Enstrom 480B Training Manual

TH-28/480/480B Training Manual 5/29/2008 2007 Edition 2 For Training Purposes Only

Introduction 3 Aircraft Description 4 Construction details 7 Aircraft Systems 17

Electrical 17 Caution and warning 23 Instruments 29 Rotor Systems 37 Fuel System 42 Power Train 45 Flight Controls 50 Power Plant 57

Operation Procedures 63

Aircraft Servicing 63 Performance Data 67

Engine 74 Emergency Procedures 84

Weight & Balance 98

Pilot Notes 105

Table of Contents


INTRODUCTION Enstrom TH-28/480 Series Helicopter Pilot Training Course Objectives The purpose of this course is to prepare an experienced helicopter pilot for a smooth transition into the Enstrom Turbine powered helicopters. This course includes descriptions and theory of operation for the systems, and the location of the system components. The course also includes the description of the pilot pre-flight procedures and the pilots are expected to perform these pre-flight inspections.


AIRCRAFT DESCRIPTION The TH-28/480 helicopter is a 3 bladed, single engine helicopter manufactured by the Enstrom 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


Characteristics Helicopter Description The Enstrom 480B helicopter is a single-engine; turbine-powered helicopter certified for day and night VFR flight, that can be equipped for IFR flight. The 480B was developed for light commercial, municipal, and military uses and was certified to FAR 27 standards in 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 low noise signature.


The Enstrom 480B features a three-bladed, fully articulated main rotor system which has over 3,000,000 flight hours and which has never had a catastrophic failure or thrown a blade. The tail rotor is two bladed and completely unblocked for exceptional effectiveness. Due to the high inertia rotor design, the helicopter possesses outstanding auto-rotational capabilities. There have been no fatalities from any accident involving 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 gear and airframe feature an integrated energy absorbing system. The 480 comes standard with high skid landing gear which consists of aluminum skid tubes and nitrogen air-oleo struts to cushion ground contact. Replaceable hardened steel 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 be procured and operated for minimum costs. The basic modular design is simple and inexpensive to manufacture. The helicopter does not require hydraulic boost, electric boost pumps, or a stability augmentation system. The entire control system consists of mechanical linkages. The avionics package is designed for easy installation and accessibility and the 480 is configured with five hinged doors and five removable panels for maintenance accessibility. The limited number of fatigue critical parts, the long overhaul intervals, and the low maintenance hour/flight hour ratio resulting from high reliability and easy maintenance combine to yield low operating and support costs.


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


Cabin Floor and Backwall

Keel Attach Structure


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


Landing Gear Main Landing Gear: The main landing gear consists of two tubular aluminum skids attached 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 movement during ground contact maneuvers. Due to their design, the drag struts will sustain landings with significant forward movement of the helicopter; however, landing with rearward 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.


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

The ground handling wheels must be removed before flight.

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


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


Cabin Doors The two cabin doors are composite reinforced structure with transparent plexiglass windows in the upper section. Ventilation is supplied by sliding vent windows to draw fresh 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 air ventilation system or a bleed air heating system depending on the optional equipment installed 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.


Seats The pilot and copilot/passenger seats are adjustable for fore and aft positioning and are easily removed from the aircraft to facilitate maintenance on the seats or the cockpit area. 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, and an inertia reel mounted on the back of the seat pedestal are an integral part of the seat. Passenger Seats: The passenger seats are mounted to the pylon assembly through the 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 inertial reel.


Inertia Reel Shoulder Harness An inertia reel and shoulder harness is incorporated in all seats. There is no independent control to manually lock the harness. With the shoulder straps properly adjusted, the reel strap will extend to allow the occupant to lean forward; however, the reel 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 tension on 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 harness inertial reel locking up and preventing the crew member from leaning forward. For this reason the shoulder straps must be adjusted so that the y splice is close to the seat occupants neck.


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 SHP for 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 277 SHP continuous.

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


AIRCRAFT SYSTEMS

ELECTRICAL SYSTEMS Description - Starter/Generator Systems Rolls Royce helicopter engines are cycle limited (start limited). Overhaul of the hot section, (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, total number of starts, pilot name and the purpose of the flight. For the pilot / owner, it is worth considering removing the start counter and substituting a trip log in its place.


Battery shelf, Master and Starter Relays and GPU Plug The starter system is used to start the aircraft powerplant. Using either battery power or external power and with the battery switch turned ON to supply electrical power to the main electrical terminal strip, the starter system is engaged by pressing the start switch located on the pilot's collective stick control head. When the start switch is engaged, the start relay coil is energized and electrical power is supplied to the starter side of the starter/generator and electrical power is supplied to the ignition exciter and the start counter.

The starter ignition lock is key operated and must be ON for power to be applied to the starter-generator, ignition exciter, and the start counter. The starter/generator cannot be engaged with the starter ignition lock in the OFF position. However, once the helicopter 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 OR WITH THE KEY REMOVED.


