Enstrom 480B TM

5/23/2018 Enstrom 480B TM

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TH-28/480/480B Training Manual5/29/2008 2007 Edition 2

For Training Purposes Only

Introduction 3

Aircraft Description 4

Construction details 7

Aircraft Systems 17

Electrical 17Caution and warning 23Instruments 29Rotor Systems 37Fuel System 42Power Train 45Flight Controls 50

Power Plant 57

Operation Procedures 63

Aircraft Servicing 63Performance Data 67Engine 74Emergency Procedures 84Weight & Balance 98

Pilot Notes 105

Table of Contents


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INTRODUCTION

Enstrom TH-28/480 Series Helicopter Pilot Training CourseObjectives

The purpose of this course is to prepare an experienced helicopter pilot for a smoothtransition into the Enstrom Turbine powered helicopters.

This course includes descriptions and theory of operation for the systems, and thelocation of the system components.

The course also includes the description of the pilot pre-flight procedures and the pilotsare expected to perform these pre-flight inspections.


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AIRCRAFT DESCRIPTIONThe TH-28/480 helicopter is a 3 bladed, single engine helicopter manufactured by theEnstrom Helicopter Corporation and certificated by the FAA under FAR Part 27.

Turning Radius

The turning radius is about 23 feet when pivoted on the wheels about the mast.

Principal Dimensions


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Characteristics

Helicopter Description

The Enstrom 480B helicopter is a single-engine; turbine-powered helicopter certified forday and night VFR flight, that can be equipped for IFR flight. The 480B was developedfor light commercial, municipal, and military uses and was certified to FAR 27 standardsin February 2001. Its predecessor, the 480, was designed between 1988 and 1993,and was certified to FAR 27 standards in 1994.

It is a relatively quiet helicopter that was certified to meet FAR 3 Appendix J noise

limits. The main and tail rotors are relatively slow turning which contribute to the lownoise signature.


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The Enstrom 480B features a three-bladed, fully articulated main rotor system whichhas over 3,000,000 flight hours and which has never had a catastrophic failure orthrown a blade. The tail rotor is two bladed and completely unblocked for exceptionaleffectiveness. Due to the high inertia rotor design, the helicopter possessesoutstanding auto-rotational capabilities. There have been no fatalities from any accidentinvolving an Enstrom 480 model of helicopter. ( As of the date of this manual revision)

In the event of a mishap, the 480 is extremely crashworthy. The basic landing gearand airframe feature an integrated energy absorbing system.

The 480 comes standard with high skid landing gear which consists of aluminum skidtubes and nitrogen air-oleo struts to cushion ground contact. Replaceable hardenedsteel skid shoes are installed on each skid to resist wear on hard surfaces.

In addition to being a versatile and crashworthy helicopter, the 480 is designed to beprocured and operated for minimum costs. The basic modular design is simple andinexpensive to manufacture. The helicopter does not require hydraulic boost, electricboost pumps, or a stability augmentation system. The entire control system consists ofmechanical linkages. The avionics package is designed for easy installation andaccessibility and the 480 is configured with five hinged doors and five removable panelsfor maintenance accessibility.

The limited number of fatigue critical parts, the long overhaul intervals, and the lowmaintenance hour/flight hour ratio resulting from high reliability and easy maintenancecombine to yield low operating and support costs.


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CONSTRUCTION DETAILS

Fuselage

The fuselage is the forward section of the airframe extending from the nose to theforward end of the tailcone. The primary fuselage structure consists of the keelassembly (two longitudinal beams with transverse bulkheads) which is attached to awelded steel tubular truss structure called the pylon. All of the major components ofthe aircraft are attached to the pylon. The keel assembly is the main supportingstructure for the cabin and forward landing gear cross tube. The pylon forms thesupporting structure for the cabin, fuel cells, transmission, engine, aft landing gearcross tube, and the tailcone. The cabin shell is of composite construction withreinforcing where necessary to add structural stiffness.


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Cabin Floor and Backwall

Keel Attach Structure


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Tailcone

The tailcone is bolted to the aft end of the pylon. It is a tapered, semi-monocoquestructure comprised of skins, bulkheads, longerons, and stringers. The tailcone supportsthe tail rotor, tail rotor transmission, horizontal and vertical stabilizers, and the tail rotorguard. It houses the tail rotor drive shaft and some electronic equipment.


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Landing Gear

Main Landing Gear: The main landing gear consists of two tubular aluminum skidsattached to the airframe by means of the forward and aft cross tubes through four air-oil oleo struts. The struts cushion ground contact during landing.

Drag struts give the gear stability and strength and prevent fore and aft movementduring ground contact maneuvers. Due to their design, the drag struts will sustainlandings with significant forward movement of the helicopter; however, landing withrearward movement may overload the structure and cause its collapse.

Replaceable hardened steel skid shoes are installed on each skid to resist skid wear on

hard surfaces.


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Ground Handling Wheels

Each landing gear skid tube has provisions for installing ground handling wheelassemblies. Each skid has two lugs that the wheel assemblies are installed on. Eachassembly has a manually operated over-centering device to lift the skids clear of theground.

The ground handling wheels must be removed before flight.

When the helicopter is placed up on the wheels for ground movement, care must betaken to support the tail hoop to prevent inadvertent contact with the ground. It isadvisable to place the tail rotor in the horizontal position to prevent damage to the tailrotor blades in the event that the helicopter tips on to the tail as the wheels areinstalled close to the helicopter center of gravity.


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Crew Compartment

The crew compartment contains the pilot and copilot/passenger seating, a complete setof dual flight controls, a lower radio console, and an instrument panel all enclosed bythe composite cabin.


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Cabin Doors

The two cabin doors are composite reinforced structure with transparent plexiglasswindows in the upper section. Ventilation is supplied by sliding vent windows to drawfresh air into the cabin. Positive retention door latches are used to secure the doors.

Cabin ventilation is provided by pop-out vents, sliding vent windows, a ram airventilation system or a bleed air heating system depending on the optional equipmentinstalled on the aircraft. There is also a ceiling mounted air circulation fan installed in

many 480Bs.

A Freon air-conditioning system is available for the 480Bs as optional equipment.


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Seats

The pilot and copilot/passenger seats are adjustable for fore and aft positioning and areeasily removed from the aircraft to facilitate maintenance on the seats or the cockpitarea. Both seats use a composite bucket mounted on a pedestal assembly. A four-point restraint system with a release buckle, adjusters in the lap and shoulder belts, andan inertia reel mounted on the back of the seat pedestal are an integral part of theseat.

Passenger Seats: The passenger seats are mounted to the pylon assembly throughthe cockpit bulkhead and fold up to the stowed position when not in use. The

passenger seats use a three-point automotive style single shoulder strap with an inertialreel.


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Inertia Reel Shoulder Harness

An inertia reel and shoulder harness is incorporated in all seats. There is noindependent control to manually lock the harness. With the shoulder straps properlyadjusted, the reel strap will extend to allow the occupant to lean forward; however, thereel automatically locks when the helicopter encounters an impact force of 2 to 3 "G"deceleration. To release the lock, it is necessary to lean back slightly to release tensionon the lock.

If the Pilot and co-pilot shoulder straps are adjusted too loosely, the webbing splice will

catch in the slot in the top of the seat back. This has the effect of the shoulder harnessinertial reel locking up and preventing the crew member from leaning forward. For thisreason the shoulder straps must be adjusted so that the y splice is close to the seatoccupants neck.


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Engine Assembly

The TH-28/480 is equipped with an Allison designed, Rolls Royce built, 250-C20W free-turbine, turboshaft engine rated at 420 SHP but derated in this installation to 285 SHPfor a five minute take-off rating, and 256 SHP for maximum continuous operation.

In the 480B, the engine is rated at 305 SHP for a five minute take-off rating and at 277SHP continuous.

Refer to the Rolls Royce 250-C20 Operation and Maintenance Manual) for a completedescription of the engine assembly and its sub-components.


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AIRCRAFT SYSTEMS

ELECTRICAL SYSTEMS

Description - Starter/Generator Systems

Rolls Royce helicopter engines are cycle limited (start limited). Overhaul of the hotsection, (mini-turbine) is required at 3000 starts or 1750 hours, which ever occurs first.Because start counters can be inaccurate, it is recommended that the pilot keep careful

records of starts.

