Navajo Ground School

171
Navajo Training Program Sunday, June 12, 2011 Navajo Chieftain Initial Training Course

Transcript of Navajo Ground School

Page 1: Navajo Ground School

Navajo Training Program Sunday, June 12, 2011

Navajo Chieftain Initial Training Course

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Navajo Training Program Sunday, June 12, 2011

Aircraft History

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History Of The Navajo

The Piper PA-31 Navajo is a family of cabin-class, twin-engine aircraft designed and built by Piper Aircraft for the general aviation market, most using Lycoming engines.

Targeted at small-scale cargo and feeder liner operations and the corporate market, the aircraft was a success.

It continues to prove a popular choice, but due to greatly decreased demand across the general aviation sector in the 1980s, production of the PA-31 ceased in 1984

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Navajo Chieftain

In September 1972 Piper unveiled the PA-31-350 Navajo Chieftain, a stretched version of the Navajo B with more powerful engines and counter-rotating propellers to prevent critical engine handling problems. The fuselage was lengthened by 2ft 0in (0.61m), allowing for up to ten seats in total.

Variants of the Lycoming TIO-540 developing 350 hp (261 kW) were fitted to the Chieftain, with an opposite-rotation LTIO-540 installed on the right-hand wing; MTOW was increased to 7,000 lb (3,175 kg)

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General Characteristics

Crew: 1 or 2

Capacity: 8 passengers

Length: 34 ft 7½ in (10.55 m)

Wingspan: 40 ft 8 in (12.40 m)

Height: 13 ft 0 in (3.96 m)

Wing area: 229 ft² (21.3 m²)

Empty weight: 4,114 lb (1,866 kg)

Max takeoff weight: 7,000 lb (3,175 kg)

Power plant: 2× Lycoming TIO-540-

J2BD,six-cylinder horizontally-

opposed, air cooled 350 HP each

Propellers: Three blade, metal, fully

feathering, Hartzell propeller

Performance

Maximum speed: 234 knots at

15,000 ft

Cruise speed: 214 knots econ cruise

at 20,000 ft (6,100 m)

Stall speed: 74 knots flaps down

Range: 1,103 nm

Service ceiling: 27,200 ft

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Navajo Chieftain Aircraft Systems

Description and Operation

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Description & Operation

The Piper Navajo is a twin-engine, non pressurized, multi purpose aircraft with tricycle retractable landing gear.

The aircraft are certified for normal category and approved to operate in Day, Night VFR and IFR conditions.

Certified for Known icing conditions when equipped with appropriate optional equipment.

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Airframe

All Aluminum construction

Conventional semi monocoque structure

Cantilever all metal wing with an I beam MAIN SPAR which extends through the aircraft through the fuselage

A rear SPAR is also incorporated for the attachment of trailing edge control and high lift devices

A Front spar completes the BOX section of the Wing.

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Nose Section

Nose Section houses the weather Radar.

In addition there is a forward baggage compartment which can hold 200 Lbs.

Also contained within the nose compartment is access to the main Battery, Oxygen System, Janitrol Heater and Brake fluid reservoir.

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A bulkhead is provided to protect avionics in the aft section, the hydraulic power pack assembly, starter vibrator, weather radar and gyro systems.

The Battery is contained forward of the baggage area.

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•Access to the baggage area requires a key. Unlocking a key will activate the handle.

•Prior to flight the handle needs to be returned to the stored position and the lock re-engaged.

•A courtesy light illuminates when the compartment is opened.

•Power to the light is supplied directly from the battery bus and is hot wired. The Master Switch need not be on.

* Caution should be taken concerning the light switch. Failure in the on position could deplete the battery overnight.

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Cabin Lighting and Air Outlets

Individual controlled air outlets and reading lights are installed overhead of six cabin seats.

Ventilation outlets are mounted above the seats in the cabin area and flight compartment

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Cabin Access

•There are as many as two doors to access the aircraft (Crew Door and the Main Door)

•The main door is composed of two hinged doors.

•One opening upward containing a window

•Second opening downward housing the air stair. The lower portion of the door is supported by two covered cables. BOTH CABLES must be in place before weight is applied to the door.

•A Cabin Door Unsafe Annunciator will illuminate of not secured properly.

•Close top door first. Slide the latch. Then close bottom portion of door. Ensuring that the aft seat seatbelt is not hanging out the door.

A cargo door is optional..

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Load Limits

The cabin area can be configured for cargo loading with the floor bearing strengths and areas for each section as described.

Per foot of track 200 Lbs.

Per Track 900 Lbs.

Per Tie Down Ring 200 Lbs.

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Aft Fuselage, Empennage and wings

•The Empennage is a cantilever system which incorporates the Horizontal Stabilizer, Elevator, Vertical Stabilizer and Rudder.

•The 406Mhz ELT is stored in aft section of the aircraft with the antenna mounted on the starboard side of the aircraft.

•The tail contains anti ice boots on both the horizontal and vertical stabilizers.

•The Elevator and Rudder both have their own adjustable in flight trim tabs

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Wing Section •

metal, cantilever, semi-monocoque structure.

•Each wing section incorporates an I-beam main spar, that extends into the fuselage. Two spars are bolted together with high strength butt-plates, effectively creating a continuous main spar.

•A Rear Spar is also used for the attachment of trailing edge high lift devices (Flaps) and control devices (Ailerons)

•A Front Spar is also incorporated into the wing box design.

•All three spars connect to the fuselage.

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Flap Assembly

•Flaps are powered by a single motor located under the floor.

•Power to each flap is transmitted via a flexible shaft.

•Asymmetrical Flap extension is prevented by limit switches.

•Max asymmetry is 5 Degrees (Calaco)

•Max Asymmetry 9 Degrees (Duke).

•The Flap Annunciator indicates asymmetry.

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Ailerons •The ailerons are all metal and are both statically and aerodynamically balanced.

•Aerodynamic balance is located at the outboard end of each aileron.

•A trim tab is incorporated in the right aileron and is adjustable from the cockpit pedestal.

The engine nacelles are an integral part of the wing and provide an aerodynamic structure for engine mounting. Two wing locker baggage compartments can also be housed in the nacelle.

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Landing Gear

Fully retractable, hydraulic, tricycle configuration landing gear system.

Each, nose and main gear assembly employees straight, air-oil strut suspension and is actuated hydraulically through a mechanical control handle in the cockpit.

Gear handle in the gear down Neutral position

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Emergency Features • A fully removable Emergency window is located on the right side of the passenger

cabin at the first window aft of the copilots side window.

• To remove the emergency exit window, remove the Plexiglas cover over the handle; then pull the handle and push out on the window.

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Turbo Charged Engines

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Engine Description

• The Navajo Chieftain is powered by

• 2 Turbocharged Lycoming six cylinder, horizontally opposed engines

• 350 BHP horse power at 2575 RPM.

• The engines are counter rotating

• Maximum manifold pressure is measured in inches HG.

• The maximum MP is set during a maintenance action and is affected by the seasons.

• It will be re-adjusted during

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• 2575 Max RPM

• Manifold pressure decreases @ 0.64 Per 1,000 ft. above 15,000

• Pressure decreases at 2.2 in. Hg. Per 1,000 above 22,300 to 24,000 ft.

Limitations

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Engine Controls The control pedestal is located between the pilot and copilot station in the cockpit. The controls are in groups of two, Throttle, Propeller, and Mixture. In addition there is a friction lock to prevent the levers to creep when set.