The generator system is used to supply 28 Vdc electrical power to the main electrical terminal strip and to recharge the battery after a battery powered start. The generator system is controlled by the generator control unit (GCU). The GCU performs the following functions: voltage regulation, overvoltage protection, reverse current protection, 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 switch is placed in the on position, the GCU connects the starter-generator to the main electrical 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.


GCU plug The ground power plug is installed on the inside forward bulkhead of the right side engine compartment just in front of the battery. Beginning in 2007 there is an optional access door available.

When the APU is being used to assist starting the 480 helicopters, the battery switch must be in the on position. It is important for the pilot to instruct the ground crew not to disconnect the APU power untill after the start is completed and the pilot gives 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 electrical supply.


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


At flight idle RPM and above, the voltage regulator portion of the GCU maintains the correct generator output voltage by varying the generator field current. If the generator voltage exceeds 32.0 Vdc .5 Vdc, the internal overvoltage sensor will cause current to flow in the trip coil of the generator switch and trip the switch to the OFF position. This removes the current from the generator field and power from the generator relay-actuating coil, disconnecting the starter/generator from the main electrical terminal strip. The reverse current portion of the GCU de-energizes the generator relay when the generator output voltage falls below the battery voltage.

The over-current protection circuitry will cause current to flow in the trip coil of the generator switch when the generator maximum output current rating is continuously exceeded for 10 seconds 2 seconds. This trips the generator switch to the off position removing the current form the form the generator field and the power from the generator relay-actuating coil. The circuitry in the GCU will illuminate the generator caution light (DC GEN) in the caution panel any time the generator voltage is less than the battery voltage, the generator switch is OFF, or the generator is not connected to the main electrical terminal strip. The GCU will also flash the generator field circuitry if required. N1-N2-NR-TOT The N1-N2-NR-TOT switch installed in the top right section of the instrument panel will connect the N1, N2/NR and TOT gauges directly to the battery in the event of a complete electrical power failure in the helicopter.


CAUTION AND WARNING SYSTEMS

Description - Caution and Warning Systems The caution system consists of a caution panel, two annunciator/switches (one on the 480), a test/dim switch, and 14 input circuits. The warning system consists of 3 individual warning lights and their associated input circuits. The caution system is used to provide a visual indication that a fault condition has occurred. The caution panel, located in the instrument panel, has 15 individual worded segments which when illuminated identify specific fault conditions. When a fault occurs, the associated segment on the caution panel illuminates and flashes at a 2 Hz rate, and the MASTER CAUTION annunciator/switches, located on the left and right side of 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 reset to a steady (on) condition. As each fault condition occurs, it is indicated by the same sequence of events as described above. Only a new fault will flash until it is acknowledged. The MASTER CAUTION annunciator/switches will only be illuminated by faults associated with the caution panel; the warning system will not activate the MASTER CAUTION annunciator/switches.


SEGMENT COLOR DESCRIPTION OF FAULT

ENG CHIP AMBER Engine scavenge oil has ferrous metal fragments MAIN XMSN CHIP AMBER

Main transmission chip detector has detected ferrous metal fragments

TAIL CHIP AMBER Tail rotor gearbox chip detector has detected ferrous metal fragments

ENG OIL TEMP AMBER Engine oil temperature is above 107 degrees C MAIN XMSN HOT AMBER

Main transmission oil temperature is above 107 degrees C

DRIVE BRG HOT AMBER

Either the fwd or aft lower pulley bearings are above 120 degrees C

ENG OIL PRESS AMBER

Engine N1 RPM is above 78.5% and engine oil pressure is below 90 psi

ENG INLET AIR AMBER

Engine inlet swirl tube particle separator partially 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 LOW AMBER Less than 5 gallons remaining

ENG DEICE GREEN Engine anti-ice has been activated

SPARE AMBER Unused segment


Note: An optional external fuel filter can be installed on the TH-28 (S/N, 3007 and subsequent) 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 bypass switch incorporated in the filter assembly. The segment for the impending bypass will be labeled A/F FILTER. Note: The BATT HOT and BATT TEMP segments of the caution panel will only be functional if an optional Nicad battery is installed.

Note: The ENG DEICE segment of the caution panel will not flash or cause the MASTER CAUTION annunciator/switch to illuminate or flash. The SPARE segment of the caution panel will not cause the MASTER CAUTION annunciator/switch to illuminate or flash. Functional Test: Caution Panel Place the caution panel test/dim switch in the TEST position. Check that the MASTER CAUTION annunciator/switch is illuminated and flashing and that all the caution panel segment 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 illuminated for approximately 5 seconds and then extinguish due to programmed continuity sensors in each detector circuit. Reset the MASTER CAUTION annunciator/switches by pressing in on the annunciator/switch. Check that the MASTER CAUTION annunciator/switches extinguish and the illuminated caution panel segments are in a steady bright condition.