Most commercial operators keep a trip log which records: Date, hour meter, totalnumber of starts, pilot name and the purpose of the flight. For the pilot / owner, it isworth considering removing the start counter and substituting a trip log in its place.


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Battery shelf, Master and Starter Relays and GPU Plug

The starter system is used to start the aircraft powerplant. Using either battery poweror external power and with the battery switch turned ON to supply electrical power tothe main electrical terminal strip, the starter system is engaged by pressing the startswitch located on the pilot's collective stick control head. When the start switch isengaged, the start relay coil is energized and electrical power is supplied to the starterside of the starter/generator and electrical power is supplied to the ignition exciter andthe start counter.

The starter ignition lock is key operated and must be ON for power to be applied to thestarter-generator, ignition exciter, and the start counter. The starter/generator cannotbe engaged with the starter ignition lock in the OFF position. However, once thehelicopter is running, the key can be removed if necessary to access a door lock.

DO NOT FLY THE HELICOPTER WITH THE KEY IN THE OFF POSITION ORWITH THE KEY REMOVED.


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The generator system is used to supply 28 Vdc electrical power to the main electricalterminal strip and to recharge the battery after a battery powered start. The generatorsystem is controlled by the generator control unit (GCU). The GCU performs the

following functions: voltage regulation, overvoltage protection, reverse currentprotection, over current protection, generator failure indication for the caution panel,and generator field excitation.

The voltage regulator portion of the GCU can be adjusted. When the generator switchis placed in the on position, the GCU connects the starter-generator to the mainelectrical terminal strip via the generator relay. When the starter-generator is on-line,the dual volt/ammeter monitors generator current output via the generator shunt.

Helicopters built after November 2007 incorporate a cooling duct that directs cooling air

from the transmission inlet scoop to the generator.


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GCU plug

The ground power plug is installed on the inside forward bulkhead of theright side engine compartment just in front of the battery. Beginning in2007 there is an optional access door available.

When the APU is being used toassist starting the 480 helicopters, the battery switch must be in the onposition. It is important for the pilot to instruct the ground crew not todisconnect the APU power untill after the start is completed and the pilotgives the disconnect signal.

If the auxiliary power is being used to start an aircraft with a low battery,serious engine damage can result if there is an interruption of electricalsupply.


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Description - Generator Control Unit (GCU)

The GCU, located on the bottom right hand side of the oil cooler/blower shelf, or underthe cabin floor after S/N 5043, performs the following functions: voltage regulation,overvoltage protection, reverse current protection, over-current protection, generatorfailure indication for the caution panel, and generator field excitation.


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At flight idle RPM and above, the voltage regulator portion of the GCU maintains thecorrect generator output voltage by varying the generator field current. If thegenerator voltage exceeds 32.0 Vdc .5 Vdc, the internal overvoltage sensor will causecurrent to flow in the trip coil of the generator switch and trip the switch to the OFFposition. This removes the current from the generator field and power from thegenerator relay-actuating coil, disconnecting the starter/generator from the mainelectrical terminal strip.

The reverse current portion of the GCU de-energizes the generator relay when thegenerator output voltage falls below the battery voltage.

The over-current protection circuitry will cause current to flow in the trip coil of thegenerator switch when the generator maximum output current rating is continuouslyexceeded for 10 seconds 2 seconds. This trips the generator switch to the offposition removing the current form the form the generator field and the power from thegenerator relay-actuating coil. The circuitry in the GCU will illuminate the generatorcaution light (DC GEN) in the caution panel any time the generator voltage is less thanthe battery voltage, the generator switch is OFF, or the generator is not connected tothe main electrical terminal strip. The GCU will also flash the generator field circuitry ifrequired.

N1-N2-NR-TOT

The N1-N2-NR-TOT switch installed in the top right section of the instrument panel willconnect the N1, N2/NR and TOT gauges directly to the battery in the event of acomplete electrical power failure in the helicopter.


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CAUTION AND WARNING SYSTEMS

Description - Caution and Warning Systems

The caution system consists of a caution panel, two annunciator/switches (one on the480), a test/dim switch, and 14 input circuits. The warning system consists of 3individual warning lights and their associated input circuits.

The caution system is used to provide a visual indication that a fault condition hasoccurred. The caution panel, located in the instrument panel, has 15 individual wordedsegments which when illuminated identify specific fault conditions. When a faultoccurs, the associated segment on the caution panel illuminates and flashes at a 2 Hzrate, and the MASTER CAUTION annunciator/switches, located on the left and right sideof the instrument panel, will illuminate and flash at the same rate.

When the fault is acknowledged by pressing the MASTER CAUTION annunciator/switch,the MASTER CAUTION light will extinguish and the fault on the caution panel will resetto a steady (on) condition. As each fault condition occurs, it is indicated by the samesequence of events as described above. Only a new fault will flash until it is

acknowledged. The MASTER CAUTION annunciator/switches will only be illuminated byfaults associated with the caution panel; the warning system will not activate theMASTER CAUTION annunciator/switches.


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SEGMENT COLOR DESCRIPTION OF FAULT

ENG CHIP AMBEREngine scavenge oil has ferrous metalfragments

MAIN XMSNCHIP

AMBERMain transmission chip detector hasdetected ferrous metal fragments

TAIL CHIP AMBER

Tail rotor gearbox chip detector has

detected ferrous metal fragments

ENG OIL TEMP AMBEREngine oil temperature is above 107degrees C

MAIN XMSNHOT

AMBERMain transmission oil temperature is above107 degrees C

DRIVE BRGHOT

AMBEREither the fwd or aft lower pulley bearingsare above 120 degrees C

ENG OILPRESS

AMBEREngine N1RPM is above 78.5% and engineoil pressure is below 90 psi

ENG INLETAIR

AMBEREngine inlet swirl tube particle separatorpartially blocked

A/F Filter AMBER Airframe fuel filter bypass is impending

DC GEN AMBER DC Generator failure

FUEL FILTER AMBER Fuel filter bypass is impending

FUEL LOWAMBER Less than 5 gallons remaining

ENG DEICE GREEN Engine anti-ice has been activated

SPARE AMBER Unused segment


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Note: An optional external fuel filter can be installed on the TH-28 (S/N, 3007 andsubsequent) and the 480 (S/N, 5003 and subsequent). If the external filter is installed,

the SPARE segment on the caution panel will be connected to the impending bypassswitch incorporated in the filter assembly. The segment for the impending bypass willbe labeled A/F FILTER.

Note: The BATT HOT and BATT TEMP segments of the caution panel will only befunctional if an optional Nicad battery is installed.

Note: The ENG DEICE segment of the caution panel will not flash or cause theMASTER CAUTION annunciator/switch to illuminate or flash. The SPARE segment of

the caution panel will not cause the MASTER CAUTION annunciator/switch to illuminateor flash.

Functional Test: Caution Panel

Place the caution panel test/dim switch in the TEST position. Check that the MASTERCAUTION annunciator/switch is illuminated and flashing and that all the caution panelsegment lights are illuminated and with the exception of the ENG OIL PRESS, DC GEN,ENG DEICE, and possible the FUEL LOW, all the segments are flashing. Release the

switch and reset the MASTER CAUTION annunciator/switches.

The ENG CHIP, MAIN XMSN CHIP, and TAIL CHIP segments should only be illuminatedfor approximately 5 seconds and then extinguish due to programmed continuity sensorsin each detector circuit.

Reset the MASTER CAUTION annunciator/switches by pressing in on theannunciator/switch. Check that the MASTER CAUTION annunciator/switches extinguishand the illuminated caution panel segments are in a steady bright condition.


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Warning Systems

The warning system consists of three independent red warning lights located at the topof the instrument panel. When each light is activated, it comes on steady and fullbright with no dimming capability. These lights are for conditions that requireimmediate action.

LIGHTS COLOR DESCRIPTION OF FAULT

ROTOR RPM RED Main rotor RPM below 334 RPM

ENGINE OUT RED Engine N1below 58%

FIRE RED The fire detection system has detected eithera fire or an extreme overheat condition ineither the upper or lower enginecompartment.

