SISTEMAS DE MOTORES DE COMBUSTION INTERNA PRATT & WHITNEY JT8D-17

Maria Nataly López Maya Edgar Joe Zavaleta Escobedo Grupo: 6AM1

Contents Introduction ………………………………………………………………………………..3 Power plant - description and operation………………………….…….….……………4 Engine cowling - description and operation 71………………………………………...6 Engine fuel deicing system - description and operation………………………………8 Fuel flow indicating system - description and operation……………………………11 Ignition system …………………………………………………………………………..14 Engine anti-icing - description and operation………………………………………...19 Engine exausting system………………………………………………………………..21 Thrust reverser system……………………………………………………………..…...23 Thrust reverser - hydraulic actuators and sync shafts……………………………….25 Engine starting…………………………………………………………………………...26 Generator cooling - description and operation………………………………………..28 Engine control system - description and operation…………………………..………30 Engine indicating – description…………………………………………………………32 Thrust reverser - description and operation…………………………………………..34 Engine oil system - description and operation…………………………………..……36 Pneumatic starting system - description and operation……………………………...40 Bibliography………………………………………………………………………………42

Introduction The JT8D engine is an axial flow , dual wheel and fully shaped product. Its design began in April 1960 . The JT8D engines - 7 to 17 are used in the following aircraft : B 727 and 737 , McDonnell Douglas DC -9 , Aerospatiale Super Caravelle and Saab Virgin The JT8D engine has proven itself to be a highly durable and reliable engine, having completed more than 673 million dependable flying hours since entering service. Once deemed the workhorse of the industry, more than 14,750 JT8D engines have flown. Today, there are 2,400 engines still in use.The eight models that make up the JT8D family cover a thrust range from 14,000 to 17,000 pounds. The JT8D-200 series, which entered service in 1980, offers 18,500 to 21,700 pounds of thrust, and is the exclusive power for the popular MD-80 series aircraft. The JT8D-200 builds on the family's excellent reliability and low maintenance costs while meeting noise and emissions regulations.Pratt & Whitney has developed a new low-emissions combustion system, or E-Kit, that is FAR 25-certified to ensure the JT8D-200 engine stays current with environmental regulations. The E-Kit reduces JT8D-200 engine NOx emissions by 25 percent and exceeds all ICAO standards for new production engines..

Ilustración 1

Main components:

      

Low pressure compressor ( six stages) High pressure compressor (seven stages) Combustion section (nine combustion chambers) High pressure turbine (one stage) Low pressure turbine (nine stages) Scape pipes Gear box

All ed by nine bearings POWER PLANT - DESCRIPTION AND OPERATION The 727 airplane is powered by three JT8D turbofan engines (figure 1) mounted on the aft fuselage area. Two engines are strut mounted, in conventional nacelles, one on each side of the fuselage. A center engine is mounted aft of the fuselage structure. Air is ducted to this engine through an inlet forward of the vertical fin just above the fuselage. Each engine is secured to the engine mount fittings at three points. Two cone bolts, attached to a double flange at the fan discharge intermediate case, and one cone bolt attached to a double flange at the fan discharge turbine exhaust outer duct, secure the engines to the forward and rear mount fittings. Access to the side engine exterior components is provided by hinged removable upper and lower cowl s which may be opened from either side. Center engine access is made through four hinged removable side cowl s. Access to the center engine inlet and nose dome is gained through an access above the aft airstairs in the center engine inlet duct. B. The major accessories fitted to each engine include a constant speed drive unit, a pneumatic starter, and N1 and N2 tachometer generators. A generator is mounted on the constant speed drive unit. Hydraulic pumps are installed on engines No. 1 and 2. Fire detection and fire extinguishing systems are provided in each engine area. C. The engine starting system provides a means of rotating the N2 compressor to establish a flow of air through the engine. Rotation of the N2 compressor also drives the engine fuel pump and fuel control to meter fuel, under pressure, to the combustion chamber. An ignition system provides a high voltage discharge for ignition of the fuel/air mixture. Each engine is fitted with a self-contained oil system to provide cooling and lubrication of engine gears and bearings. Transmitting devices, installed on each engine, actuate warning lights and indicators in the control cabin to provide indications of engine performance.

Ilustración 2

D. Air is bled from the low pressure and high pressure compressors and the fan discharge duct to operate various airplane systems. A thrust reverser, attached to the exhaust section of each engine provides a means of retarding the forward speed of the airplane after touchdown. E. A four compartment drain tank is mounted on the under side of each engine. Leakage fluids from the seals and drive pads of the engine accessories are drained, through tubing, to the tank. On engines with the fuel pressurizing and dump valve drain port not plugged, fuel discharged from the valve also drains through tubing to the tank. A drain line is installed between the oil tank scupper and the drain tank, to collect oil spilled during servicing. In flight, air flowing past a drain mast, on the engine cowling, creates a negative pressure which sucks the fluids overboard through a discharge manifold. Jiffy drains are provided on the under side of each tank compartment to permit draining the tank while the airplane is on the ground.