Warning Systems The warning system consists of three independent red warning lights located at the top of the instrument panel. When each light is activated, it comes on steady and full bright with no dimming capability. These lights are for conditions that require immediate action.

LIGHTS COLOR DESCRIPTION OF FAULT

ROTOR RPM

RED

Main rotor RPM below 334 RPM

ENGINE OUT

RED

Engine N1 below 58%

FIRE

RED

The fire detection system has detected either a fire or an extreme overheat condition in either the upper or lower engine compartment.

ROTOR RPM

ENGINE OUT

FIRE


Functional Tests - Warning System Turn on the master switch: Check that the ROTOR RPM and ENGINE OUT warning lights are illuminated and the associated audio horns are not activated. Place the place the caution panel test/dim switch in the TEST position and check that the 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 warning light extinguishes when the N1 passes through 58%. Increase the power to bring the NR up to 334 rpm and check that the ROTOR RPM warning light extinguishes at 334 1 rpm. Place the caution panel test/dim switch in the TEST position and check that the ROTOR RPM, ENGINE OUT, and FIRE warning lights illuminate. Release the switch. Check the operation of the ROTOR RPM and ENGINE OUT warning lights during the shutdown procedure


Indicator Caution Activation

TORQUE

NO

Exceeds 72 PSI for 1 second

N1 TACH NO Exceeds 105%

TOT NO

Exceeds 10 seconds between 810 927C during start. Exceeds 810C for 5 seconds during normal operations.

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 YES Less than 50 PSI for 5 seconds Exceeds 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 oil pressure. ** The indicator panel is independent of the FUEL LOW segment in the caution panel.


INSTRUMENTS Dual Tach The N2 (power turbine) RPM is generated by a tach generator mounted of the left side of the engine accessory gearbox and the Rotor RPM is generated by a magnetic sensor installed in the forward section of the MRGB. The rotor and power turbine tachometer (dual tach) is a digital system powered by the aircraft 28-volt electrical system.


Rotor Limitations Minimum Transient Rotor Speed

The minimum allowable transient rotor speed following engine failure or sudden power reduction for practice forced landing is 300 RPM. This is a transient limit and positive corrective action (lowering the collective) must be taken immediately by the pilot to regain 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 N2 See 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


Early 480B Beginning 2007 Engine Instruments

Description Helicopters built during 2005, 20006 and 2007 are fitted with instrument gauges supplied by Horizon/Ultra which incorporate a diagnostic check which is performed when power is first switched on. The indicator light will start out red, and the needle will swing to the full right position before settling at the correct reading. When the diagnosis program is completed, the indicator lamp will also indicate correctly, red or green, corresponding to the gauge reading. Helicopters delivered after 2007 are supplied with the Ahlers gauges which have the indicator lamp but which do not perform the needle swing function test procedure.


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 signal from an engine mounted pressure transducer to indicate engine power. The indicator is powered by the aircraft electrical system through the TORQUE, or TRQ circuit breaker. The Horizon/Ultra gauges microprocessor performs a power on self-test when power initialized and monitor self reasonableness. The Ahlers gauges do not incorporate the self-test function. Both types of torque gauge incorporate a red indicator light that will illuminate when 72 PSI is exceeded for 1 second. (480B) Oil Pressure/Oil Temperature

The engine oil temperature and pressure indicator is a microprocessor based dual indicator which uses a temperature bulb located in the engine oil reservoir for engine oil temperature indications and a pressure transducer connected to the engine oil pressure line on the engine. The microprocessor in the indicator will illuminate the red indicator lights when the engine oil temperature exceeds 107C for 5 seconds, the oil pressure is less than 50 PSI for 5 seconds or the oil pressure exceeds 130 PSI for 5 seconds.


TOT

The turbine outlet temperature (TOT) indicator is a microprocessor based instrument that uses DC voltage through the TOT thermocouple harness to indicate the turbine outlet temperature in degrees Celsius. The indicator receives and averages temperature indications from four thermocouples mounted radially around the engine between 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 the TOT 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 be driven directly from the aircraft battery by selecting BAT on the N1-N2-NR-TOT switch. The microprocessor in the indicator will illuminate the red indicator light when the maximum TOT exceeds 10 seconds between 810C and 927C during the start, or if the TOT 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


Transmission Oil Temperature

The transmission oil temperature indicator is a microprocessor based instrument that displays transmission oil temperature in degrees Celsius by means of an electrical resistance type temperature bulb which is located on the left front bottom of the main rotor transmission. The microprocessor will illuminate the red indicator light in the instrument 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 that uses AC voltage produced by the right side tach generator to indicate the N1 turbine speed in terms of percent RPM. The microprocessor in the indicator will light the red light when the N1 speed exceeds 105%. The 480B systems are active which requires that they be powered by the aircraft electrical system. In the event of a main electrical buss failure, the instrument can be driven directly from the aircraft battery by selecting BAT on the N1-N2-NR-TOT switch.