ROTOR

RPM

ENGINEOUT

FIRE


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Functional Tests - Warning System

Turn on the master switch:

Check that the ROTOR RPM and ENGINE OUT warning lights are illuminated and theassociated audio horns are not activated.

Place the place the caution panel test/dim switch in the TEST position and check thatthe FIRE warning light illuminates. Release the switch.

Release the collective friction and raise the collective controls off of the down stop.

Check that the low rotor and engine out audio horns activate.

Run up the aircraft I/A/W the operator's manual. Check that the ENGINE OUT warninglight extinguishes when the N1passes through 58%.

Increase the power to bring the NR up to 334 rpm and check that the ROTOR RPMwarning light extinguishes at 334 1 rpm.

Place the caution panel test/dim switch in the TEST position and check that the ROTORRPM, ENGINE OUT, and FIRE warning lights illuminate. Release the switch.

Check the operation of the ROTOR RPM and ENGINE OUT warning lights during theshutdown procedure


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Indicator Caution Activation

TORQUE NO Exceeds 72 PSI for 1 second

N1 TACH NO Exceeds 105%

TOT NO

Exceeds 10 seconds between 810 927Cduring start.

Exceeds 810C for 5 seconds during normaloperations.

DC VOLTS NO Exceeds 30 Vdc for 5 seconds

DC AMPS NO Exceeds 110 amps for 5 seconds

ENG OIL TEMP YES* Exceeds 107C for 5 seconds

ENG OIL PRESS YESLess than 50 PSI for 5 secondsExceeds 130 PSI for 5 seconds

XMSN OIL TEMP YES Exceeds 107C for 5 seconds

FUEL QTY YES** Less than 35 lbs for 5 seconds

* The caution panel monitors both N1 speed and engine oil pressure for controlling the

ENG OIL PRESS segment. The indicator caution activation only uses engine oilpressure.

** The indicator panel is independent of the FUEL LOWsegment in the caution panel.


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INSTRUMENTS

Dual Tach

The N2 (power turbine) RPM is generated by a tach generator mounted of the left sideof the engine accessory gearbox and the Rotor RPM is generated by a magnetic sensorinstalled in the forward section of the MRGB.

The rotor and power turbine tachometer (dual tach) is a digital system powered by theaircraft 28-volt electrical system.


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Rotor Limitations

Minimum Transient Rotor Speed

The minimum allowable transient rotor speed following engine failure or sudden powerreduction for practice forced landing is 300 RPM. This is a transient limit and positivecorrective action (lowering the collective) must be taken immediately by the pilot toregain at least 334 RPM (minimum power off rotor RPM).

ROTOR

385 RPM Red Radial Max Power Off

334-385 RPM Green Arc Continuous Operation(Including Autorotation)

334 RPM Red Radial Minimum Power OFF

POWER TURBINE

113% RPM Red Arrowhead 15 Second Max Transient N2See Flight Manual Page 1-16.2

103% RPM Red Radial Maximum N2 Continuous

101-103% RPM Green Arc Normal Operating Range

101% RPM Red Radial Minimum N2 Continuous


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Early 480B Beginning 2007

Engine Instruments

Description

Helicopters built during 2005, 20006 and 2007 are fitted with instrument gaugessupplied by Horizon/Ultra which incorporate a diagnostic check which is performedwhen power is first switched on. The indicator light will start out red, and the needlewill swing to the full right position before settling at the correct reading. When thediagnosis program is completed, the indicator lamp will also indicate correctly, red orgreen, corresponding to the gauge reading.

Helicopters delivered after 2007 are supplied with the Ahlers gauges which have theindicator lamp but which do not perform the needle swing function test procedure.


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Torque Indicator

TORQUEMETER

72 PSI Red Radial Max for Takeoff

65-72 PSI Yellow Arc 5 Minute limit

0-65 PSI Green Arc Continuous operation

The engine torque indicator is a microprocessor based indicator which uses the signalfrom an engine mounted pressure transducer to indicate engine power. The indicator ispowered by the aircraft electrical system through the TORQUE, or TRQ circuit breaker.

The Horizon/Ultra gauges microprocessor performs a power on self-test when powerinitialized and monitor self reasonableness.

The Ahlers gauges do not incorporate the self-test function. Both types of torquegauge incorporate a red indicator light that will illuminate when 72 PSI is exceeded for1 second. (480B)

Oil Pressure/Oil Temperature

The engine oil temperature and pressure indicator is a microprocessor based dualindicator which uses a temperature bulb located in the engine oil reservoir for engine oiltemperature indications and a pressure transducer connected to the engine oil pressureline on the engine.

The microprocessor in the indicator will illuminate the red indicator lights when theengine oil temperature exceeds 107C for 5 seconds, the oil pressure is less than 50 PSIfor 5 seconds or the oil pressure exceeds 130 PSI for 5 seconds.


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TOT

The turbine outlet temperature (TOT) indicator is a microprocessor based instrumentthat uses DC voltage through the TOT thermocouple harness to indicate the turbineoutlet temperature in degrees Celsius. The indicator receives and averagestemperature indications from four thermocouples mounted radially around the enginebetween the N1(gas producer) and N2 (power turbine) sections of the turbine.

On the TH-28, and some 480 aircraft the TOT system is passive, and on the 480B theTOT system is active. The active system requires aircraft power to be supplied for the

system to indicate. In the event of a main electrical buss failure, the instrument can bedriven directly from the aircraft battery by selecting BAT on the N1-N2-NR-TOTswitch.

The microprocessor in the indicator will illuminate the red indicator light when themaximum TOT exceeds 10 seconds between 810C and 927C during the start, or if theTOT exceeds 810C for 5 seconds during normal operations.

TURBINE OUTLET TEMPERATURE

927C Red Diamond Maximum Temperature, (1 Sec-Starting Only)

843C Red Arrowhead Maximum Transient Limit, (10 Sec on Start)

810-843C Maximum 6 Sec during transient power only

810C Red Radial Maximum starting and takeoff (5 minutes)

737-810C Yellow Arc Maximum 5 Minutes

0-737C Green Arc Continuous Operation


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Transmission Oil Temperature

The transmission oil temperature indicator is a microprocessor based instrument thatdisplays transmission oil temperature in degrees Celsius by means of an electricalresistance type temperature bulb which is located on the left front bottom of the mainrotor transmission. The microprocessor will illuminate the red indicator light in theinstrument when the transmission oil temperature exceeds 107C for 5 seconds.

Gas Producer Tachometer (N1)

The gas producer Tachometer,(N1 Gauge) is a microprocessor based indicator thatuses AC voltage produced by the right side tach generator to indicate the N1 turbinespeed in terms of percent RPM. The microprocessor in the indicator will light the redlight when the N1 speed exceeds 105%. The 480B systems are active which requiresthat they be powered by the aircraft electrical system. In the event of a main electricalbuss failure, the instrument can be driven directly from the aircraft battery by selectingBAT on the N1-N2-NR-TOTswitch.

Fuel Quantity

The fuel quantity system is a capacitance type quantity indicating system and consistsof a fuel quantity indicator, a signal conditioner, and a fuel quantity probe. Theinstrument is a microprocessor based unit that compensates for the non-linear shape ofthe fuel tank to maintain accuracy between the full fuel and the empty fuel quantityreadings. The red light in the indicator will light if the fuel quantity indicates less than35 lbs for 5 seconds.

The right fuel tank has a float switch that activates the low fuel light on the annunciatorpanel when the fuel level reaches between 5 and 8 gallons fuel remaining. The light on

the annunciator panel and the light in the instrument are not interconnected.


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Hour Meter

The hour meter, located in the left side of the center pedestal records helicopter flighttime. It is activated when the engine is running through an oil pressure switch whenthe collective is raised and therefore records helicopter time off the ground.

There is a second optional hour meter available that records time anytime that theengine is running through the oil pressure switch.

Start Counter

The Rolls Royce helicopter engines are start limited (cycle limited). The hot section ofthe engine must be overhauled at 3000 cycles or 1750 hours, which ever occurs first.