1. Throttle Control

2. Propeller Control

3. Mixture Control

4. Friction Lock

1

2

3

Friction Lock

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Alternate Air Control •An automatic door which is magnetically held shut will open should the air filter or intake become blocked. Suction from the turbo will open the door.

•If the turbo is operative but the automatic function fails the door can be opened manually using the Manual Alternate Air Pull Handles.

•This Assumes that the turbo is functioning. Should the Turbo seize there is another alternate air door.

Manual Alternate Air Magnetic Door

Caution: Do not open on ground when engines are running.

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Normally Aspirated Alternate Air Intake

Should the turbo seize the engine will automatically revert to normally aspirated.

the alternate air intake "opens" when the manifold pressure drops due to turbo failure ,combined with continued engine operation lowering intake manifold pressure .

The aircraft will be reduced to about 75% of its max power and will be altitude dependant.

In addition the mixture with the turbo failed will be TOO rich. In the event of a turbo failure lean engine until engine runs smooth.

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Cowl Flap Controls • The Cowl flaps are electrically operated using switches located at the base of the centre pedestal.

•Each Cowl flap has its own independent motor.

•Each switch is spring loaded to the neutral position.

•When Selected up the cowl flaps move toward fully closed. Intermittent stops are allowed

•Selecting down on the switch moves the cowl flaps toward full open.

•A position indicator indicates the position of the associated cowl flap.

•CAUTION Cowl flaps should not be selected from Full Close to Full Open IN FLIGHT. The excessive cooling can cause Cylinder Heads to Crack.

Video Clips

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Cowl Flap Continued •During ground operations regardless of temperature the cowl flaps are to be open.

•There is little to no ram air provided by the spinning propeller while the engine is running on the ground. Cooling is brought about by the negative pressure caused by the air flowing over the cowl flap.

•The cowl flaps open in flight will reduced speeds by apprx 5-8 knots. Monitor engine parameters to ensure that cylinder head temperatures are around the nominal 380 degree mark.

before target altitude will cause the aircraft to level at the desired altitude. Set cruise power in short order to prevent excessive cylinder head temps.

Video Clips

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Engine Starting System •The Navajo engines incorporate a 24 volt D.C. gear driven starter unit.

•The starter System includes the battery master switch, a rocker type [LEFT or RIGHT] starter switch and interconnected starter solenoid, and circuit breakers.

•The starters are energized by engaging the [ENGINE STARTER] switch located in the overhead switch panel between the engine magneto switches.

•To operate, push the switch to the [LEFT or RIGHT] position and hold down until the selected engine has started. After engine start release the switch allowing it to return to is spring loaded center [OFF] position.

is sent to the Left Mag to retard timing. See more in Ignition

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Engine Instruments

Engine Status is monitored using

•Manifold Pressure Dual Needle Gauge

•Tachometer Dual Needle Gauge

•Fuel Flow Single Needle Gauge

•EGT Exhaust gas temperature Single Needle Gauge

•Oil Pressure Clustered with CHT and Oil Temp (Per Engine)

•Oil Temperature - Clustered with CHT and Oil Temp (Per Engine)

•Cylinder Head Temperature - Clustered with CHT and Oil Temp (Per Engine)

•Fuel Pressure

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Manifold Pressure Gauge

•Green Arc 20 to 41 Inches

•Cautionary Range 41-49 Inches

•Limitation 49 inches.

•Turbo supplies pressure beyond naturally aspirated levels. Due care should be taken when increasing throttle to allow the to spool slowly and evenly.

Engine Tachometer

•Indicates engine and propeller RPM since there is no gear between the engine and propeller. The instrument scale ranges from 0 to 3500 RPM with a green arc from 500 to 2400 RPM and a red line at 2575.

•Each engine will have sweet spots in terms of RPM where it runs smoother with less vibration. So RPMs are not hard values. 2400 may give slight vibration where 2440 is smooth.

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•Calibrated in gallons per hour (GPH)

•The scale is 10 GPH and 46 GPH.

•There are no colored markings on this gauge.

•Minimum Fuel Flows at Takeoff Must be observed or a RTO is required.

•Insufficient fuel flow at TO will result in high CHT and Detonation and possible engine failure!

•Fuel pressure measured in Psig

•Minimum fuel pressure indicated by a red radial 34 psig.

•Normal operating Range 34-45 Psig

•Max Fuel Pressure 55 Psig

Fuel Flow Indicator

Fuel Pressure Gauge

Psig = Pounds Per Square Inch GAUGE

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Exhaust Gas Temp.

•The [EGT] gauge is of significant value in both practical and economical operation of the engine.

•The thermal probes are self powered using thermocouples.

•There is one EGT sensor on the exhaust manifold. Allowing for the monitoring of all 6 Cylinders

Instrument scale range is from 700⁰F to 1800 ⁰F.

Normal Operating Range 0 ⁰F to 1650 ⁰F

There are Three Different Limitations For EGT.

1350 Minimum For Descents

1550 Maximum For Climbs

1650 Maximum For Cruise.

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EGT

85%

•Piper allows for leaning at this power setting. Mixtures should be leaned to 1500 degrees or 30 GPH which ever occurs first.

Ensure cylinder head temperatures are within nominal range a few moments after leaning to

ensure that sufficient fuel is being supplied.

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Oil Temperature

•Oil Temperature is taken form the output side of the oil cooler prior to entering the engine lubrication channels.

•Scale ranges from 50⁰F to 250⁰F.

• Yellow arc Caution Range - 50⁰ F and 120⁰ F.

•Green arc Normal Operating Range is 120 ⁰F to 245⁰F.

•Red line indicates maximum allowable operating temperature at 245⁰F.

• During cold weather starts ensure the engine speed is kept at idle until oil pressure has had a chance to build.

• RPM is restricted to 1200 or less until oil temperature is in the green arc.

• Oil is consumed at the average rate of .25 qts per hour. Always ensure sufficient oil before operating the engines.

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Cylinder Head Temperature

•The cylinder head temperature is taken from Cylinder #6 on the aft most Cylinder of both Engines.

•Temperature is taken here because it will be the hottest of the cylinders.

•Gauge Scale 100⁰F to 500⁰F

•Green arc Normal Operating Range 100⁰F to 475⁰F

•Red line Maximum never exceed 475⁰F.

Typical temperatures in cruise should be found to be between 375-400 degrees. Temperatures that exceed this should be carefully monitored and reported upon landing. Take steps to cool engines to below 400 degrees to maximize longevity.

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Oil Pressure Gauge

•Oil pressure is measured by an oil-filled instrument line from a sample tap to the engine oil pressure gauge in the cockpit.

There is no transducer or mechanical interface.

Gauge Scale - 0 to 200 psi

Minimum operating red line - 25 psi

Yellow caution arc - 25 psi to 60 psi

Green arc 60 psi to 90 psi

Secondary Yellow arc 90 psi to 100 psi

Red Line Maximum Pressure - 100 psi.

Fittings can be inspected inside the cowling. Failure of the sample line inside the cockpit will result in oil flowing into the cockpit unrestrained.

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Oil Pressure Gauge

Due to the design of the oil pressure sensing of the Navajo and the temperatures at which we operate in the winter it can take quite a while before oil pressure begins to indicate.

If abnormally long. Advise maintenance at your base so that they may flush the line.

We run 20w50 oil all year long. This on occasion (-35 in YTH) will result in higher than normal pressures at startup. Be patient and vigilant; they should reduce once oil has had a moment to warm up.

ENGINE RPM IS TO BE KEPT AT IDLE UNTIL OIL PRESSURE IS AT THE TOP OF YELLOW ARC.