Fuel Quantity

The fuel quantity system is a capacitance type quantity indicating system and consists of a fuel quantity indicator, a signal conditioner, and a fuel quantity probe. The instrument is a microprocessor based unit that compensates for the non-linear shape of the fuel tank to maintain accuracy between the full fuel and the empty fuel quantity readings. The red light in the indicator will light if the fuel quantity indicates less than 35 lbs for 5 seconds.

The right fuel tank has a float switch that activates the low fuel light on the annunciator panel 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.


Hour Meter The hour meter, located in the left side of the center pedestal records helicopter flight time. It is activated when the engine is running through an oil pressure switch when the collective is raised and therefore records helicopter time off the ground. There is a second optional hour meter available that records time anytime that the engine is running through the oil pressure switch. Start Counter The Rolls Royce helicopter engines are start limited (cycle limited). The hot section of the 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 each time that the starter button is activated to track engine cycles. Any time that the engine is motored without the intention of making an engine start, the IGN EXCITE circuit 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 is connected to the airspeed, altimeter, and VSI.

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

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


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 recommended that autorotation speeds be limited to 85 KIAS or Vne, whichever is lower.


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 plates separated by an aluminum cylindrical spacer. Through bolts hold these items together along with steel spline adapters.

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

Dampers

480B Hub


The rotor blades are secured to each universal block on the hub through a forged aluminum grip which is in turn secured to a steel

spindle assembly through a retention nut or an optional tension-torsion strap assembly and supporting bearings. On some of the early 480s, centrifugal blade loads are carried by Lamiflex elastomeric bearing assemblies. On most of the 480 series aircraft and the 480Bs, tension-torsion strap assemblies mounted between the blade grip and the spindle are used.

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

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


The main rotor blades are of hollow construction consisting of an extruded leading edge spar, with a 7-degree twist, to which is bonded upper and lower aluminum skins. The blade root is composed of a bonded doubler assembly. A cap is bonded to the tip of each blade in which there are provisions for spanwise and cordwise balance weights. Two tracking tabs are riveted to the trailing edge of each blade.


Tail Rotor Assembly The tail rotor assembly is a two bladed, wide cord, teetering, delta hinged rotor assembly.

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

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


Tail Rotor Guard A tubular aluminum tail rotor guard is installed on the aft end of the tailcone. It acts as a warning to the pilot upon an inadvertent tail-low landing and aids in protecting the tail rotor 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 rotor strike during a hard landing or auto-rotation flare.


FUEL SYSTEM

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


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 the cockpit.

Each fuel cell is equipped with sump drains plus the system is equipped with a low point drain 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 cell is filled by crossfeeding action during refueling.

On the 480 series aircraft the fuel cells are filled with blue/green open cell foam. The purpose of this foam is to help maintain the shape of the cell as the fuel is used and to prevent the fuel from sloshing around in the tank. If the fuel nozzle contacts the plastic tube during servicing, it can push the plastic tube down into the foam or it can dislodge some of the blue / green foam inside the tank which will then show up in the fuel sumps for a short while.

Blue Foam

Plastic tube to measure fuel quantity


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 are disturbed.

To physically measure the fuel level in the tanks. A dip stick can be inserted through the fuel filler opening and into the clear plastic tube that penetrates into the foam. The top reference line on the dipstick must be held against the filler opening. The fuel level can be 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 must be 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 into the fuel flow totalizer.

IMPORTANT!

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


POWER TRAIN

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


Drive System Lower Pulley Drive System The lower pulley drive system consists of the lower pulley drive shaft and hub, lower pulley, "H"- strut, and flex pack couplings. The lower pulley drive shaft, located in the hollow center of the lower pulley assembly, is connected to the power output shaft and the lower pulley assembly with couplings. The lower pulley shaft is also used to drive the oil cooler blower fan by means of a hub attached 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 used to laterally align the lower pulley to the engine.


Thermocouples are installed in the lower pulley bearing housings to provide temperature 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 to the pinion bearing support truss and the main rotor transmission at the upper end. The flex 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 the power output shaft and the lower pulley drive shaft.

IMPORTANT!

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

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


Description - Overrunning Clutch The overrunning clutch is installed on the front side of the engine accessory gearbox. The outer housing of the overrunning clutch forms the driving portion of the clutch and is driven by the engine power output shaft.

In the driving direction, the sprags engage and connect the outer housing to an inner drive housing which transmits the engine torque to a splined drive shaft that passes through the center of the engine power output shaft and the outer housing of the clutch assembly to the rear of the engine accessory gearbox where it is coupled to the lower 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 will disengage thus disconnecting the engine from the rotor drive system. The overrunning clutch is a sealed unit and contains its own lubrication separate from the engine.

Overrunning Clutch


Oil level Sight Glass

There is a sight glass installed on the forward end of the overrunning clutch (SDB

T-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 the clutch requires service earlier than each 20 hours of use, Enstrom recommends investigating the engine for double-lip seal leakage.

