to operate the pilot's and copilot's airspeed indicators, altimeters, and vertical speed indicators. The pitot systems consists of two electrically-heated pitot heads, each installed on a mast projecting from the left and right side of the aircraft nose section. The left side serves the pilot's airspeed indicator and the right side serves the copilot's airspeed indicator. Each pitot head is connected to its respective airspeed indicator by a flexible hose and a moisture trap is installed in the lowest part of each line adjacent to the mast. The static systems consist of four static vents, two on each side of the aircraft nose section, one above the other. The lines from each static vent are routed to the static manifold where the lines from the upper vents interconnect and the lines from the lower vents interconnect. Static pressure from the lower vents is routed from the manifold to the static valve by a single line where it is switched to the pilot's flight instruments with the valve set to NORM. Static pressure from the upper vents is routed by one line directly to the copilot's flight instruments and by a second line to the static valve. When the static valve is set to EMER, static pressure to the pilot's flight instruments is switched from the lower vents to the upper vents and the static pressure from the lower vents is blanked off.

Pitot Heater Operational Test 1. Conect external power scource to aircraft. 2. Set BATTERY - OFF - EXTERNAL switch to EXTERNAL and switch on DC MASTER switch. 3. Insure PITOT HTR L and R circuit breakers are engaged. 4. Switch on pitot heater and check that pitot head warms up. CAUTION: DO NOT LEAVE PITOT HEATER SWITCHED ON FOR MORE THAN ONE MINUTE. 5. Switch off DC MASTER switch and BATTERY - OFF - EXTERNAL switch. Disconnect external power scource.



Pitot-static system layout



ELECTRICAL POWER Aircraft 28-volt dc electrical power is provided by two starter-generator systems in conjunction with two 24-volt batteries (main and auxiliary). The dc system is a single wire type installation using the airframe structure as ground return. The main battery supplies all dc power necessary when the generators are off-line and is also used to start the engines. The auxiliary battery provides an independent source of power for the engine start control and ignition circuits. Aircraft alternating current (ac) is provided as follows: Two 250 VA static inverter system OR Two 400 VA static inverter system (Special Order) When equipped with 250 VA static inverters, each inverter gives 115/26-volt, 400 Hz single phase outputs from a 28-volts dc input. With transformers fitted, the 250 VA system can provide an additional 26-volt output to cater for extra ac instruments. When equipped with 400 VA static inverters, each inverter gives a 115-volt, 400 Hz single phase output from a 28-volt dc input. The 400 VA system must be fitted with transformers to provide a 26-volt ac output. An external 28-volt dc power source can be connected to the aircraft for engine starting and test purposes.

Dc circuits are protected by thermal push-pull circuit breakers and the ac circuits by fuses. APPLICATION OF EXTERNAL POWER The external power receptacle is located on the left side of the fuselage beside the baggage compartment door. A spring loaded flap is fitted over the power receptacle to protect it from dirt and weather. (1) Insure that EXTERNAL / BATTERY switch is set to OFF. (2) Connect 28 - volt dc power source to aircraft external dc power receptacle. (3) Insure circuit breakers are engaged. (4) Set EXTERNAL / BATTERY switch to EXTERNAL. (5) Set DC master switch to MASTER; set BUS TIE switch to NORMAL. CAUTION: Do not commence work on aircraft electrical equipment with electrical power connected, unless work consists of tests requiring use of electrical power.



Location of electrical system switches in the cockpit and external power receptacle



DC MASTER SWITCH. A two-position dc master switch is located on the overhead console. The switch has a center-off position marked OFF and an on position marked DC MASTER. This switch controls the power supply to all buses, except the main battery bus, in conjunction with the external/battery switch and the bus tie switch; when the dc master switch is at OFF, no power is supplied to these buses regardless of the position of the external/battery switch. At DC MASTER position, the switch connects power from an external power source or the battery, (as determined by the external/battery switch position) to left and right buses or left bus only (as determined by the bus tie switch position), or the generators to their respective buses. EXTERNAL/BATTERY SWITCH. A three-position external/battery switch, located on the overhead console has positions marked EXTERNAL, OFF, and BATTERY. It connects the various sources of electrical supply to the buses dependent upon the position of the dc master switch. At EXTERNAL the switch connects an applied external power source the system and isolates the battery. At BATTERY, the battery powers the electrical system if the generators are inoperative or their voltage output is less than that of the battery. One or both generators are connected to the system at BATTERY position if the generator output exceeds that of the battery.