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Fuel Injection System

•Fuel injectors are continuous pressure/ flow type.

•The system is controlled by the FACU (Fuel Air Control Unit)

•Air passes over the butterfly valve and generates a negative pressure in the venturi.

•As the throttle is moved forward it allows more fuel which is under pressure to flow to the injectors.

•This allows for engine priming. Advancing the throttle allows fuel to flow through the SPIDER to the injectors.

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Precise and automatic fuel leaning is afforded over a wide range of altitudes and

power settings without addressing the need for constant manual alterations of

the mixture control.

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Fuel Injection Pump

•The high pressure fuel pump is driven by the engine.

•The high pressure pump delivers pressure between 34-55 PSI to the injectors.

•In order to prevent cavitations of the high pressure pump low pressure boost pumps ensure a constant flow of fuel at 10 PSI to the high pressure fuel pump.

•From the fuel pump pressurized fuel goes to the FACU and is then distributed through the spider lines to the intake port of each cylinder.

•Vapor is routed through a vapor return line back to the recirculation chamber of the pump.

Located under an access panel in the bottom of the wing

outboard of the engine is the Emergency Boost Pump.

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Fuel Injection Pump

•Cavitation

•The formation of vapour bubbles of a flowing liquid in a region where the pressure of the liquid falls below its vapour pressure.

•Inertial cavitation is the process where a void or bubble in a liquid rapidly collapses, producing a shock wave. Such cavitation often occurs in control valves, pumps, propellers, impellers.

•Inertial Cavitation will destroy a fuel pump. Ensure that the main fuel pumps are never starved for fuel.

•Unless in event of emergency do noT

Select Emergency Fuel Shutoff

Cavitation of high pressure fuel pumps can occur if the boost pumps fail to supply 10psi of boost pressure. In addition pulling the fuel shutoff valve will starve the fuel pump inconsistently and result in cavitation.

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Low Pressure Boost Pumps

Each engine is fitted with a low pressure boost pump which provides 10PSI of positive pressure from the moment the master is turned on.

This 10PSI can be used for priming but is primarily in place to provide a constant flow of fuel to the high pressure engine driven fuel pump and thus prevent cavitation.

If the fuel boost pump CB is pulled the annunciator will illuminate or if the pump has failed. With the CB in and the master switch on the fuel pressure should be 10PSI and the annunciator should be extinguished.

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Low Pressure Boost Pumps

During initial checks fuel is turned off so at to check cross feed. Once fuel is turn off a few moments later the FUEL BOOST INOP annunciator will illuminate indicating no flow to the pump.

Once fuel is restored the annunciator will extinguish.

This annunciator will illuminate when a tank is run dry. Quickly switch to full tank and activate emergency boost pump to prevent engine surging. (3-5 seconds after lamp before engine is starved for fuel at cruise power)

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High Pressure Injector Pump Positive displacement VANE

type pump.

Engine driven

High pressure with greater capacity than engine demand.

Fuel enters at swirl chamber vapour separator

Pump outlet fuel pressure passes through an orifice affected by turbo output pressure and before entering relief valve chamber thus making pump delivery pressure proportional to engine speed.

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Fuel Air Control Unit

The purpose of the fuel air control unit is control the amount of air admitted into the inlet manifold and meter the fuel in correct quantity to provide correct ratio.

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Fuel Air Control Unit

The location of the air throttle valve is in the intake manifold and is mechanically connected to the fuel metering valve housed in the fuel control unit. Moving the air throttle valve thus determines the position of the fuel metering valve and ensures the correct amount of fuel is provided.

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Fuel Manifold Valve (Spider)

Manifold valve receives fuel from the metering unit.

When pressure reaches a predetermined level fuel flows to injectors.

Serves as a positive cut fuel cutoff when the engine is shutdown.

Normally located on the top of the engine somewhere.

The vent for the fuel spider is overboard. If you see fuel leaking from this vent it means the manifold is faulty.

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Fuel Injection System

Multi nozzle continuous flow

Fuel flow is proportional to fuel pressure.

Matches engine requirements

Changes in throttle

Engine speed

Deck pressure

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Injection Nozzles

•A single fuel injector per Cylinder

•The air bleed type injector is used in the Navajo.

•When the engine is running flow from the nozzles is continuous and fuel enters the engine when the intake valve opens.

•Since the size of the injector nozzle is fixed fuel pressure determines the amount of fuel that enters the engine. Therefore accurate fuel flow can be determined by the fuel pressure at the manifold.

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Engine Oil System

•Wet Sump Type

•12 U.S. quart capacity.

•Each engine Has its own Dipstick. Ensure the appropriate dipstick is present

Oil dipsticks are dual scale

and interchangeable. Ensure

you read oil level for

appropriate side.

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•A gear driven oil pump, powered by the engine accessory section provides oil at 25 to 100 psi to all moving parts, the propeller, and the turbocharger waste gate controller.

Normal operating pressure is 60 to 90 psi. Oil pressure below the 25 psi red line minimum is hazardous to engine operation.

The low pressure may be caused by an inaccurate or inoperative gauge, a fouled or obstructed oil pressure gauge line, an insufficient supply of oil to feed the pump, a low oil volume at a high engine operating temperature, an inoperative or damaged oil pump, a blockage in the lubrication system or in extremely cold soaked engines or congealed oil in the lubrication and or engine gauge system.

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•Oil pressure is required to operated the waste gate controller. Low oil pressure will

applying full power due to insufficient pressure to close waste gate.

•Additionally. If oil pressure is too low the engine will fail to make full power because the waste gate will fail to close entirely under WOT (Wide Open Throttle)

Oil for cooling Turbo Bearing

Waste Gate Controller Below Turbo.

Before shutdown engines must be run at idle for no less than 2 minutes to allow for turbo cool down. Failure to do so will

failure. Mixture will be rich to maximize cooling.

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•High oil pressures are generally the result of cold or high viscosity states of the oil.

•Engine oil is screened immediately before leaving the sump and channeled to an oil filter for the final filtration process.

•Oil is cooled via a ram air heat exchanger (pictured left) using a thermostatic control bypass valve to route oil through or around the oil cooler.

•Normal operating range is 120° F to 245° F

•Minimum Temp 50° F

•Maximum Temp 245° F.

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Ignition System

•Each engine is equipped with dual ignition.

•Both mags are housed in a common unit.

•Each Mag has its own starter Vibrator

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Magnetos

Magneto operates on the principle that a rotating magnetic field created by a rotating magnet induces an electrical voltage in the primary winding. This voltage sets up a magnetic field that envelops the secondary windings. When the breaker points close the magnetic field around the primary coil collapses as voltage goes to ground. This generates a high voltage in the secondary coil. This voltage is routed to the spark plugs.

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•Each magneto operates entirely independent of the other and failure of one system has no effect on the other.

•Both mags share a common shaft. Failure of this shaft will result in engine failure.

•When in the ON position the magnetos are live or Ungrounded.

•The

starter switch is depressed.

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Vibrator This is required to start the engine. Does not operate once the engine is running and the starter switch is disengaged. The vibrator does two things. First It send sends electrical pulses 3V (DC) to the primary coil. The pulses cause a magnetic field to form and collapse

in each cylinder when the retard points are activated. Otherwise all voltage goes to ground.

effectively retard ignition from about 20 degrees before TDC to TDC allowing for easy starts.

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•The retard point is the centre lead into each of the mag on each engine.