Sight Glass


FLIGHT CONTROLS Note: the pilot controls are normally on the left side in a 480 and on the right side in a TH28. The flight controls include three primary systems: the collective, cyclic, and anti-torque or 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.


Collective Control System The collective control system is comprised of dual collective controls mechanically interconnected and linked to the main rotor swashplate through a series of push-pull tubes, a torque tube, a bellcrank, and a collective walking beam at the base of the main rotor transmission.

Both collective controls have interconnected twist grip throttles and a switch box mounted forward of the throttles. The pilots collective incorporates the landing light controls, 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 control switch, and a N2 power turbine governor increase/decrease (GOV INCR/DECR) switch.


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

Idle Detent Button

Starter Button

Throttle

Landing Light Angle Adjustment

Governor Beeper

Landing Light

Switch


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

The control may be positioned in any intermediate position for any desired level of friction. The collective friction system is designed so that positive locking of the collective controls cannot be obtained at the maximum friction point. Safety of flight considerations require that the pilot be able to instantly overcome the established friction 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 of the collective bellcrank in the engine compartment. The collective trim system consists of a spring capsule, bracketry, and an adjusting link.


Cyclic Control System Description Note: The cyclic control system is a fully mechanical control system which is linked to the swashplate through a series of interconnected push-pull tubes, a torque tube, and bellcranks. Both longitudinal and lateral control systems are totally independent with no intermixing before the individual inputs reach the swashplate.

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


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 center of the grip, used to control the four way cyclic trim system. The cyclic trim system maintains the position of the cyclic control stick and reduces rotor feedback to zero. Cyclic Grip

Trim Switch

Freq. Transfer

Radio & Intercom Transmit Switch

Transponder Ident.


Cyclic Trim

The system consists of a cyclic trim switch located at the top of each cyclic grip, a pair of electrically operated jack screw actuators that vary spring tension produced by the longitudinal and lateral trim units, and a pair of trim switch units which reverse the direction 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 an electrically 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 engine compartment. When a trim switch is moved off of center to any one of the four trim directions, power applied through the TRIM circuit breaker energizes one of the trim motors to apply trim spring force in the desired direction. By momentarily moving the switch, very small trim increments may be obtained. Trim force cannot be applied in two directions simultaneously; when both longitudinal and lateral trim corrections are desired, it is necessary to apply first one and then the other. The cyclic trim system does not limit travel of the cyclic control; the pilot may override the trim forces at any time.


POWER PLANT Description The Enstrom TH-28 and 480 series helicopters utilize an Allison designed, Rolls Royce built 250-C20W reverse-flow freepower-turbine turbo-shaft engine. The engine maximum rating is 420 shaft horsepower. In the 480B, the engine is rated at 277 shaft horsepower continuous and 305 shp for 5 minutes. This is a transmission based limitation. In the TH28 and early 480 aircraft, the engine is derated to 256 shaft horsepower maximum continuous power and 285 shp for 5-minutes.


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 engine compartment from the transmission compartment. Beginning on helicopters delivered in 2008 and on helicopters that have the high temperature modification installed, the baffle 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 compartment which 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 helicopters delivered in 2008, additional ventilation holes have been added in the lower engine cowl panels and the bottom engine cover access hole has been opened up to assist in engine cooling.


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


Air Induction System The upper plenum chamber is equipped with dual full flow swirl tube inertial type particle separators. Engine inlet air is directed into the upper plenum chamber through a 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 through one of two bleed air driven venturi-type ejectors that exit at the aft side of the upper plenum. 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 bleed air shutoff valve can be opened by placing the SCAV AIR control handle in the ON position. The inlet air moves from the upper plenum chamber to the lower plenum chamber 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 engine may not deliver rated power when they are engaged. The lower plenum chamber is mounted directly behind the engine and is connected to the engine by a bell mouth inlet. The inlet is attached to the engine and a foam-rubber gasket 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 any moisture that might happen to accumulate during operation or while the engine is not running. A fitting is also installed in the left side for connection to the ENG INLET 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 the right side of the engine compartment and is accessible through the right side engine access panel.


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

The blower assembly is located aft of the lower drive pulley. The assembly consists of a 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 exhaust ducts.