ENGINE COWLING - DESCRIPTION AND OPERATION 71 A. The engines are covered by cowl s to provide a smooth airflow over the engine and to protect exterior engine components from damage. The cowl s for the side engines include a nose cowl, upper and lower hinged removable cowl s, and a fixed cowl . The center engine is covered by four hinged

removable side cowl s. A nose dome is also fitted on each engine. Titanium and steel fireshields are bonded to the inside of the outer skin in areas in which a fire may occur. The cowl s are designed to be drip-free. Drain holes and drain tubes in the lower surfaces of the cowl s collect any engine leakage fluids in a drain manifold to be exhausted overboard in flight. 

Center Engine Cowl s

The center engine cowl s form part of the aft fuselage contour. The forward left and right s fair with aft fuselage skin. The aft s fair with the thrust reverser shroud ring and the thrust reverser actuator fairing. Each is hinged at the top by fittings on the center engine rib. Three hook latch fasteners opposite s at the underside of the engine. Two safety latches at the top of each prevent an open cowl from being removed until the latches are depressed. The safety latches automatically trip into the locked position when the s are closed. An indicator on each safety latch shows when the latch is in the locked or unlocked position. "Hold open" rods, installed in each , allow the s to be propped in the open position. Safety pins are provided to lock the rods in either the open or stowed position. B. The left forward cowl is provided with an access door to facilitate servicing the engine oil tank. On preferred cowl installations an access door is also provided in the left forward cowl to manually operate the starter shutoff valve. Exhaust ports are provided on the left aft cowl to accommodate engine gearcase breather air and on the left forward cowl , to exhaust fuel heater air and constant speed drive oil cooler air. An exhaust port on the right aft cowl is provided to exhaust generator cooling air. An overboard drain mast, installed on the right aft cowl , evacuates the engine fluid drain tank during flight. Three blowout doors, one on the right forward cowl and two on the right aft cowl , relieve excessive pressure which could develop within the cowling during a fire. 

Side Engine Removable Cowl s

The side engine upper and lower cowl s fair with the engine nose cowl, with a fixed cowl attached to the inboard side of the engine, and with the thrust reverser shroud ring Six hook Latch fasteners attach the upper and lower cowl s to the fixed cowl . Six other hook latch fasteners the upper and lower cowl s together. Three forward and three aft pin latches ensure positive attachment of the cowls and serve as hinges permitting each to be opened from either side. "Hold open" rods for each are stowed on the engine, the nose cowl aft structure, and on each

. Ilustración 3

B. Openings are provided in the lower cowl for the engine gearbox breather, the constant speed drive cooling air exhaust, the generator cooling air exhaust, an overboard drain mast, and an oil tank access . On preferred installation, a starter override and CSD access door is also fitted to the lower cowl . The upper cowl is fitted with a blowout door. 

Side Engine Nose Cowl

The nose cowl is bolted to the forward flange of the side engine inlet case. It is shaped to obtain a smooth airflow over the engine and to provide an optimum airflow to the engine compressor inlet (Fig. 3). An anti-icing air inlet is located on the rear face of the nose cowl at approximately the 9 o’clock position. Three pin latches, two on the inboard side and one outboard, the nose cowl to the upper and lower cowl s. A constant speed drive cooling air scoop is located on the underside of the nose cowl. The nose cowl is anti-iced by engine bleed air (Ref Chapter 20, Ice and Rain Protection). 

Side Engine Fixed Cowl

A four piece fixed cowl , mounted on the inboard side of each side engine, fairs with the upper and lower cowl s. The fixed cowl s the side engine upper and lower cowl s. 

Nose Dome

A nose dome is mounted on each engine to provide a smooth airflow over the front accessory drive housing. The nose dome covers the N1 tachometer generator. An inlet pressure sensing prove (Pt2) is mounted in the center of the nose dome.

Ilustración 4

ENGINE FUEL DEICING SYSTEM - DESCRIPTION AND OPERATION A. Engine fuel deicing system detects the presence of ice in the fuel and provides controlled heating of fuel to melt the ice. Fuel generally contains suspended water droplets when the temperature of the fuel falls below the freezing point of water. These suspended water droplets freeze and form ice. Ice eventually clogs the engine main fuel filter and restricts normal fuel flow to the engine. B. The engine fuel deicing system for each engine consists of a fuel deicing heater, fuel deicing air valve, fuel deice control switch, valve-in-transit light, fuel filter pressure switch, fuel icing warning light, and the necessary tubing. The fuel heater, fuel filter pressure switch, and fuel deicing air valve are mounted on the engine. The fuel deice control switch, valve-in-transit light and fuel icing warning light are on the third crewman’s lower . C. When ice clogs the engine main fuel filter, pressure differential across the filter builds up, the fuel filter pressure switch closes and the fuel icing warning light illuminates. Turning on the fuel deice control switch permits high pressure compressor air to through the fuel deicing heater air tubes and heat the fuel. The warm fuel es through the fuel filter and melts the ice. When the ice is melted, the fuel filter pressure switch opens and de-energizes the fuel icing warning light, the fuel deicing control switch is then turned OFF. The fuel should be heated intermittently. 