•These points can become worn and make starting difficult. A slight turn of the centre lead or retard point could allow you to start the aircraft and make it home for maint.

•The maximum drop is 175 rpm

•The maximum difference between mags is 50 rpm.

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Ignition System Schematic

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Induction Air System The primary source of induction air for the turbocharger and the fuel injection system is through a filtered inlet on the lower right side of the engine nacelle.

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Dual Controller Turbo System

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Density Controller Ground boosted turbo

charger Produces greater than atmospheric MAP and maintains with an increase in altitude

Density controller senses air density to prevent turbo outlet pressure from the set maximum level.

Only active at wide open throttle.

• • Should MP exceed the pre-set value density

controller will cause an oil pressure bleed valve to open which will in turn open the waste gate and lower deck pressure

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Differential Pressure Controller

Controls waste gate or turbo outlet pressure at all throttle settings other than wide open.

At WOT the bleed valve is closed

As throttle closes bleed valve gradually opens.

Pressure drops and waste gate opens.

• • Should MP exceed the pre-set value density

controller will cause an oil pressure bleed valve to open which will in turn open the waste gate and lower deck pressure

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•In normal operation air flows over air filter and into the intake manifold then onto the turbo compressor.

•In the event of blockage an alternate air door is located on the upper right side of the engine.

•The door is magnetically held in place and is opened with excessive negative pressure.

•Air from this intake is not filtered prior to consumption.

•In addition to automatic function the alternate air prior to the turbo can be activated manually.

It is important to NOTE that alternate air should not be checked on the ground, to do so. Could result in turbocharger impeller damage.

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•The air in the induction system flows directly into the impeller area of the turbocharger.

•The waste gate requires a minimum oil pressure of 50 psi to function

•Compressed air from the turbo is expelled at the intake orifice of fuel injection system, differential pressure and density controller unit.

•This unit means the engines air consumption and meters an appropriate amount of fuel.

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Induction Continued

The unit measures engine air consumption by use of a venturi tube using the flow forces to control fuel flow to the engine.

The induction air then flows past the throttle body and into the induction manifold to each intake valve port.

At the intake port of the cylinder a metered/measured charge of fuel is introduced into the base of the intake valve as the seat of the intake valve begins to open and throughout its intake stroke.

After combustion, the exhaust gas is emitted through the exhaust valve into the exhaust port and manifold of the cylinder.

The exhaust gases are introduced into a manifold divider, part of the gases are channeled under exhaust pressure into the turbocharger turbine impellers to drive the turbocharger and out into the exhaust stream.

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The majority of the exhaust gases pass over the turbine induction manifold and flow past the waste gate controller where engine oil lines deliver commands from the density air controller and differential pressure unit to the exhaust bypass valve assembly which positions the waste gate valve.

The Waste Gate Unit controls the rotational velocity of the turbocharger centrifugal compressor section. The Density Controller controls the waste gate under full power application. The Differential Controller controls the operation of the waste gate under partial power conditions by increasing or decreasing oil pressure in the exhaust bypass valve assembly.

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Navajo Training Program Sunday, June 12, 2011 Induction Air System Schematic

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Propellers and Synchronizer

The Navajo is equipped with Hartzell all-metal, three bladed, constant speed, full feathering propellers.

A combination of nitrogen or air pressure, a spring, propeller blade counter weights, and governor. Regulate oil pressure controlling the pilot valve by the governor flyweights control the pitch of the blades.

No gearbox or gear reduction system is incorporated in this arrangement.

Propeller speed is selected by positioning the propeller pitch control levers in the flight cockpit.

Mechanical positioning of the governor through control linkage repositions the spring tension on the fly-weights in the governor body

Prop sync has been

removed from all of our

aircraft.

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A pilot valve is operated to allow more or less engine oil to enter the journal of the crankshaft.

Therefore, flow to the end of the crankshaft and into the propeller hub or out as appropriate, will reposition the propeller blades.

The position of the blades are altered in response to this pressure gradient, therefore, maintaining a constant engine RPM. The governor system is so designed that if oil pressure is lost the propellers will begin to feather.

CAUTION: Should the propellers begin to go to the feather position

because of lack of oil and the pilot does not immediately retard the propeller control lever to the feather position, locking pins located on the propeller, would at approximately 700 RPM lock the propeller and feathering would be impossible.

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72

Propeller Governor

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Chapter 3 Flight Controls

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Ailerons and Elevators The ailerons are connected to the flight controls by cables routed through the fuselage by bell crank and pulleys. The rudder is also connected to the aileron system by a spring assembly causing the rudder to automatically assist the turn when the ailerons are actuated.

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The elevator control surfaces are connected by a torque tube which is, in turn, connected to a bell crank in the tail section. The e bell crank is actuated by cables running through the empennage from the pilot and copilot control yoke.

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Aileron Trim

A trim tab provides aileron trim located on the right aileron. This tab is actuated by a push-pull rod attached to a jack screw type actuator in the wing. The actuator is driven by cables attached to the trim control knob on the control pedestal in the cockpit.

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Rudder Trim

The rudder trim is located on the control pedestal. This wheel controls a trim tab on the lower portion of the rudder, it also is connected by push-pull rod attached to a jack screw type actuator and cables that run through the fuselage.

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Elevator Trim

The elevator trim tab is located on the right elevator and is actuated by a push-pull rod attached to a jack screw actuator in the horizontal stabilizer. The actuator is driven by cables attached to the trim wheel in the cockpit control pedestal.

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Trim Controls

An optional electrical elevator trim, incorporating an emergency disengage switch is provided on the pilots control wheel and is located on the left hand side of the yoke. In addition to the manual application, an additional switch can be provided on the copilot yoke. Should a switch be located

On the copilot yoke, the pilot’s switch will override the copilot commands, should the pilot apply opposite control. Refer to the model concerning specific information for your aircraft.

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Nose Wheel Steering

The nose wheel is connected to the nose wheel by a bungee type spring assembly and cables. The nose gear engages the steering system when compressed by the weight of the aircraft during ground operations. As the landing gear is retracted on takeoff, the nose gear straightens for storage in flight.

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Ground Towing

The nose wheel can be turned up to 20 degrees left or right of center, using the rudder pedals. For steering angles greater than 20 degrees up to a maximum of 40 degrees differential power can bi used. While minimum turning distance can be achieved, excessive ware on tires would result.

Caution: should be taken when the aircraft is being towed, turning the aircraft

to the red turn limit is easily ignored. Further turning can result in damage to the nose wheel collar. A through ground inspection should be made in the presents of ground crew to ensure the collar was not broken or damaged prior to ground towing

THE SCISSORS WILL BE DISCONNCETED FOR ALL TOWING!!!!

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Flaps

Flaps on the Navajo’s are conventional flap design and are actuated by a pair of flexible cables attached to a jackscrew transmission assembly. Each wing flap is connected to one electric drive motor located in the fuselage center section just behind the pilot and copilot seats.

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The flap motor controls the wing flaps by manipulating the flap control lever located in the cockpit on the right side of the control pedestal. This selector switch provides variable positioning of the wing flaps. The increments are 0%, 15% or 40%, however as little as 2% increments can be applied.

A visual flap position indicator is provided for each model just above the selector. A test switch is provided to check the flap system.

PA31-350

Up to flap 25 can be extended at 162 knots or less.

Flap settings greater than 25 degrees are restricted the white

arc.