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, ASTM D-1655 Jet B, Jet A, and A1 (See note 1 below) JP-1, Diesel #1, or Arctic Diesel DF-A (VV-F-800B) conforming to ASTM D-1655, Jet A or Jet A1

With anti ice additive conforming to MIL-I-27886

Emergency

MIL-G-5572E AVGAS (without TCP)

All Grades, Maximum 6 hours operation per overhaul period. With anti ice additive

Cold Weather

MIL-T-5624 JP-4 ASTM D-1655 Jet B Avgas-Jet fuel mixture

(See note 2 below)


Note 1: All fuels used in the TH-28/480 shall contain Anti-icing and Biocidal Additive conforming to MIL-I-27686. The additive provides anti-icing protection and also functions as a biocide to microbial growths in helicopter fuel systems. Icing inhibitor conforming to MIL-I-27686 shall be added to all commercial fuel, not already containing an icing inhibitor, during refueling operations, regardless of the ambient temperatures. Refueling operations shall be accomplished in accordance with accepted commercial procedures. Commercial product "PRIST" conforms to MIL-I-27686. Note 2: The AVGAS-jet fuel mixture is an alternate fuel which may be used if starting problems 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 jet fuel. 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 lead content. (The 2.0 ml/gal max. lead content grade 100/130 AVGAS is known as 100L in European areas). The commercial jet fuel may be kerosene; JP-5 or commercial Jet A conforming 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 to exactly follow the additive mixing procedure during refueling can result in incorrect additive concentrations, fuel system contamination, and possible engine stoppage


Engine Lubrication Oils Note: Rolls Royce recommends not mixing oils within a series unless absolutely necessary. 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 Stability (HTS)


Royal Lubricants Company Aeroshell / Royco Turbine Oil 254

Shell International Petroleum Co. Aeroshell Turbine Oil 560

Mobil Oil Mobil Jet Oil 254


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 IS PERMITTED WITHOUT A TIME PENALTY. SEE THE ROLLS ROYCE MAINTENANCE MANUAL FOR SPECIFIC INFORMATION.


PERFORMANCE DATA The following general conditions are applicable to the performance data: 1. All airframe and engine controls are assumed to be rigged within allowable

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

continuous. 3. No two aircraft are exactly the same; however variations are considered to be

small and cannot be individually accounted for. 4. The data presented presumes that all instruments and systems have been

properly maintained, are in proper working condition and are calibrated. Torque Available The torque available chart shows the effects of altitude and temperature on engine power available. The primary use of the chart is to provide the pilot information on the maximum power available either as a function of the helicopter limits or the flight 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 N1 may 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 air usage on performance. By using maximum torque available chart, and the hover chart, the pilot can determine the power margin for an intended operation.


Hover Performance


Cruise Performance

The cruse performance charts show the torque pressure and fuel flow for level flight at various pressure altitudes, airspeeds and gross weights. The two charts on the page are not connected together. To obtain the expected fuel flow, knowing the torque, the left chart must be used to obtain the airspeed. This airspeed is then entered into the chart on the right and the fuel flow can be obtained. There are charts for sea level, 3000, 6000, and 9000 feet pressure altitude.


Power Assurance Check The power assurance chart provides the pilot with a method to assure that the installed engine will provide the expected power required to achieve the performance presented in the RFM and also to monitor the engine performance over time. The pilot should come to a stabilized hover, record the pressure altitude, OAT, Torque and 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 the helicopter 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 a stable hover. It may be necessary to hover for at least 2 minutes to allow the engine to completely stabilize before taking engine performance numbers

During a hover, The TOT, Torque, Pressure Altitude (29.92), and Temperature are recorded. These parameters can be plotted on the Power Assurance chart to see that engine is performing within the recommended parameters. If the actual TOT reading is higher than the calculated TOT of the Engine Power Assurance Chart in the TH-28/480 Operators Manual and pilot error has been ruled out, the engine will not meet the performance figures called out in the Performance Data section of the TH-28/480 Operators Manual. Turbine engines normally degrade or lose power through engine operation (Refer to the Trend Check Procedure section of the Rolls Royce 250-C20 Operation and Maintenance Manual for further information). A gradual increase in TOT that is above the normal engine degradation trend line is normally caused by a dirty compressor. Performing a compressor wash will usually return the TOT indications to the normal engine degradation trend line. If the actual TOT readings significantly increase or decrease over a short period of time, a potential problem may exist with the engine or an airframe system.


Height Velocity Diagram


Engine

Engine Power Controls The engine power control system is a mechanical linkage/ cable system, actuated by a twist-grip on the collective sticks, which provides manual control of the power lever on the 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 below the flight idle position causing an accidental engine shutdown. The release does not have to be pushed for engine start or run-up but does have to be pushed for engine shutdown. Rolling the throttle back past the stop has the effect of shutting off the fuel flow to the engine, thereby shutting it down.


An electrically operated linear actuator operates a lever connected to the power turbine governor. 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 N2 RPM as the engine load fluctuates with changes in the main rotor pitch. It consists of a series of linkages and bellcranks that connect the collective control to the governor on the engine an accumulator that is installed in the engine PC system. The linkage is designed in such a way as to cause the governor to anticipate the requirement for increases or decreases in power requirements to help prevent the rotor speed from lagging behind the pilot inputs.