EXTERNAL POWER RECEPTABLE. An external power receptacle is located on the fuselage left side aft of the cargo doors; it is covered by a spring-loaded access panel. When external power is connected, the external/battery switch is at EXTERNAL, and DC MASTER switch at DC MASTER, an external power relay is energized (provided the generators are not operating) which connects the external power to the left and right dc buses (provided the bus tie switch is at NORMAL). DC VOLTMETER. A dc voltmeter is mounted to the right of the emergency panel. The voltmeter dial is marked DC VOLTS and has a scale graduated in 1-volt increments from 0 to 30, with numerals at 10-volt intervals. It is connected to the left dc bus and indicates the voltage at the bus, supplied from any of the power sources. It is protected by a VM circuit breaker on the main circuit breaker panel.



Main battery installation and auxiliary battery



EXTERIOR LIGHTING SYSTEM The exterior lighting system comprises position lights, landing lights, and an anti-collision light; wiring is installed for the addition of a special order taxi light and wing inspection lights. A lower anti-collision light and the necessary wiring, and strobe lights and the necessary wiring, can also be installed on special order. All exterior lights are powered from the left or right dc buses. POSITION LIGHTS. The conventional position (or navigation) lights consists of red and green wing tip lights and a white tail light. The lights are controlled by a switch marked POSN on the overhead console. The circuit is powered from the left dc bus and protected by a POSN LTS circuit breaker on the main circuit breaker panel. ANTI-COLLISION LIGHTS. Two white strobe type anti-collision lights are part of the position light assembly installed in each wing tip. A red rotating anti-collision beacon light is mounted on the top of the vertical stabilizer. The white strobe lights are controlled by a switch marked ANTI COLL on the overhead console; the circuit is powered from the left dc bus and is protected by the ANTI COLL LT circuit breaker on the main circuit breaker panel. The red rotating anti-collision beacon light is controlled by a switch labeled BEACON on the overhead console; the circuit is powered from the left dc bus and is protected by the BEACON LT circuit breaker on the overhead circuit breaker panel. On special order a lower rotating anti-collision beacon light can be mounted on the underside of the rear fuselage. It is controlled by be

BEACON switch on the overhead console; the circuit is powered from the right dc bus protected by the ANTI COLL LT LOWER circuit breaker on the overhead circuit breaker panel. On earlier aircraft, without strobe lights, the rotating anti-collision beacon light is controlled by the ANTI COLL switch on the overhead console. LANDING LIGHTS. Two 250-watt landing lights are installed, one in each wing leading edge outboard of the engine. The lights are controlled by two switches marked LANDING LIGHTS, LEFT and RIGHT, on the overhead console, and the circuits are powered from the left and right dc buses and protected by LDG LT L and LDG LT R circuit breakers on the main circuit breaker panel. TAXI LIGHT. On special order a 100-watt taxi light can be installed on the nosewheel fork. The light is controlled by a switch marked TAXI on the overhead console and the circuits are powered from the right dc bus and protected by a TAXI LT circuit breaker on the main circuit breaker panel. The taxi light is standard equipment on later airplanes. WING INSPECTION LIGHTS. On special order, two 40-watt wing inspection lights are installed, one on the outboard side of each engine nacelle. The lights illuminate the leading edges of the wings and are controlled by a switch marked WING INSP LT on the overhead console. The circuit is powered from the right dc bus and protected by a WING INSP LT circuit breaker on the main circuit breaker panel.