The start counter is located in the left side of the center pedestal and it records eachtime that the starter button is activated to track engine cycles. Any time that theengine is motored without the intention of making an engine start, the IGN EXCITEcircuit breaker should be pulled to avoid recording a cycle on the start counter.

Airspeed Indicator/Altimeter/VSI

The pitot system is connected to the airspeed indicator and the forward static system isconnected to the airspeed, altimeter, and VSI.

There are 4 static ports on the 480 series aircraft. The forward two static ports areconnected to the instruments, and the two on the tail cone are reference air for theEngine Inlet Air Caution Segment pressure differential pressure switch.

The forward two static ports have a set of vortex generators that correct the airflowpast the vent.


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Flight Instruments

Airspeed Indicator

AIRSPEED LIMITATIONS

125 Kts Red Line Max Power On Vne

85Kts Barber Pole Max Autorotation Vne

Note: In order to avoid excessive rates of descent in autorotation, it is recommendedthat autorotation speeds be limited to 85 KIAS or Vne, whichever is lower.


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ROTOR SYSTEMS

Main Rotor

The main rotor system is a three bladed, high inertia, fully articulated rotor system.The main rotor hub assembly is composed of two opposing forged aluminum hub platesseparated by an aluminum cylindrical spacer. Through bolts hold these items togetheralong with steel spline adapters.

Three steel universal blocks are mounted on roller bearing units that permit flappingand lead-lag motions. Laminated phenolic pads are used to limit blade travel in boththe lead-lag and flapping axes. A thrust nut on the bottom of each universal blocktransfers vertical blade forces to both hub plates through the universal block.

Dampers

480B Hub


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The rotor blades are secured toeach universal block on the hubthrough a forged aluminum grip

which is in turn secured to a steelspindle assembly through a retention nut or an optional tension-torsion strap assemblyand supporting bearings. On some of the early 480s, centrifugal blade loads are carriedby Lamiflex elastomeric bearing assemblies. On most of the 480 series aircraft and the480Bs, tension-torsion strap assemblies mounted between the blade grip and thespindle are used.

Closed circuit hydraulic dampers are incorporated between each flapping pin and therotor hub to limit the lead-lag velocity of the blades. Because the hydraulic dampershave no centering spring, they are quite limber; this, coupled with the large heavyblades causes the ground rock that is often experienced while the helicopter rotor

system is spooling up.

A single retention pin connects theblade root to the grip and a non-adjustable drag brace connects thetrailing edge of the blade to the grip.


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The main rotor blades are of hollow constructionconsisting of an extruded leading edge spar, with a 7-degree twist, to which is bonded upper and loweraluminum skins. The blade root is composed of abonded doubler assembly.

A cap is bonded to the tip of each blade in which there

are provisions for spanwise and cordwise balanceweights. Two tracking tabs are riveted to the trailingedge of each blade.


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Tail Rotor Assembly

The tail rotor assembly is a twobladed, wide cord, teetering, deltahinged rotor assembly.

The flyweights on the blade retention plates unload the tail rotor twisting forces in flightso that the pilot does not need to carry left pedal in cruse power settings. They areweighted so that when the aircraft is being flown at approximately 50lbs torque, thepedals are neutralized and the slip ball centered.

For this reason, the aircraft requires very little left pedal in hover and in climb, andsignificant right pedal in low power situations.


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Tail Rotor Guard

A tubular aluminum tail rotor guard is installed on the aft end of the tailcone. It acts asa warning to the pilot upon an inadvertent tail-low landing and aids in protecting the tailrotor from damage while the helicopter is on the ground.

IMPORTANT!

The tail rotor guard will not prevent damage to the tail rotor in the event of a tail rotorstrike during a hard landing or auto-rotation flare.


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FUEL SYSTEM

Description

The fuel system consists of two 45-gallon bladder type fuel cells mounted on either sideof the main rotor transmission. Each cell is housed in a composite fuel cell structureand is interconnected to the other fuel cell through a 2" (51mm) fuel crossover line inthe lower forward corner of the fuel cell and a " (13mm) overboard vent crossoverline located at the top of each fuel cell.


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The " (19mm) main fuel supply lines, located at the lowest point in each fuel cell,interconnect at a "tee" to supply fuel to the engine equally from each fuel cell. The

main fuel shutoff valve is incorporated onto the "tee" and is manually operated from thecockpit.

Each fuel cell is equipped with sump drains plus the system is equipped with a low pointdrain at the fuel shutoff valve. A capacitance fuel quantity probe and a low fuel-warning switch are mounted in the right hand fuel cell.

The refueling port is located in the top of the left hand fuel cell. The right hand fuel cellis filled by crossfeeding action during refueling.

On the 480 series aircraft the fuel cells are filled with blue/green open cell foam. Thepurpose of this foam is to help maintain the shape of the cell as the fuel is used and toprevent the fuel from sloshing around in the tank.

If the fuel nozzle contacts the plastic tube during servicing, it can push the plastic tubedown into the foam or it can dislodge some of the blue / green foam inside the tankwhich will then show up in the fuel sumps for a short while.

Blue Foam

Plastic tube tomeasure fuelquantity


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NOTE: Pilots are encouraged to advise line personnel to take care to not

push the fuel nozzle into the tank so that the plastic tube and foam aredisturbed.

To physically measure the fuel level in the tanks. A dip stick can be inserted through thefuel filler opening and into the clear plastic tube that penetrates into the foam. The topreference line on the dipstick must be held against the filler opening. The fuel level canbe read in pounds and gallons on the stick.

The fuel quantity display system consists of a capacitance probe and a quantity-

indicating gauge.

Fuel management can be assisted with the use of a fuel flow monitoring system.

If the aircraft is equipped with a fuel flow measuring system, the actual fuel level mustbe entered into the computer manually each time the aircraft is fueled.

It is recommended that the fuel level verified before the actual quantity is entered intothe fuel flow totalizer.

IMPORTANT!

Note: Avoid using anti-icing/biocidal additives packaged in aerosol cans. Failure toexactly follow the additive mixing procedure during refueling can result in incorrectadditive concentrations, fuel system contamination, and possible engine stoppage.


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POWER TRAIN

Description - Power Train

The power train includes the main rotor transmission, upper pulley, "H"- strut, lowerpulley, lower pulley drive shaft, drive belt, overrunning clutch, power output drive shaft,tail rotor drive shafts, and the tail rotor transmission.


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Drive System

Lower Pulley Drive System

The lower pulley drive system consists of the lower pulley drive shaft and hub, lowerpulley, "H"- strut, and flex pack couplings. The lower pulley drive shaft, located in thehollow center of the lower pulley assembly, is connected to the power output shaft andthe lower pulley assembly with couplings.

The lower pulley shaft is also used to drive the oil cooler blower fan by means of a hubattached to the aft end of the lower pulley drive shaft. The lower pulley has two

positioning links attached to the right side of the bearing housings. The links are usedto laterally align the lower pulley to the engine.


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Thermocouples are installed in the lower pulley bearing housings to providetemperature input for the drive bearing hot caution panel segment (DRIVE BRG HOT).

The "H"- strut, used to tension the drive belt, is connected to the lower pulley and tothe pinion bearing support truss and the main rotor transmission at the upper end. Theflex pack couplings consist of multiple thin stainless steel plates bolted to the drive

flanges. The flex pack couplings will allow up to 1.5 of misalignment between thepower output shaft and the lower pulley drive shaft.

IMPORTANT!Note: Alignment of the lower pulley drive shafts is critical to the integrity of the drivesystem. If the Belt tension adjustment is turned a total of turn, the drive systemalignment must be checked.

There is a component page in the airframe log book to record adjustments to the belttension and lower pulley alignment.


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Description - Overrunning Clutch

The overrunning clutch is installed onthe front side of the engine accessorygearbox. The outer housing of theoverrunning clutch forms the drivingportion of the clutch and is driven bythe engine power output shaft.

In the driving direction, the sprags engage and connect the outer housing to an innerdrive housing which transmits the engine torque to a splined drive shaft that passesthrough the center of the engine power output shaft and the outer housing of theclutch assembly to the rear of the engine accessory gearbox where it is coupled to thelower pulley drive shaft.