Flap 15 used in the approach config at 5nm from fix or airfield

will allow aircraft to slow to Vlo/Vle of 153

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Split Flap Warning System

The electronic system includes an amplifier, three rheostats, two power solenoids, and one motor. The amplifier provides a regulated voltage supply for the three external rheostats, circuit logic to analyze the system. The rheostats provide information to operate the system. Should an imbalance occur the sensing system will shutdown if the right flap rheostat does not agree with the left one within 5% at all times

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An Annunciator light will illuminate when an imbalance

occurs or a component of the amplifier fails

Caution: Serial numbers 31-7712001 through 78121129 were manufactured with a time delay relay instead of an electronic control system

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Chapter 4 Hydraulics/Landing Gear-Brakes

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Nose Landing Gear

The nose landing gear is

conventional with an air-oil oleo

type strut. It retracts aft into a

wheel well in the nose section of

the aircraft. Two doors while

retraction follow the nose gear

into the well. Rollers just above

the shimmy damper engage

slotted brackets on the gear doors

are operated as the gear retracts

and extends.

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A shimmy dampening device and steering linkage is provided so that the nose wheel is steerable while taxiing. While straightened during extension and retraction

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Nose Gear Assembly

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The main landing gear also incorporates a air-oil oleo type strut that retracts inboard into the wheel wells. The outer doors operate by mechanical linkage, and hydraulically actuated and remain open until the gear extend or retracts. The doors are opened hydraulically and during the extension cycle the gear handle when cycled in the up position, opens the doors. When the doors are fully opened

pressure continues to build in the system a [cracking pressure] of approximately 1,800 psi then directs hydraulic fluid to the gear piston type actuator, which retracts the landing gear. Upon the completion of this cycle the landing [outer] gear doors begin to close and the cycle is complete. See following pages for detailed schematic.

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The landing gear are held up by up lock hooks. The cycle is then considered to be static until a command restarts the cycle. Before landing a test can be conducted by lifting the landing gear handle and timing it’s return to neural. This indicates the systems operation.

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Main Landing Gear Assembly

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System Operation

When the selector is moved to either the [UP DOWN] position, it is locked in place by action of the handle release valve at the power pack., acting against the release mechanism detent. The handle will remain in this position until the hydraulic fluid pressure in the actuator reaches a present pressure. At this time the lever will return to [NEUTRAL].

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Movement of the selector handle, electrically actuates the door solenoid valve to its open position [should there be an electrical failure the solenoid is held closed by electrical power therefore will open without electrical power]. Then hydraulic fluid circulating through the open valve causes the doors to open. As they the gear priority valve remains closed as less pressure is required to operate the doors than is required to operate the gear. After the gear doors have completely opened, pressure continues to build enough to cause the gear priority valve to open. This permits the flow of fluid through the gear selector valve and the to the gear actuating cylinder. This causes the gear to either extend or retract, depending on the position of the gear selector handle.

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Annunciators

Located on the instrument panel, above and to the right of the gear selector control are one red and three green indicator lights. The red light will show an indication. When the gear is not located in either the up or down position. The green lights will show when each individual gear is down and locked.

The Red Gear Annunciator will illuminate when the gear is neither full up , or

full down, or the inboard doors are not completely closed. Pilots shall action appropriate checklist for resolving the gear issue.

Caution: Be sure that the gear [NOT LOCKED] light is out prior to exceeding the

maximum gear operating speed.

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Landing Gear Emergency Extension

Located between the pilot and copilot seat under the floor panel, is a hand pump uses for the emergency extension of the land in gear should a hydraulic failure occur.

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Check valves prevent the fluid from backing up through the engine driven pumps into the reservoir. In the event of a hydraulic fluid leak, a standpipe in the reservoir prevents the fluid level from dropping below the minimum quantity required for emergency gear extension by using the hand pump.

To operate the system lift of the cover by turning a small [D] type handle counter clockwise and lifting the access

cover labeled Emergency Gear Extension.

Place the gear selector in the down position, pull the pump handle out and pump the handle until the three green position lights illuminate. Approximately 50 full strokes is required to achieve full landing gear extension.

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Landing Gear Warning Horn

When power is reduced below approximately 10 to 12 inches of manifold pressure with the gear selector in the [UP] position, the horn will sound.

When the selector is in the [UP} position the horn will only sound when the throttles are retarded to that mentioned value, however, the horn will sound regardless of the landing gear being in the [UP] position when the flap lever is moved beyond approach flaps. This cannot be silenced until the excessive flap is reduced or the landing gear is down and locked.

An additional safety switch is installed on the left landing gear main. Should the landing gear handle be inadvertently placed in the [UP] position while the aircraft is on the ground, the warning horn will sound and the gear will not retract.

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Brake System The brakes are actuated by individual master cylinders mounted on the left set of rudder pedals. A reservoir, accessible through an access door located on the top of the nose section of the fuselage, supplies fluid to each master cylinder. The industry standard hydraulic fluid is MIL-H-5606. The fluid is routed through

lines and hoses to the master cylinders, then to the brake system.

The parking brake may be actuated by applying pressure to the top of the rudder pedals, then pulling out the [T] handle located below the pilots yoke. To release the parking brake apply toe pressure and at the same time push in the brake handle.

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Chapter 5 Fuel System

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Fuel Cells

Fuel is stored in two bladder type flexible cells in each wing.

The inboard cell hold 56 U.S. gallons each

The outboard cells hold 40 U.S. gallons each, providing a total of 192 gallons, of which

187.3 gallons are usable.

There are [Optional] nacelle fuel cells, each holds 27 U.S. gallons to a total of 246 U.S. gallons.

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Fuel Drains

Drains on the filters are at the base of the filters and are accessible through access doors in the lower portion aft of the wing leading edge.

Fuel drains for checking contamination are located at the rear inboard corner of the fuel cells, on the fuel on the fuel filters and the lowest point of the

crossfeed system. The quick drains for the cross feed is located on the left wing just forward of the main spar at the bottom of the wing. There are 6 drains in all.

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Fuel Pumps

Emergency fuel pump switches are located above the windscreen on the copilot side. These pumps are uses for takeoff and landings. Should an engine driven pump fail during takeoff, the emergency pump would provide fuel flow to the engine.

Note: Caution should be used when switching off the electric pump. Should the

engine quit, immediately turn on the pump and land as soon as practicable.

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Boost Pumps

Are electric, continuous duty, in-line type pumps and are located between the fuel filter and the electric fuel pump.

These pumps are provided to maintain fuel under pressure to the other fuel pumps, improving the altitude performance of the fuel system.

Each pump is controlled by a separate circuit breaker and are on when the master switch is turned on.

Fuel boost pump warning lights are located in the Annunciator panel and will illuminate should a pump fail.

Each are controlled by a sensor switch located above the firewall shutoff, forward of the cross feed line. Should the pressure drop below 10 psi the sensor will indicate a disruption and the light will illuminate.

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Fuel Selection and Cross feed

During normal operation, each engine is supplied with fuel from its own respective fuel system. Fuel controls on the right control fuel supply from the right tanks and left from the left tank. The cross feed is located in the center of the valve arrangement.

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Fuel Gauges

The gauges are located on the overhead panel between the pilot and copilot. The amp. Meter for the [LEFT and RIGHT] alternator is located in the center between the two fuel quality gauges on the PA31-350.

Note These fuel gauges are NOT accurate. Use time and cross check with fuel gauges. Do Not Rely On Them!

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The two fuel quality gauges are mounted in the overhead switch panel. As positioned each gauge represents its respective fuel tanks. [LEFT] and [RIGHT]. When the inboard tanks are selected the gauge reads the amount of that tank, when to outboard is selected it read that amount. When a fuel tank is selected, its corresponding micro switch is actuated, which completes the circuit between the fuel senders and its fuel quantity gauge.