Location of External Lights Switches on the Overhead Panel



FIRE DETECTING SYSTEM A fire detection system is provided for each engine. In the event of an engine fire, thermal detector unit(s) installed in the engine nacelle actuate and complete an electrical circuit to illuminate a warning light in the appropriate fire extinguisher handle, located on the emergency panel in the flight compartment. Each engine fire detecting system consists of four thermal units connected to two wire loops, a fire warning light, magnetic and thermal circuit breakers, the associated electrical wiring, and a circuit test switch which is common to both engine systems. On aircraft with Mods S.O.O 6123 or 6/1277 embodied, and alarm bell is installed. FIRE DETECTING THERMAL UNITS. The four fire detecting thermal units (bi-metallic strips) for each engine, and their connecting wire loops, are located on the forward and rear fireseals in the engine nacelle, and on a nacelle frame at station 54.60. The two forward thermal units actuate at a temperature of 4500F, and the two rear thermal units at a temperature of 3000F.

FIRE DETECTION WARNING LIGHTS AND SWITCHES. The fire detection lights are integral with the fire extinguisher handles located on the emergency panel, the handles are annotated LEFT ENGINE and RIGHT ENGINE. Four circuit breakers, two for each engine system, are provided on the circuit breaker panel; two are thermal circuit breakers labeled FIRE DET L and R, and two are magnetic circuit breakers labeled FIRE DET FAULT IND L and R. When the fire warning bell is installed, a two-position fire bell mute switch is provided. The switch labeled FIRE BELL MUTE, is located below the pilot's flight instrument panel (below the main circuit breaker panel on aircraft with Mod S.O.O 6123 or the retrofit version of Mod 6/1277). FIRE DETECTING SYSTEM MAGNETIC CIRCUIT BREAKERS. Two of the system circuit breakers, one in each engine circuit, are of the magnetic type, having a single set of switch contacts with a center moving arm mechanically linked to its magnetic coil.



Fire detecting system - components locations



FIRE EXTINGUISHING SYSTEM The fire extinguishing system consists of two independent systems, one for each engine. Each system has a handle located in the flight compartment, which, when pulled, electrically discharges a fire extinguisher bottle. The extinguishing agent from the bottle is directed by a discharge nozzle into the appropriate engine accessory compartment. Each engine fire extinguishing system consists of a fire extinguisher bottle with discharge pipes and indicator discs, and a fire extinguisher handle and switch. FIRE EXTINGUISHER BOTTLES. A fire extinguisher bottle containing FREON 1301, is installed in each engine nacelle behind the firewall. A main discharge pipe extends from a valve on the bottom of the bottle through the firewall to the engine accessory compartment. Two additional pipes, one connected to the bottle and the other to the main discharge pipe, terminate at indicator discs on the side of the nacelle.

FIRE EXTINGUISHER BOTTLE INDICATING DISCS. Two fire extinguisher bottle indicating discs are provided for each bottle, one yellow, and the other red; they are located on the inboard side of each engine nacelle. FIRE EXTINGUISHER HANDLES. A fire extinguisher handle identified FIRE PULL, is provided for each engine system. Each handle is interconnected with a switch mounted on the rear of the emergency panel. OPERATION The handle operated switch of each system is supplied with 28-volt dc from its respective FIRE EXT L or FIRE EXT R 5-ampere circuit breaker. When the FIRE PULL handle is pulled, the switch is operated, and the circuit is completed to the appropriate bottle valve which is opened by the action of a squib assembly. The fire extinguishing agent is then expelled through the valve and main discharge pipe into the engine accessory compartment. The yellow indicator disc, which is backed up by a piston assembly in the pipe connected to the main discharge pipe, is pushed out when the extinguisher operates. When the fire bottle is discharged due to thermal expansion, the red disc is blown out.



Fire extinguishing system - components location



HAND-OPERATED FIRE EXTINGUISHERS A hand-operated fire extinguisher is located on the sloping bulkhead behind the copilot's seat. On installations incorporating Mod 6/1542, the extinguisher is located on the aisle floor beside the copilot's seat. The extinguisher contains a frangible disc which, when broken, allows bromochlorodifluoromethane (BCF) contents to discharge under pressure. On special order a water/glycol hand extinguisher may be installed, in which case it is located on the right-hand cabin door.

EMERGENCY EQUIPMENT FIRST AID KIT. A first aid kit is stowed on the cabin wall aft of the right cabin door or on the divider panel immediately forward of the left cabin doors.



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