Output Shaft

In the overrunning direction, the inner drive shafting, being driven by the rotor system,will rotate faster than the outer housing of the overrunning clutch and the sprags willdisengage thus disconnecting the engine from the rotor drive system. The overrunningclutch is a sealed unit and contains its own lubrication separate from the engine.

OverrunningClutch


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Oil level Sight Glass

There is a sight glass installed on the forward end of the overrunning clutch (SDBT-027). The level of the oil in the overrunning clutch should be performed during pre-flight and if an air bubble is present, the clutch should be serviced before flight.

The flight hours should be recorded at each service of the overrunning clutch and if theclutch requires service earlier than each 20 hours of use, Enstrom recommendsinvestigating the engine for double-lip seal leakage.

Sight Glass


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FLIGHT CONTROLS

Note: the pilot controls are normally on the left side in a 480 and on the right side in aTH28.

The flight controls include three primary systems: the collective, cyclic, and anti-torqueor directional controls. The aircraft also has fixed horizontal and vertical stabilizers

mounted on the tailcone to provide additional stability and attitude control during high-speed flight.


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Collective Control System

The collective control system is comprised of dual collective controls mechanicallyinterconnected and linked to the main rotor swashplate through a series of push-pulltubes, a torque tube, a bellcrank, and a collective walking beam at the base of the mainrotor transmission.

Both collective controls have interconnected twist grip throttles and a switch boxmounted forward of the throttles. The pilots collective incorporates the landing lightcontrols, N2 power turbine governor increase/decrease (GOV INCR/DECR) switch(governor beeper), and the engine starter switch.

The copilot's collective switch box only has two switches; a landing light attitude controlswitch, and a N2power turbine governor increase/decrease (GOV INCR/DECR) switch.


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The pilot's collective control incorporates a collective friction system located on thecollective control. The collective friction system consists of a simple stop bracket thatincorporates both the up and down collective stops and a knob/lever assembly used toclamp two friction disks to the stop bracket

Idle DetentButton

StarterButton

Throttle

Landing Light Angle

Adjustment

GovernorBeeper

LandingLight

Switch


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Collective Friction

When the Friction lever is moved to the horizontal position by rotating the knob/leverassembly 90 degrees in the forward direction, the friction is fully applied.

The control may be positioned in any intermediate position for any desired level offriction. The collective friction system is designed so that positive locking of thecollective controls cannot be obtained at the maximum friction point. Safety of flightconsiderations require that the pilot be able to instantly overcome the establishedfriction without any further pilot action to adjust it in the case of engine failure.

Collective control forces are reduced by means of a collective trim system located aft ofthe collective bellcrank in the engine compartment. The collective trim system consistsof a spring capsule, bracketry, and an adjusting link.


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Cyclic Control System

Description

Note: The cyclic control system is afully mechanical control system which

is linked to the swashplate through aseries of interconnected push-pulltubes, a torque tube, and bellcranks.Both longitudinal and lateral controlsystems are totally independent withno intermixing before the individualinputs reach the swashplate.

Non-rotating control inputs are transmitted to the rotating controls via a universal jointtype swashplate at the base of the transmission. Inputs are mixed at the swashplateand transmitted through a set of three push-pull tubes though the center of the mast topitch change walking beams at the top of the hub. The motion is then transmittedthrough pitch change links to the blade pitch horns located on the leading edge of eachblade.


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The aircraft is equipped with a cyclic stick, located directly in front of the pilot seats.The switches mounted on the cyclic grip assembly are all non-functional (before the

installation of optional equipment) except the four way toggle switch at the top centerof the grip, used to control the four way cyclic trim system. The cyclic trim systemmaintains the position of the cyclic control stick and reduces rotor feedback to zero.

Cyclic Grip

Trim Switch

Freq. Transfer

Radio & IntercomTransmit Switch

TransponderIdent.


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Cyclic Trim

The system consists of a cyclic trim switch located at the top of each cyclic grip, a pairof electrically operated jack screw actuators that vary spring tension produced by thelongitudinal and lateral trim units, and a pair of trim switch units which reverse thedirection of the current operating the actuators.

The cyclic trim switches each have five positions which are: normally OFF in the center,and momentary FORWARD, AFT, LEFT, AND RIGHT. Both trim mechanisms include anelectrically operated reversible motor and a cylindrical spring assembly connected to the

cyclic control linkage and both are mounted on the cabin bulkhead in the upper enginecompartment.

When a trim switch is moved off of center to any one of the four trim directions, powerapplied through the TRIM circuit breaker energizes one of the trim motors to apply trimspring force in the desired direction. By momentarily moving the switch, very small trimincrements may be obtained.

Trim force cannot be applied in two directions simultaneously; when both longitudinaland lateral trim corrections are desired, it is necessary to apply first one and then theother. The cyclic trim system does not limit travel of the cyclic control; the pilot may

override the trim forces at any time.


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POWER PLANT

Description

The Enstrom TH-28 and 480 series helicopters utilize an Allison designed, Rolls Roycebuilt 250-C20W reverse-flow freepower-turbine turbo-shaft engine. The enginemaximum rating is 420 shaft horsepower. In the 480B, the engine is rated at 277shaft horsepower continuous and 305 shp for 5 minutes. This is a transmission basedlimitation.

In the TH28 and early 480 aircraft, the engine is derated to 256 shaft horsepowermaximum continuous power and 285 shp for 5-minutes.


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Engine Compartment Cooling

The engine and transmissions are cooled by a fan cast into the transmission pulley.There is a baffle installed at the bottom of the transmission to separate the enginecompartment from the transmission compartment. Beginning on helicopters deliveredin 2008 and on helicopters that have the high temperature modification installed, thebaffle has been removed.

The fan pulls air through the transmission and engine area and over the transmission oil

cooler tubes to cool the MRGB oil.

The engine oil is cooled by a fan installed just forward of the baggage compartmentwhich draws air in to a vent on the left side forward of the baggage compartment door,through the fan, and out through the oil cooler on the right side of the helicopter.

Beginning on helicoptersdelivered in 2008,additional ventilation holeshave been added in the

lower engine cowl panelsand the bottom enginecover access hole has beenopened up to assist inengine cooling.


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The high temperature cooling kit also includes ventilated engine side cowls which raisesthe maximum operational ambient air temperature from 106F (41C) to 122F (50C).


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Air Induction System

The upper plenum chamber is equipped with dual full flow swirl tube inertial typeparticle separators. Engine inlet air is directed into the upper plenum chamber througha series of swirl tubes which impart a centrifugal spin to the air as it enters the tubes,thereby inertially separating the heavier foreign matter.

The particulate matter falls down into a collector and is then purged overboard throughone of two bleed air driven venturi-type ejectors that exit at the aft side of the upperplenum. Operation of the scavenge ejectors is manually controlled by a handle

mounted in the cockpit.

During takeoff, hovering, or cruise operations in dusty atmospheric conditions the bleedair shutoff valve can be opened by placing the SCAV AIR control handle in the ONposition. The inlet air moves from the upper plenum chamber to the lower plenumchamber via two (2) transfer ducts located on either side of the aircraft.

Note: Use of the engine Anti-ice and Scavenge Air will increase the TOT. The enginemay not deliver rated power when they are engaged.

The lower plenum chamber is mounted directly behind the engine and is connected tothe engine by a bell mouth inlet. The inlet is attached to the engine and a foam-rubbergasket on the inlet provides the seal between the inlet and the lower plenum chamber.The lower plenum chamber has drain holes located at its lowest point to drain anymoisture that might happen to accumulate during operation or while the engine is notrunning.

A fitting is also installed in the left side for connection to the ENGINLET AIR caution light differential pressure switch.

Engine Oil System

The engine oil system consists of an engine oil reservoir, oil cooler, blower assembly;scavenge oil filter, and connecting lines and fittings. The oil reservoir is located on theright side of the engine compartment and is accessible through the right side engineaccess panel.


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Care must be taken to be sure that the cap is straight and flush before it is

latched to prevent leakage.

The oil cooler is located on the right side of the aircraft and is accessible through the oilcooler access panel. The scavenge oil filter with an integral impending bypass pop-outindicator, located at the bottom of the filter bowl, is located on the right side of theaircraft and is accessible through the step access panel, and the right engine accesspanel.