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

Understanding the operation of and the relationship between the governor and the fuel control will help prevent the pilot from getting into some situations which can lead to accidents. The governor reads the N2 RPM (output-shaft speed of the engine) and schedules the fuel to the fuel control to keep the N2 speed, (also the rotor speed) constant. The governor is sensing RPM and will always attempt to maintain this RPM. Because turbine 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 compensates for changes in the N2 (Rotor Speed), and that both changes in collective and tail rotor inputs will affect N2. However, the droop compensator system is only connected to the collective and not to the tail rotor, so there are possible situations where large left tail rotor inputs can spike the torque as the governor tries to react to, and compensate for changes in the Main Rotor RPM. The engine will not respond to a tail rotor input until the 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 governor cannot react to the pilot inputs with the resulting loss of the aircraft.


During autorotation, and especially in the flare, the rotor RPM will be on the high end of the 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 the rotor. So at the maximum point of the flare, just as the pilot needs to roll the throttle back on for the recovery, the governor is still trying to slow the rotor by cutting back the fuel!

When the pilot rolls the throttle back on, the N1 will be slow, and due to the lag in the turbine 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 this only aggravates a bad situation! The addition of left pedal drags the rotor RPM down even more. When the governor does react to the rotor RPM which is now low due to the autorotation recovery and the addition of left tail rotor pedal, the governor is behind and the result is a spike in torque as it tries to catch up.

Enstrom Helicopter recommends that the throttle not be rolled off during practice autorotations. The pilot should begin a power recovery by increasing collective pitch slightly during the flare to give the engine a jump start on delivering the power that will be 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 will dramatically reduce the torque necessary for the recovery


Starting Procedures

IMPORTANT!

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

Hot Starts

For the most part, hot starts are a result of not enough air passing through the engine when fuel is added. Usually this is a result of either a low battery or improper starting 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 so there are two consequences of low battery.

1. The engine does not achieve enough RPM before light-off and the TOT

exceeds the manufactures limits due to insufficient air flow. 2. The engine hangs in the 30 % range and wont carry through. (This can also

be the result of an overheated starter) (Also known as a stagnated start)


Starting procedures There are three or four common procedures in general use for starting Rolls Royce 250 C-20 B and W series, engines. Any of these procedures that the pilot is comfortable with can be used for starting the Enstrom 480 helicopters. The pilot needs to be aware that due to the design of the 480 collective throttle and detent system, if the throttle is rolled off and held against the idle stop detent position, the detent button cannot be depressed. Pressure against the detent position must be relieved before the detent button can be depressed. The following procedure is an alternate method of starting the helicopter that can be used 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 within acceptable tolerances. NOTE

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

The reason for not opening the throttle to the detent during the start is that in 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 lock button and abort the start. Also, not passing the detent position gives the instructor in the co-pilot seat the ability to abort the start in case the student does not recognize an impending hot start.

If the throttle is advanced past the idle position before the engine reaches 58%, there will be a surge in torque as the engine passes 58% that might be uncontrollable. This is because the governor is sensing that the operating RPM is less than the throttle position is calling for and it will schedule full fuel to catch up. The C-20 engine will develop maximum torque at situations where the RPM is low and increasing.


Recommended starting procedure follows

1. Master on 2. 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 throttle 6. (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 and start again.)

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

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

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

10. Monitor N1 and TOT when turning the generator switch on. If N1 decays below 60% or TOT approaches 750C, turn the generator off and increase N1 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 on the 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 peak TOT is reached, push and hold it down until after peak has been passed.


Recommended procedure for starting the engine when warm or hot

1 Master on 2 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 is at 150C.

7 After the engine starts, stop opening the throttle. 8 After the TOT peaks, tweak the throttle open just a bit more to prevent

inadvertent shut-down. 9 When the N1 reaches 58% carefully advance the throttle until the detent button

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

generator. 11 Monitor N1 and TOT when turning the generator switch on. If N1 decays below

60% or TOT approaches 750C, turn the generator off and increase N1 speed with the throttle to 70% and reset the generator switch to on.

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

13 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 peak TOT is

reached, push and hold it down until after peak has been passed.


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. In this case, waiting five minutes and reinitiating the start will usually be successful as sitting 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 fuel control and is field adjustable.

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

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

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

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

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

To obtain an optimized start, move the throttle to begin fuel flow as the N1 RPM accelerates through 12 15% N1. Do not wait for N1 RPM to peak out before initiating fuel 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 gas producer turbine life. (N1 wheels and Nozzles) The probes for the TOT system are installed radially around the engine between the gas producer and the power turbine sections and may not reflect the actual temperatures being experienced by the first stage turbine wheel and nozzle assembly. The recommended quick warm starts actually increase cooling air flow in the combustion section to help cool the gas producer turbine


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 it through the start. The earlier (lower RPM) that the start is initiated, the more of the start 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 may not be enough boost from acceleration to carry the start, and the result is that the start may stagnate. This is due to the natural tendency of the battery to deliver lower voltage when 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 and making the start in one event, than in two. It is possible to get a stagnated start with a warm 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 or 20 minutes and use an APU for the start. Initiate the fuel as the N1 passes 12% for maximum boost from the acceleration.