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Fuel Flow and Fuel Pressure Gauges

A dual indicating fuel flow and fuel pressure gauges are located on the cockpit panel. The fuel flow is located above and to the right of the control pedestal, while the fuel pressure gauge is located to the right of the copilot control column. Both gauges gave two indicators, one for each engine. When the master switch is on and the fuel mixtures advanced, a pressure of approximately 10 psi should be indicated. Should the circuit breaker for the boost pump be pulled the gauge will read zero. This procedure is recommended to check the pumps.

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Fuel Warning Lights

Fuel flow warning lights illuminate when an impending fuel flow interruption is sensed. The lights are activated by a sensing prop mounted near each inboard fuel tank outlet. Should the flow be interrupted and power loss occur the warning light will illuminate. The warning light will illuminate for approximately 10 seconds and remain on if the cause is not corrected.

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Fuel Boost Pump Warning Lights

Warning lights for fuel boost pump failure are located on the Annunciator panel. The respective light illuminates when the boost pressure fall below 3 psi. Therefore, in a full power continuous climb from takeoff to high altitude under conditions of high ambient temperature, high climb rate, and extremely volatile fuel, the boost pump may not maintain sufficient pressure to the engine driven fuel pump. Should this condition occur and the boost pump inoperative light illuminates and remains on, the pilot should switch on the [EMERGENCY] fuel pump

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A warning light on the PA31-350 is located on the fuel control panel is incorporated in the firewall shut off system to indicate the one or both of the shut off valves are not fully open. This procedure is listed in the emergency Engine Securing Procedures.

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Fuel System Schematic

Placing the fuel mixture levers in “IDLE CUT-OFF” shuts off the fuel supply, so it should never be necessary to set the fuel selectors in “OFF” position unless the aircraft will not be used for an extended period of time.

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Fuel Additive

The Navajo requires 100 or 100 low lead fuel. Normal preflight fuel draining

procedures will generally remove most excess water from the fuel tanks However,

small amounts of water may remain in solution within the gasoline. This isn‟t

normally a problem unless you encounter high humidity conditions on the ground

followed by flight to high to high altitudes with low temperatures. On these rare

occurrences, it is permissible to use an anti-icing fuel additive.

Ethylene Glycol Monomethyl Ether (EGME), may be used. EGME is better known

by the commercial name “Prist”. The product absorbs dissolved water in the fuel

and also lowers the freezing temperature.

Caution: The EGME (Prist ) is a very aggressive chemical and should not exceed

0.15% of fuel volume, as it can destroy the fuel tank. It will even attack the

protective primer, sealants and O-rings in the fuel system, and should never come

into contact with the aircraft paint or finish. It must be thoroughly mixed with the

fuel.

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Chapter 6 Electrical System

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Left-Hand Circuit Breaker Panel

The Navajo electrical system is a 28-volt DC system with the negative lead of each power source grounded to the main aircraft structure.

Standby power is provided by a 24-volt, 17 Ah battery

Operating power is provided by two 28-volt, 70 ampere alternators.

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Battery The battery is rated at 24-volts and 17/25 ampere-hours. The 24-volts rating means that the battery can store 24-volts of power when it is fully charged.

The 25 ampere-hours mean that a brand new, fully charged battery can deliver one ampere of current for 25 hours without being recharged.

It should also be able to deliver 25 amperes for one hour or any combination of amperes and hours that add up to 25.

This rating system can help you determine how long your battery should last if both alternators fail.

One most remember, however, that this only works accurately with a brand new, fully charged battery. You most make allowances for older batteries and partially charged batteries.

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Battery

In operation the ALT Load gauge work very well.

To ensure that the Alternators can hold the load be sure to reference the volt meter.

As you load up the alternator during a run-up the voltage may drop. It should remain within the green arc.

Should the alternator not be able to maintain sufficient output the voltage on the system will drop to that of the battery or less.

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The condition of a lead acid battery will deteriorate at a fairly consistent rate. It is recommended that the battery water level be checked every 50 flight hours or every 30 days, whichever occurs first. However, since battery water consumption increases with warmer weather, these checks should be made more frequently in warm climates. Use only distilled water when servicing is required.

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Alternators Two 28 volt engine driven alternators supply power to all DC load buses and charge the battery.

The standard alternators are each capable of producing a continuous current of 70 amperes.

The alternator switch on the left side panel is a split rocker type and gives the pilot control over the field of the respective alternator.

The alternators are paralleled by the use of voltage regulators to control field voltage of both units.

Also incorporated in the system is an overvoltage relay. Its function is to open and remove field voltage to the unregulated alternator in the event of a failure of the voltage regulator.

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Electrical System Schematic

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Master Switch and Battery Control

The battery is connected directly to the hot battery bus. Therefore, the baggage compartments and exit lights are powered regardless of the battery switch position. A faulty switch or if the lights were left on could result in the battery being depleted.

1. Master Switch

2. Alternator Switches

3. Emergency Avionics Switch

When the master switch is selected to the on position, a relay is energized and power is provided to the battery bus. Should there be no indication of electrical power to the system, the battery must be removed and re-charged. A voltage meter is located on the pilots panel and should be checked to ensure sufficient voltage is available prior to engine start. Low voltage may damage the starters due to overheating during initial start.

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Engine Starter Switch

When the spring loaded type rocker ENGINE STARTER switch is selected, current closes the starter relay. This switch is held until engine start. It should be noted [Starters should not be energized for more then 30 seconds of continuous cranking]. Allow sufficient time to cool between starting attempts

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The alternator switch should be on during engine start. Battery current is

allowed to flow through the alternator switch, overvoltage relay, and voltage

regulator to excite the alternator field, which the produces power.

The alternator is producing 28 volts while the battery is supplying 24 volts.

Therefore, current flow reverses permitting the system to be powered, and

the battery charged.

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External Power Receptacle

To conserve battery power an external power source can be used. This allows for assisted starts during cold weather, or when testing electrical equipment on the ground. The receptacle is located on the lower left side of the nose of the aircraft. The master switch should be OFF before inserting the external power plug into the receptacle. Only a source that can provide 24 volts and 300 amperes is approved. Although a relay protects the system in case of overvoltage, damage can result to the wiring between the receptacle and the relay

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On the PA31-350, an ammeter is located between the fuel quantity gauges.. Should the ammeter indicate to the left of center, the battery is being discharged.

The battery is being charged when the opposite indication exits. During alternator failure or single engine operation, this feature can be used to determine the electrical load and how much the load should be reduced by the pilot.

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The output of each alternator can be checked by depressing the Press-To-Test

buttons located on either side of the ammeter. When each button is pushed

separately, alternator output is indicated. When flying in heavy rain, the electrical

load on the right alternator must be reduced to 40 amperes or less to ensure

against alternator belt slippage.

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Lighting System

External lighting control switches are located on the overhead switch panel. They

control wing and tail navigation lights, anti-collision lights, taxi lights and optional

wing deice lights.

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The taxi an landing lights have a two position ON and OFF switch.

Should the taxi light be in the on position when the landing gear is

retracted, the light will not illuminate. However, it will automatically

illuminate when the landing gear is extended as it is wires through the

gear down micro switch. Other switches are self explanatory as to

their use.

It recommended that anti-collision lights not be operated on the ramp

in close proximity to ground personal. During flight, operating the anti-

collision lights through, fog or haze, could cause spaceial

disorientation, or vertigo.