The blower assembly is located aft of the lower drive pulley. The assembly consists ofa fan mounted on a drive shaft which is mounted on a platform, a connecting drive

shaft between the lower pulley and the fan drive shaft, and air intake and exhaustducts.


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OPERATIONAL PROCEEDURES

AIRCRAFT SERVICING

Fuel

Approved Standard, Alternate, and Emergency Fuels

TYPE SPECIFICATION LIMITATIONS

Primary MIL-T-5624 JP-4 & JP-5,MIL-T-83133 JP-8, ASTMD-1655 Jet B, Jet A, and

A1 (See note 1 below)

JP-1, Diesel #1, or ArcticDiesel DF-A (VV-F-800B)conforming to ASTM D-1655, Jet A orJet A1

With anti ice additiveconforming to MIL-I-27886

Emergency MIL-G-5572E AVGAS(without TCP)

All Grades, Maximum 6 hoursoperation per overhaul period.

With anti ice additive

Cold WeatherMIL-T-5624 JP-4 ASTM D-1655 Jet B Avgas-Jet fuelmixture

(See note 2 below)


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Note 1: All fuels used in the TH-28/480 shall contain Anti-icing and Biocidal Additiveconforming to MIL-I-27686. The additive provides anti-icing protection and also

functions as a biocide to microbial growths in helicopter fuel systems. Icing inhibitorconforming to MIL-I-27686 shall be added to all commercial fuel, not already containingan icing inhibitor, during refueling operations, regardless of the ambient temperatures.Refueling operations shall be accomplished in accordance with accepted commercialprocedures. Commercial product "PRIST" conforms to MIL-I-27686.

Note 2: The AVGAS-jet fuel mixture is an alternate fuel which may be used if startingproblems are encountered in areas where JP-4 or commercial Jet B cannot be obtained.The mixture shall be one part by volume AVGAS to two parts by volume commercial jetfuel. The AVGAS shall conform to MIL-G-5572C, grade 80/87, or grade 100/130 with a

maximum of 2.0 ml/gal lead content. Do Not use grade 100/130 with 4.6 ml/gal leadcontent. (The 2.0 ml/gal max. lead content grade 100/130 AVGAS is known as 100L inEuropean areas). The commercial jet fuel may be kerosene; JP-5 or commercial Jet Aconforming to MIL-T-5624, grade JP-5 or ASTM D-1655, Jet A or A1.

IMPORTANT!Note: Avoid using anti-icing/biocidal additives packaged in aerosol cans. Failure toexactly follow the additive mixing procedure during refueling can result in incorrectadditiveconcentrations, fuel system contamination, and possible engine stoppage


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Engine Lubrication Oils

Note: Rolls Royce recommends not mixing oils within a series unless absolutelynecessary.

Warning: Mixing of oils from different series is not permitted. (MIL-L-7808 or MIL-L-23699). Refer to the Rolls Royce 250-C20 series operation and maintenance manual.

Approved Domestic Commercial Oils for MIL-L-7808

Manufacturer Manufacturers Designation

American Oil American PQ Turbine Oil 8365

Brayco Oil Brayco 880H

EXXON Company EXXON Turbo Oil 2389

Mobil Oil Mobil Avrex S Turbo 256, Mobil RM-201A

Mobil Oil Mobil RM-184A

Stauffer Chemical Stauffer Jet 1

MIL-PRF-23699F Series High Thermal Stabili ty (HTS)


Royal Lubricants Company Aeroshell / Royco Turbine Oil 254

Shell International Petroleum Co. Aeroshell Turbine Oil 560

Mobil Oil Mobil Jet Oil 254


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Approved Domestic Commercial Oils for MIL-PRF-23699(Formerly MIL-23699)


Mobil Oil Mobil Jet II

NYCO S.A. Turbonycoil 600 (TN600)

Royal Lubricants Company Aeroshell / Royco Turbine Oil 500

EXXON Company EXXON TURBO OIL 2380

Stauffer Chemical Staufer Jet 11 (Castrol 205)

Caltex Petrolium Corp. Caltrex RPM Jet Engine Oil 5

Chevron International Oil. Co. Cheveron Jet Engine Oil 5

American Oil and Supply Co. American PQ Lubricants 6700

Castrol Inc. BRAYCO 899

Hatcol HATCOL 3211

Air BP Air BP BPTO 2380

WARNING; ONLY DISCRETIONARY MIXING OF OILS WITHIN AN OIL SERIES ISPERMITTED WITHOUT A TIME PENALTY. SEE THE ROLLS ROYCEMAINTENANCE MANUAL FOR SPECIFIC INFORMATION.


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PERFORMANCE DATA

The following general conditions are applicable to the performance data:

1. All airframe and engine controls are assumed to be rigged within allowabletolerances.

2. Normal pilot technique is assumed, control movements should be smooth andcontinuous.

3. No two aircraft are exactly the same; however variations are considered to besmall and cannot be individually accounted for.

4. The data presented presumes that all instruments and systems have beenproperly maintained, are in proper working condition and are calibrated.

Torque Available

The torque available chart shows the effects of altitude and temperature on enginepower available. The primary use of the chart is to provide the pilot information onthe maximum power available either as a function of the helicopter limits or theflight conditions.

Operation of the engine anti-ice, scavenge-air or bleed air will result in higher TOT,N1 speed and fuel flow. Because the engine is de-rated and torque limited, torque

will normally be the limiting consideration, however in some conditions TOT or N1may approach operational limits and limit the power available.

See table 4-1 in the RFM for exact information as to the effects of engine bleed airusage on performance.

By using maximum torque available chart, and the hover chart, the pilot candetermine the power margin for an intended operation.


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Hover Performance


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Cruise Performance

The cruse performance charts show the torque pressure and fuel flow for level flight atvarious pressure altitudes, airspeeds and gross weights. The two charts on the page arenot connected together. To obtain the expected fuel flow, knowing the torque, the leftchart must be used to obtain the airspeed. This airspeed is then entered into the charton the right and the fuel flow can be obtained. There are charts for sea level, 3000,6000, and 9000 feet pressure altitude.


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Power Assurance Check

The power assurance chart provides the pilot with a method to assure that the installedengine will provide the expected power required to achieve the performance presentedin the RFM and also to monitor the engine performance over time.

The pilot should come to a stabilized hover, record the pressure altitude, OAT, Torqueand TOT, and then land and plot the actual data on this power assurance chart. If the

actual TOT is less than or equal to the TOT determined from this chart, then thehelicopter be expected to achieve the pre-flight calculated performance for the flight.

The conditions for using this chart are 103% N2 engine RPM, and 372 rotor RPM at astable hover. It may be necessary to hover for at least 2 minutes to allow the engine tocompletely stabilize before taking engine performance numbers

During a hover, The TOT, Torque, Pressure Altitude (29.92), and Temperature arerecorded. These parameters can be plotted on the Power Assurance chart to see thatengine is performing within the recommended parameters.

If the actual TOT reading is higher than the calculated TOT of the Engine PowerAssurance Chart in the TH-28/480 Operators Manual and pilot error has been ruledout, the engine will not meet the performance figures called out in the PerformanceData section of the TH-28/480 Operators Manual.

Turbine engines normally degrade or lose power through engine operation (Refer to theTrend Check Procedure section of the Rolls Royce 250-C20 Operation and MaintenanceManual for further information). A gradual increase in TOT that is above the normalengine degradation trend line is normally caused by a dirty compressor.

Performing a compressor wash will usually return the TOT indications to the normalengine degradation trend line.

If the actual TOT readings significantly increase or decrease over a shortperiod of time, a potential problem may exist with the engine or an airframesystem.


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Height Velocity Diagram


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Engine

Engine Power Controls

The engine power control system is a mechanical linkage/ cable system, actuated by atwist-grip on the collective sticks, which provides manual control of the power lever onthe fuel control unit.

A flight idle stop located above the throttle twist grip is incorporated in the pilot's

collective stick. The stop prevents the engine power setting from being reduced belowthe flight idle position causing an accidental engine shutdown. The release does nothave to be pushed for engine start or run-up but does have to be pushed for engineshutdown. Rolling the throttle back past the stop has the effect of shutting off the fuelflow to the engine, thereby shutting it down.