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 the TOT gauge. There is a scale change at 850C on the gauge. 850C is the last mark before 900C.)

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


EMERGENCY PROCEDURES

This section describes the foreseeable helicopter and systems emergencies and presents the procedures to be followed. Emergency procedures are given in checklist form when applicable. Definition of Terms Immediate Emergency Actions. Those actions that must be performed immediately in an emergency procedure are underlined. These immediate emergency actions must be committed to memory. Note: The urgency of certain emergencies requires immediate and instinctive action by the pilot. The most important single consideration is helicopter control. All procedures are 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 the doors will not open, break or kick out the door windows, overhead windows, or windshields as the situation requires.


Engine Failure The indications of an engine failure, either a partial power loss or a complete power 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 engine failure by cross checking all of the engine instruments. Note: The first indication of an engine failure will normally be an uncommanded left yaw of the nose of the helicopter. The engine-out warning horn is activated by the N1 tach-generator. If there is a horn, and NO left yaw, before entering autorotation, verify that the engine is actually not providing power and that the problem is not actually an instrument failure. If the pilot instinctively reacts to a warning horn or light by immediately bottoming the collective, it is recommended that after the glide is established, the pilot try gently raising the collective while monitoring the torque gauge. If the torque reading shows an increase the pilot would be advised 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 or it 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 a favorable landing area; however, under these conditions the pilot should always be prepared for a complete power failure at any time.


After an engine failure in flight, an engine restart may be attempted if time and altitude permit. 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 of descent, potential landing areas, and crew assistance available. Under ideal conditions, approximately 30-45 seconds is required to regain powered flight from the time the attempted start is begun if the start is commenced with an engine that is not windmilling. If the engine start button is depressed immediately after autorotation has been established, powered flight can usually be resumed within a matter 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 N1 below 15%. Note: Unless there is a reason to believe that the engine has failed due to some obvious mechanical failure, always attempt relight immediately after entering 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 relight attempt, 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 failure and attempted relight has not exceeded 5 seconds. There will be a slight surge as the engine comes back on line but it will be well controlled and will not damage the engine or 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 is reestablished, perform a power on approach and landing without delay if the engine was not intentionally shutdown.


Engine Restart - During Flight 1. Establish Autorotative glide 2. Attempt Start:

Throttle Closed Starter Button Depress Throttle - Idle (N1 15% or greater) TOT and N1 Monitor Starter Button - Release (at 58% N1) Throttle - Advance to Full ON (N2/NR needles rejoined) Powered Flight - Resume

3. Land Immediately - After the engine is started and powered flight is reestablished, perform a power on approach and landing without delay if the engine was not intentionally shutdown. Autorotation Rotor RPM recovery becomes very slow below 300 rotor RPM and the rotor RPM cannot be recovered below 240 rotor RPM. Never allow the rotor RPM to fall below 300 RPM in flight. In practice autorotations, if the rotor RPM falls below 300 rotor RPM, and the aircraft is not very close to touchdown (less than 1 foot), IMMEDIATELY bring the engine back on line to recover from the maneuver. NOTE: Normally this situation will not occur if autorotation practice is done with the throttle at the full open position. Minimum rate of descent and maximum glide are obtained at minimum rotor RPM. See Chapter 4, Performance, in the flight manual, for appropriate airspeeds and further information. Minimum Transient Rotor RPM - The minimum transient rotor speed can and may be as low as 300 rotor RPM during the initial response time following engine failure and during the initial recovery control inputs.


Note: Although operation below 334 RPM may be unavoidable during the initial stages of the autorotation, it should be minimized and immediate corrective action is required. Minimum Rotor RPM - The minimum steady state rotor speed is 334 rotor RPM. This is the minimum allowable sustained rotor speed in steady flight, power off.

Minimum Flare Airspeed - The minimum flare airspeed is 50 KIAS. This is the minimum airspeed that will allow effective tradeoff of forward airspeed for rotor RPM in the flare prior to touchdown and efficient energy conversion to arrest the rate of descent. A 25 rotor RPM increase can be achieved from a flare at 50 KIAS. The Height-Velocity Diagram is published to assist the pilot in defining the limiting combinations of height and airspeed below which it will be impossible to maneuver the helicopter to intercept the autorotation profile prior to touchdown. The most likely outcome for an engine failure within the boundaries of the H-V diagram is a crash landing.

Note: The Height-Velocity Diagram has been established with the engine shut down and can therefore be relied upon as being accurate for the actual engine failure case.

Rotor RPM Low Rotor RPM is below normal. Lower collective as required to regain RPM and ensure that the throttle is FULL OPEN.


Caution Lights

Engine Oil - Low Pressure/High Temperature - Eng Oil Press Caution Light Illuminated

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Enstrom 480B Training Manual

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