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The nose baggage compartment light rear dome light and exit light are hot wired and controlled by individual switches. In addition an optional timer is available to operate the cabin lights for 30 second upon opening the main cabin door.

A cockpit dome light is located in the overhead panel. The push button switch is located just forward of the lens.

Reading lights are available for each passenger seat. Switches are incorporated in each overhead light assembly.

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Annunciators

Left Bank

Right Bank

The annunciator panel extends across the upper center of the instrument panel. The standard panel will indicate the status of the pneumatic system, fuel boost pumps low fuel flow, flap condition, alternator, operation, combustion heater temperature, and cabin and baggage door security.

Key:

Red…………….Requires immediate attention, hazardous condition exits

Amber…………Requires attention, possible dangerous condition exits

Green or White…………..Safe or normal configuration, routine action

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Chapter 7 Flight Instruments

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Basic Flight Instruments

The pilot and copilot instrument son the Navajo vary from aircraft to aircraft. However, the basic flight instruments utilizing the pilot/static system are, airspeed indicator, altimeter, and vertical speed indicator.

Typical Instrument Panel

In addition there are three gyroscopic instruments, attitude indicator, heading indicator, and turn coordinator. These instruments are usually driven by instrument air. In certain combinations, air or electric attitude and are optional. The copilot instruments are usually air driven, with an electrically driven turn coordinator.

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Pilot/Static System An electrically heated pitot tube is mounted on he left side on the nose section. This supplies ram air to the pilot airspeed indicator. An additional heated pitot tube is located on the right side, when a copilot airspeed indicator is provided. Electrical switches located on the overhead panel controls the pilot tubes. In some aircraft one switch is provided for both heaters. However an additional copilot switch is provided as an option. The pilot/static system should be checked along with all other anti-ice or de-ice equipment prior to flight in known icing conditions.

Pilot and Copilot Pitot Tubes Instrument Panel

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Static Ports

On the Navajo, two static ports are are provided, one on each side of the fuselage forward of the horizontal stabilizer.

Two static pressure ports on each side is an optional system if installed. The system is

plumbed to the airspeed indicator, vertical speed indicator, and altimeter. System pressure errors during uncoordinated flight are eliminated by placing interconnected static ports on each side of the aircraft.

alternate static source. To actuate, push the lever up and to the left to lock the valve in the open position.

Static Port

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Airspeed Indicator

Altimeter

The airspeed indicator measures the difference between ram air from the pilot tube and static pressure. Calibration is in both knots and statute miles per hour. A small knob for calculating true airspeed is located on the lower right hand side of the instrument.

The altimeter is a pressure instrument that measures the change of static pressure by means of an indicator attached to an aneroid and is corrected to seal level. A manually adjusted knob calibrates the instrument and is locate on the lower left corner of the altimeter.

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Vertical Speed Indicator

A vertical speed indicator is connected to the static pressure line through a calibrated leak. This instrument compensates for changes in pressure and is considered a pressure differential instrument. Pressures is established between a diaphragm and trapped static pressure within the case. It is calibrated to read changes in altitude in hundreds of feet per minute.

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Gyro System

Right Engine

The heading and attitude indicator are powered by air pressure from the pneumatic system.

The turn coordinator is electrically powered as serves as a standby system in he event of dual pneumatic system failure.

This system is divided into two independent pressure and vacuum supply.

The system utilizes a common manifold, check valve and pressure gauge.

Engine driven pumps supply the pneumatic system through an inlet, inlet filter, regulator and inline filter.

Recessed inlets, just aft of the fire wall on the bottom-outboard side of each, engine nacelle, extracts a constant supply of outside of each, engine nacelle, extracts a constant supply of outside air, which is passed through inlet filters and directed to each engine driven pneumatic pump.

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Gyro pressure gauge

Air pressure for each pump is routed to its respective pressure regulator. Excess air supplied by the pumps is expelled from the system through its regulator. Pressure from each of the pumps supplies a manifold, with an inline check valve which in the event of a pump failure the other pump can maintain sufficient pressure for the gyro and deice equipment up to the single engine service ceiling. Air pressure supplied by the system is utilized to operate the attitude gyro and the directional gyro and exhausted forward of the instrument panel through the pressure bulkhead.

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A gyro pressure gauge, is mounted in the right segment of the instrument panel, which indicates system pressure in inches of mercury. A graduated green arc on the face of the gauge, indicates pressure readings within normal operating limits.

Annunciator panel

Annunciator lights, mounted in the annunciator panel provide a visual warning to the pilot should either the right or left pneumatic source is inoperative. The lights are labeled [R PNEU INOP] AND [L PNEU INOP].

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Gyro Pressure System

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Horizontal Situation Indicator

This motion is shown on the instrument which has a 360° direct reading dial.. The dial, when set to agree with the magnetic compass, provides a positive indication free from swing and turning error. An HSI is a popular instrument and operates by electrical power and is slaved to a compass, therefore setting is unnecessary, however a visual check against the liquid compass should be performed periodically.

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Attitude Indicator

The attitude indicator is operated by the same principal as the heading indicator. It provides a constant visual reference to the attitude of the airplane relative to pitch an roll axis. A bar across the face of the indicator represents the horizon.

A miniature adjustable airplane is mounted to the case, and aligning the miniature airplane to the horizon bar simulates the alignment of the airplane to the actual horizon. Any deviation simulates the deviation of the airplane from the true horizon. The attitude gyro is marked for different degrees of bank.

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Turn coordinator

Operation of the turn coordinator can be checked by initiating a standard rate turn and cross checking the turn rate with the heading indicator.

This instrument is actually two instruments in one. The turn portion is electrically driven gyro attached to a

pointer or miniature airplane. The lower portion, [Slip Indicator] is a curved, liquid filled glass tube in which an inclinometer ball changes position according to gravitational and centrifugal forces acting on the aircraft. Inaccuracy in the slip indicator can only result if the instrument is not mounted level in the instrument panel. With the aircraft on level ground, the ball should be centered in the tube.

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Chapter 8 Environmental Systems

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Environmental Systems an Location

1. Air Conditioner Control 2. Heater Switch

3. Heater Temp. Control 4. Defrost Control

5. Heater Air Inlet Control 6. Outside Air Control

PA-31/350

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Air Distribution PA-31

Cabin and cockpit air controls are located on the right lower portion of the copilot instrument panel. The controls regulate the amount of fresh air entering the cabin during ground and flight operations.

The combustion heater circulating fan can draw air through the heater inlet, when the heater is not in use, or is used during heater operations. Selecting the

FAN position of the heater two position switch allows heating or fan. During heater operations the combustion fan is activated while the heater is in operation.

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Air Controls Air flows through air ducts along each side of the fuselage to the fresh air outlets. Cabin air is re-circulated through aisle grills located on each side of the compartment.

An Air Control fan switch is located to the right side of the overhead panel. This is a three position switch, which controls the fan, and air conditioner, and centered to the off position

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The heating system incorporates a Janitrol heater with windshield defroster,

forward and aft heating controls. Air to the heater is located on the

underside of the nose section. Air passes through the heater and

distributed to the various heater outlets selected with the Controls. It should

be noted that before the heater is operated, the heater air inlet be opened.

See Heating System Schematic next page.

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Heating System Schematic

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Heating System Schematic

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Heating System

The heater is protected from overheating by a heat limit switch. If the heater temperature reaches a predetermined setting, the limit switch opens and the heater becomes inoperative. This is indicated by the illumination of the [HTR FAIL] warning light in the Annunciator panel..