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An electrically operated linear actuator operates a lever connected to the power turbinegovernor. The linear actuator controls the power turbine (N2) RPM and is operated via

the GOVN INCR/DECR switch located on the control box on the collective sticks.

A droop compensation system is incorporated to stabilize N2RPM as the engine loadfluctuates with changes in the main rotor pitch. It consists of a series of linkages andbellcranks that connect the collective control to the governor on the engine anaccumulator that is installed in the engine PC system. The linkage is designed in such away as to cause the governor to anticipate the requirement for increases or decreasesin power requirements to help prevent the rotor speed from lagging behind the pilotinputs.

The accumulator in the PC lines acts as a spring and dampens out the governorreaction to smooth the power changes.

Understanding the operation of and the relationship between the governor and the fuelcontrol will help prevent the pilot from getting into some situations which can lead toaccidents.

The governor reads the N2 RPM (output-shaft speed of the engine) and schedules thefuel to the fuel control to keep the N2 speed, (also the rotor speed) constant. Thegovernor is sensing RPM and will always attempt to maintain this RPM. Becauseturbine engines take some time to spool up, the purpose of the droop compensator is

to lead the pilot-control inputs to compensate for this time lag.

It is important that the pilot understands that the governor reacts to and compensatesfor changes in the N2 (Rotor Speed), and that both changes in collective and tail rotorinputs will affect N2. However, the droop compensator system is only connected to thecollective and not to the tail rotor, so there are possible situations where large left tailrotor inputs can spike the torque as the governor tries to react to, and compensate forchanges in the Main Rotor RPM. The engine will not respond to a tail rotor input untilthe main rotor RPM (N2) changes, so there may be a lag in the engine response.

IMPORTANT!

Closing the throttle during autorotations can result in situations where the governorcannot react to the pilot inputs with the resulting loss of the aircraft.


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During autorotation, and especially in the flare, the rotor RPM will be on the high end ofthe scale. Especially if the throttle has been closed, the governor will try to react to this

high Rotor RPM by scheduling back the fuel to the fuel control in an effort to slow therotor. So at the maximum point of the flare, just as the pilot needs to roll the throttleback on for the recovery, the governor is still trying to slow the rotor by cutting backthe fuel!

When the pilot rolls the throttle back on, the N1 will be slow, and due to the lag in theturbine engine, there may not be enough time to power up for a successful recovery.

As the blade RPM decays, the pilot will need to add more left pedal, and of course thisonly aggravates a bad situation! The addition of left pedal drags the rotor RPM downeven more.

When the governor does react to the rotor RPM which is now low due to theautorotation recovery and the addition of left tail rotor pedal, the governor is behindand the result is a spike in torque as it tries to catch up.

Enstrom Helicopter recommends that the throttle not be rolled off during practiceautorotations. The pilot should begin a power recovery by increasing collective pitchslightly during the flare to give the engine a jump start on delivering the power that willbe needed in the recovery.

Allowing the helicopter to drift forward, at least one rotor diameter during the

autorotation recovery will assist in keeping the helicopter in clean air and willdramatically reduce the torque necessary for the recovery


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Starting Procedures

IMPORTANT!

Note: During a start the throttle must never be advanced out of the fuel off positionuntil after the starter has been energized and the desired cranking speed has beenattained

Hot Starts

For the most part, hot starts are a result of not enough air passing through the enginewhen fuel is added. Usually this is a result of either a low batteryor improperstarting procedure.

Low Battery

The C-20 engine requires starter assistance to boost the acceleration through the start

sequence. The starter will accelerate the engine to a maximum of 20% on its own.

The engine will not accelerate past about 30% without the assistance of the starter sothere are two consequences of low battery.

1. The engine does not achieve enough RPM before light-off and the TOTexceeds the manufactures limits due to insufficient air flow.

2. The engine hangsin the 30 % range and wont carry through. (This can alsobe the result of an overheated starter) (Also known as a stagnated start)


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Starting procedures

There are three or four common procedures in general use for starting Rolls Royce 250C-20 B and W series, engines. Any of these procedures that the pilot is comfortablewith can be used for starting the Enstrom 480 helicopters. The pilot needs to be awarethat due to the design of the 480 collective throttle and detent system, if the throttle isrolled off and held against the idle stop detent position, the detent button cannot bedepressed. Pressure against the detent position must be relieved before the detentbutton can be depressed.

The following procedure is an alternate method of starting the helicopter that can beused that will alleviate the difficulty of depressing the detent button to roll of the

throttle and aborting a start in the event that the TOT temperatures are not withinacceptable tolerances.NOTE

The starting sequence on the C-20-W engine with the Bendix fuel control iscompletely automatic. The only action that the operator can effect is to shutoff the fuel and terminate the start. Opening the throttle further to speedthe acceleration or even opening the throttle fully has no effect. (Until theengine reaches 58%)

The reason for not opening the throttle to the detent during the start is thatin the case of a hot-start, if the operator is holding the throttle pressure

against the detent button, it can be difficult or impossible to depress the lockbutton and abort the start. Also, not passing the detent position gives theinstructor in the co-pilot seat the ability to abort the start in case the studentdoes not recognize an impending hot start.

If the throttle is advanced past the idle position before the engine reaches58%, there will be a surge in torque as the engine passes 58% that might beuncontrollable. This is because the governor is sensing that the operatingRPM is less than the throttle position is calling for and it will schedule fullfuel to catch up. The C-20 engine will develop maximum torque at situations

where the RPM is low and increasing.


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Recommended starting procedure follows

1. Master on2. Check TOT (Do not initiate the start unless the TOT is 150C or less)3. Open the throttle past the cut-off detent and then close it back to the full

off position. (Again, past the detent)4. Press the starter button and when the Ni reaches 12% begin to slowly open

the throttle.5. When the engine starts, stop turning the throttle6. (If the engine does not start within 3 seconds, close the throttle to the full

shut off position, release the starter button, allow the system to drain andstart again.)

7. After the engine TOT peaks, tweak the throttle open just a bit to preventinadvertent shut-down.

8. When the N1 reaches 58% carefully advance the throttle until the detentbutton pops up.

9. Allow the N2 to reach a maximum indication and stabilize, and then turn onthe generator.

10.Monitor N1 and TOT when turning the generator switch on. If N1 decaysbelow 60% or TOT approaches 750C, turn the generator off and increaseN1 speed with the throttle to 70% and reset the generator switch to on.

11.Wait for the charging amperage to decrease to below 50 amps and turn onthe Avionics Master Switch.

12.Do not exceed 40PSI torque while rolling the N2 RPM to 98%.

IMPORTANT!

Note: If the throttle detent button pops up before light off, or before peakTOT is reached, push and hold it down until after peak has been passed.


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Recommended procedure for starting the engine when warm or hot

1 Master on2 Check TOT (Do not initiate the start unless the TOT is 150C or less)7 Engage the starter button and motor the engine while watching the TOT. When

the TOT has reached 150C, slowly open the throttle.8 During the cooling period, the pilot should also monitor the N1. Normally, if the

TOT was higher than 150C, the N1 will reach more than 15% before the TOT isat 150C.

7 After the engine starts, stop opening the throttle.

8 After the TOT peaks, tweak the throttle open just a bit more to preventinadvertent shut-down.

9 When the N1 reaches 58% carefully advance the throttle until the detent buttonpops up.

10Allow the N2 to reach a maximum indication and stabilize, and then turn on thegenerator.

11Monitor N1 and TOT when turning the generator switch on. If N1 decays below60% or TOT approaches 750C, turn the generator off and increase N1 speedwith the throttle to 70% and reset the generator switch to on.

12Wait for the charging amperage to decrease to below 50 amps and turn on thegenerator. Do not exceed 40PSI torque while rolling the N2 RPM to 98%.

13Do not exceed 40PSI torque while rolling the N2 RPM to 98%.

IMPORTANT!

Note: If the throttle detent button pops up before light off, or before peak TOT isreached, push and hold it down until after peak has been passed.


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Hung Starts. (Stagnated Starts)

A stagnated start occurs when the engine stops accelerating before it reaches 58% N1.If this occurs in the 20 to 30% range it is most likely due to a cold soaked engine. Inthis case, waiting five minutes and reinitiating the start will usually be successful assitting allows the engine to warm from the initial light off. If it occurs in the 40 to 50%range during the first start of the day, it is usually a scheduling problem in the fuelcontrol and is field adjustable.

If a stagnated start is experienced, shut down, wait 5 minutes and repeat thestart procedure.

The start should be complete in one minute: however, if N1 and N2 areaccelerating and TOT is within limits, the start may be continued longer thanone minute.

The engine will have warmed up by the second start and will operate normally. IF thesubsequent warm starts are cool enough, the start-acceleration can be clicked up onenotch and the cold start checked at the next flight.

Dramatic changes in altitude or temperature may necessitate occasional readjustmentsof the start-acceleration.

Rolls Royce defines a good first start as one taking less than 25 seconds from theintroduction of fuel until the engine reaches ground idle. Judging starts should beaccomplished on the first start of the day, with a fully charged battery in ambienttemperatures of 40F or above.

To obtain an optimized start, move the throttle to begin fuel flow as the N1 RPMaccelerates through 12 15% N1. Do not wait for N1 RPM to peak out before initiatingfuel flow, as this will unnecessarily utilize battery capacity early in the start cycle.

Consistent long, cool starts (35 seconds of more) can be detrimental to the gasproducer turbine life. (N1 wheels and Nozzles) The probes for the TOT system areinstalled radially around the engine between the gas producer and the power turbinesections and may not reflect the actual temperatures being experienced by the firststage turbine wheel and nozzle assembly. The recommended quick warm startsactually increase cooling air flow in the combustion section to help cool the gasproducer turbine


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There are two procedures in general use to cool and start a hot engine.

1. Cool and start the engine in one event.See the above recommended procedure for starting a warm engine.

2. Cool and start the engine in two events.a.Use the above procedure with the exception:b.Pull the ignition excite circuit breaker.c. Motor the starter until the TOT is 200 C or below.

d.Push in the ignition excite circuit breaker.e.Use normal starting procedure taking care not to introduce fuel until the

TOT is 150 C or below.

The engine requires the starter to help accelerate the engine through the 30 to 50%range, but the starter by itself can only carry the engine to around 20%.

The engine also requires the boost from the starting acceleration to help carry itthrough the start. The earlier (lower RPM) that the start is initiated, the more of thestart is carried through by the natural acceleration of the engine, and the less work that

the starter has to accomplish.

On a cold soaked engine, if the RPM is too high when the start is initiated, there maynot be enough boost from acceleration to carry the start, and the result is that the startmay stagnate. This is due to the natural tendency of the battery to deliver lower voltagewhen cold, and the additional resistance of a cold engine.

On a warm or hot engine, the battery will be drained less by cooling the engine andmaking the start in one event, than in two. It is possible to get a stagnated start with awarm or hot engine if the starter has overheated.

If a stagnated star occurs during a warm or hot start, allow the engine to cool for 15 or20 minutes and use an APU for the start. Initiate the fuel as the N1 passes 12% formaximum boost from the acceleration.


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The recommended TOT temperatures and time for start are as follows.

First Start of the day:

TOT 750C, 20 to 40 seconds from engine light-off.Maximum time one minute.

Starts with the engine at 150C

The limits for TOT during start are as follows.

1. 843C, 10 seconds on start (Red arrowhead)2. 927C. 1 second on start. (Red Diamond)

(Rolls Royce has changed the allowable time between 810C and 926C to 10 seconds.The Enstrom 480 Aircraft Hand Book will be revised to reflect this change.)

(You may notice that the red arrowhead appears to be at the incorrect position on theTOT gauge. There is a scale change at 850C on the gauge. 850C is the last markbefore 900C.)

Note: It is recommended that the pilot abort the start if the TOT is saccelerating at 850 TOT. Generally there is a rapid increase in TOTbetween 850 and 927C and if the pilot delays shutting off the fuel unclose to 900C, the TOT will exceed 927C.


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EMERGENCY PROCEDURES

This section describes the foreseeable helicopter and systems emergencies andpresents the procedures to be followed. Emergency procedures are given in checklistform when applicable.

Definition of Terms

Immediate Emergency Actions. Those actions that must be performed immediatelyin an emergency procedure are underlined. These immediate emergency actions mustbe committed to memory.

Note: The urgency of certain emergencies requires immediate and instinctive action bythe pilot. The most important single consideration is helicopter control. All proceduresare subordinate to this requirement.

Urgency to Land

Land Immediately - Perform a landing at the closest suitable landing site.

Land as Soon as Practicable - Land at the nearest suitable airport or landing facility.

Emergency Exit

To exit the cabin in the event of an emergency, first attempt to open the doors. If thedoors will not open, break or kick out the door windows, overhead windows, orwindshields as the situation requires.


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Engine Failure

The indications of an engine failure, either a partial power loss or a completepower loss are:

A left yaw caused by the drop in torque applied to the main rotor. A drop in engine (N2) RPM. The ENGINE OUT warning light and audio triggered by the N1 speed

dropping below 58%. A change in engine noise.

Immediate reaction to an engine failure or power loss is essential. After

immediate emergency actions have been accomplished, verify the enginefailure by cross checking all of the engine instruments.

Note: The first indication of an engine failure will normally be anuncommanded left yaw of the nose of the helicopter. The engine-outwarning horn is activated by the N1 tach-generator. If there is a horn, andNO left yaw, before entering autorotation, verify that the engine is actuallynot providing power and that the problem is not actually an instrumentfailure.

If the pilot instinctively reacts to a warning horn or light by immediatelybottoming the collective, it is recommended that after the glide isestablished, the pilot try gently raising the collective while monitoring thetorque gauge. If the torque reading shows an increase the pilot would beadvised to attempt to reestablish powered flight, while considering the

possibility of a false engine out indication.

Under partial power conditions, the engine may operate smoothly at reduced power orit may operate roughly and erratically with intermittent surges of power.

In instances where a power loss is experienced without accompanying engine

roughness or surging, the helicopter may sometimes be flown at reduced power to afavorable landing area; however, under these conditions the pilot should always beprepared for a complete power failure at any time.


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After an engine failure in flight, an engine restart may be attempted if time and altitudepermit. Because the exact cause of engine failure cannot be determined in flight, the

decision to attempt the restart will depend on the altitude and time available, rate ofdescent, potential landing areas, and crew assistance available.

Under ideal conditions, approximately 30-45 seconds is required to regain poweredflight from the time the attempted start is begun if the start is commenced with anengine that is not windmilling. If the engine start button is depressed immediately afterautorotation has been established, powered flight can usually be resumed within amatter of 20 to 25 seconds.

There are two alternative types of restart that will be discussed below. The first is an

immediate relight; the second is a restart from a full shutdown with the N 1below 15%.

Note: Unless there is a reason to believe that the engine has failed due tosome obvious mechanical failure, always attempt relight immediately afterentering autorotation if time and altitude permit.

Immediate Engine Relight

Although there is a formal step-by-step checklist provided below for engine restart in

flight, if circumstances such as terrain or flight condition require an immediate relightattempt, the procedure need only involve two steps;

1. Enter autorotation

2. Depress and hold the starter button.

The throttle does not have to be retarded to idle if the elapsed time between failureand attempted relight has not exceeded 5 seconds. There will be a slight surge as theengine comes back on line but it will be well controlled and will not damage the engineor drive train. To control the engine surge as it returns to flight RPM, it is recommended

that the rotor RPM be reduced to minimum (334 RPM).

3. Land Immediately - After the engine is started and powered flight isreestablished, perform a power on approach and landing without delay if the enginewas not intentionally shutdown.


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Engine Restart - During Flight

1. Establish Autorotative glide

2. Attempt Start:

Throttle Closed

Starter Button Depress Throttle - Idle (N115% or greater) TOT and N1 Monitor Starter Button - Release (at 58% N1) Throttle - Advance to Full ON (N2/NRneedles rejoined)

Powered Flight - Resume

3. Land Immediately - After the engine is started and powered flight isreestablished, perform a power on approach and landing without delay if the enginewas not intentionally shutdown

Documents

Enstrom 480B TM