The heat limit switch will automatically reset when the heater has cooled. By

depressing the START/RESET switch momentarily, the heater can be restarted.

Heater restart is indicated by the [HTR FAIL] warning light extinguishing after

the START/RESET switch is released.

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Caution: Prior to restarting the heater, ensure the following precautions are followed:

a. The heater air inlet lever must be full open prior to and during heater operation.

b. Open all heater outlets to the full open position.

c. The TEMP HEAT lever should only be half open.

d. After heater has restarted, wait approximately 5 minutes and then, if desired, the

temperature lever may be moved to a higher selection.

e. Immediately shut off heater if Annunciator comes on again after attempting to

restart. Do not operate heater again until it has been serviced by a qualified

repair station.

Warning: Operating the aircraft with a defective heater may be a serious fire hazard.

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Chapter 9 Anti-Ice/Deice

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Anti-Ice System A two position ON/OFF switch is located

on the overhead switch panel controlling

a D.C. electric heating element in the

pitot tube. If a dual pitot system is

installed, both systems have heaters.

The stall warning transmitter, located on

the right wing, is also electrically heated

and controlled by this switch.

Pitot Heat

This transmitter is heated when the pitot heat is on and the weight is off the

landing gear.

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1. Windshield Heat Switch

2. Pitot Heat Switches

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Windshield Heat The optional electrically heated windshield, used to prevent and/or remove icing and fogging, is controlled by a WINDSHIELD HEAT switch mounted on the overhead switch panel. With the engine running, a preflight check can be made by activating the control switch. The windshield is operating properly if it feels warm to the touch.

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The heated windshield should keep the surface clear of fog, however,

when encountering icing the heater and defroster should be turned

“ON”. Windshield heat must be turned “ON” prior to entering icing

conditions.

Caution: Ground operation should be kept to a minimum to prevent

overheating of the windshield. Distorted vision or small bubbles in the

plastic of the windshield may indicate an overheat condition.

The exterior surface of the windshield has a Nesa coating to prevent

static discharge. Use care when cleaning.

Note: If windshield wipers are installed, do not operate them

when ice is accumulated on either the wiper or windshield.

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Deice Systems

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The surface deice system removes ice accumulations on the leading edges of the wings, horizontal and vertical stabilizers. The deicer is essentially a fabric reinforced rubber sheet containing built-in inflation tubes. The type used in this installation have spanwise inflation tubes. Deicers are attached by means of cement to the leading edges of the surfaces being protected.

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The electrically controlled and

pneumatically operated system

consists of engine-driven

vacuum pumps, solenoid valves,

pressure switch electrical timer,

annunciator lights, and inflatable

rubber deice boots.

Through the engine driven pneumatic pumps, the system will normally apply

vacuum to the deicer boots at all times, except when the boots are being inflated.

Deicer inflation is effected by the deicer system control switch. This is a

momentary on type switch labeled SURFACE DEICE on the overhead switch

panel.

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Actuation of the momentary switch

triggers a system cycle timer,

This in turn shifts the two stage

regulators to high pressure (18 psi)

„A‟ system solenoid valve opens

sending air to the wing boots, and cuts

off air to the copilot‟s gyros (when

installed).

After six seconds, the „A‟ system

solenoid is closed and the „B‟ system

solenoid is opened to send air to the

tail boots for six second.

At the completion of the tail cycle, the

„B‟ system solenoid is opened to send

air to the tail boots for six second..

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At the completion of the tail cycle,

the „B‟ system solenoid closes, the

two stage regulators return to low

pressure (gyro pressure) and the

copilot‟s air supply resumes.

When the inflation cycle is

completed, the deicer solenoid

valves permit overboard exhaustion

of the pressurized boots. Suction is

then reapplied to the deicer boots to

hold them close to the airfoil surface.

The inflation cycle occurs only once

each time the momentary on switch

is activated

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Two blue indicator lights with press-to-test and dimming features, illuminate

when each surface deicer boot system inflates to a predetermined pressure.

Illumination of the indicator light is controlled by a pressure sensitive switch

connected to the deicer pressure lines (one in the „A‟ system, and one in the

„B‟ system).

Note: To insure good ice shedding, the boots should be clean and free of

any oils or dirt and in good condition. No special coating is required.

The pneumatic deicing system should be checked at least every 100 hours.

This check can be done on the ground. A visual inspection should be

preformed to determine the condition of the deicer boots, and any areas in

need of repair should be taken care of before continuing with the operational

check of the system

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Before takeoff, the deice boots should be cycled and visually checked

for proper operation, including the annunciator. With one engine

operating, activate the deicing system switch. The pressure will

fluctuate as the tubes inflate and deflate. Check the pneumatic

pressure gauge. If pressure is satisfactory, observe the operation of

the deicers carefully for evidence of malfunctioning. Look for tubes

which leak or fail to inflate and deflate properly. Repeat the procedure

for the other engine.

Caution: In order for the deice boots to properly crack and remove the

ice, an accumulation of at least 1/4” to 1/2” of ice must be present.

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Propeller Deice

Although this is a deice

system, it is generally used

in an anti-ice capacity to

prevent large chunks of ice

from damaging the side of

the aircraft.

Electro thermal propeller deicer pads are bonded to the leading edges of the

propeller blades. Each deicer pad has two separate heaters, one for the

outboard and one for the inboard half.

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The system is controlled by an on-off type PROP DE-ICE switch located in the overhead switch panel and the circuit is protected through a circuit breaker in the circuit breaker control panel. When the switch is actuated, power is supplied to the system timer. The propeller deice ammeter is connected in series between the switch and the timer to monitor the current through the propeller deicing system. With the propeller deicing system the ammeter needle should be within the green arc on the face of the ammeter for a normal reading.

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Power from the timer is cycled to brush assemblies which distribute power to modified starter ring gears incorporating slip rings. The current is then supplied through the slip rings directly to the electro thermal propeller deicer pads.

Deicing is accomplished by heating the outboard and then the inboard half of the deicer pads in a sequence controlled by the timer. The heating sequence of the deicer pads is according to the following cycle:

1. Outboard halves of the propeller deicer pads on the right engine (30 seconds)

2. Inboard halves of the propeller deicer pads on the right engine (30 second)

3. Outboard halves of the propeller deicer pads on the left engine (30 second)

4. Inboard halves of the propeller deicer pads on the left engine (30 second)

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When the system is turned on, heating may begin on any one of the above

steps, depending upon the positioning of the timer switch when the system was

turned off from previous use. Once begun, cycling will proceed in the above

sequence and will continue until the system is turned off.

A preflight check of the propeller deicers can be performed by turning the PRP

DE-ICE switch to “ON” and feeling the deicer pads for proper heating sequence.

The deicer pads should be warm to the touch. A less vigorous test can be made

by turning the switch “ON” and “OFF” four times and noting that the ammeter

needle goes to the green arc each time.

Caution: When conducting the above described ground test, do not operate

system longer than two complete cycles.

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If the ammeter reading is less than in the green arc this indicates that one

or more of the pads is not functioning. If propeller deice is used under this

condition the pilot can expect an uneven build up of ice with consequent

undesirable vibration.

Ice shields can be installed on both sides of the fuselage nose adjacent

to the propeller. These are installed to prevent excessive damage due to

slivers of ice shed from the propeller impacting on the skin.

Note: Propeller imbalance may be relieved by varying the RPM;

increase RPM briefly and return to desired setting, repeating if necessary.

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Course Completion

SAFETY IS NO ACCIDENT