Fuel Deicing Heater

A. The heater is an air-fuel heat exchanger and consists of a housing containing a core composed of air tubes, a series of baffles, and a fuel by valve heater is mounted on the fuel control unit between the boost and main stages of the enginedriven fuel pump.

Ilustración 5

B. The fuel flowing to the main engine fuel filter es through the heater at all times. It is heated, only when the fuel deicing air valve is open and permits high pressure compressor (13th stage) bleed air to through the heater air tubes. To obtain uniform heating of the fuel, it is baffled around the air tubes. In an event the heater becomes clogged, a by valve permits the fuel to by the heater and flow directly through the fuel filter to the engine. 

Fuel Deicing Air Valve

A. The air valve is used to control the high pressure compressor (13th stage) bleed air flow through the fuel deicing heater (Fig. 1). It consists of a butterfly valve and an electrical actuator (motor). The valve has an override handle for manually positioning the valve and this same handle can be used during maintenance to determine valve position. The air valve is operated by a fuel deice control switch, located on the third crewman’s lower . There is one air valve for each engine located at the 2 o’clock position slightly forward of the engine mount ring. 

Fuel Filter Pressure Switch

The pressure switch senses the pressure differential across the engine main fuel filter. The pressure switch is mounted directly on the filter. If the engine fuel filter becomes clogged, pressure differential across the filter builds up. When the

pressure differential reaches 4.4 to 5.8 psi, the pressure switch closes and illuminates the fuel icing warning light on third crewman’s lower . When the pressure differential again decreases, the pressure switch opens and the fuel icing warning light goes off. 

Operation

A. The engine fuel deicing system receives its power from 115-volt ac buses through circuit breakers on circuit breaker . B. When ice is present in the fuel it will eventually clog the engine main fuel filter and cause the pressure differential across the filter to increase. When the pressure differential reaches 4.4 to 5.8 psi, the fuel filter pressure switch will close and illuminate the fuel icing warning light on third crewman’s lower . Placing the fuel deice control switch to ON position opens the fuel deicing air valve. High pressure compressor bleed air (13th stage) es through the valve into the heater air tubes and heats the fuel. To obtain a uniform heating of the fuel, baffles in the heater core direct the fuel around the air tubes in a controlled flow pattern. The heated fuel flows to the fuel filter and gradually melts the ice clogging the filter. When all ice is melted, pressure differential across the filter decreases and the fuel filter pressure switch opens, deenergizing the fuel icing warning light. The hot airflow through the heater is stopped by placing the fuel deice control switch to OFF position. Whenever the fuel deicing air valve is in transition, the valve-intransit light on third crewman’s lower will illuminate. The deicing heater should be used intermittently. C. If the fuel filter becomes clogged, the fuel icing warning light will remain on, and the filter pressure differential will continue to increase until the filter by valve opens. The entire fuel filter system must be cleaned and checked after operating under these conditions. If the fuel deicing heater itself becomes clogged, a by valve permits the fuel to by the heater and flow directly through the fuel filter to the engine.

Ilustración 6

FUEL FLOW INDICATING SYSTEM - DESCRIPTION AND OPERATION A. The function of the fuel flow indicating system is to provide a visual indication in the control cabin of the fuel consumption rate in kilograms per hour for each individual engine. B. This system has three fuel flow transmitters, three fuel flow indicators, and one fuel flow power supply unit. A fuel flow transmitter is located on the forward left side of each engine. All three fuel flow indicators are located on the third crewman’s lower instrument . The fuel flow power supply unit is located on the E5-1 equipment rack on 727-100 Series Airplanes, and is located behind the third crewman's on 727-200 Series Airplanes (Fig. 1). C. When the system is energized the fuel flow power supply unit supplies constant frequency 3-phase power for driving the three fuel flow transmitters. Each transmitter senses the mass rate of fuel flowing to its respective engine and generates an electrical signal proportional to this flow. The electrical signals are transmitted to the fuel flow indicators which show the fuel being consumed in kilograms per hour by their respective engines. D. Each fuel flow transmitter measures the amount of fuel flowing to its respective engine and provides a corresponding signal to its respective indicator. The indicator uses this signal to provide a visual indication of the rate of fuel flow to the engine.



Fuel Flow Power Supply Unit

A. There are two types of fuel flow power supply units. One is electromechanical and the other newer unit has solid state circuitry. Both units furnish a constant frequency power source for the fuel flow transmitters and both units are interchangeable in the airplane. The electromechanical unit consists of a synchronous motor connected to a 3 bar commutator through a reduction gear. When the motor is energized the commutator is driven at a constant speed. As the commutator rotates it chops the dc current to produce a simulated 3-phase , 4-cycle ac current. Filters are installed in the unit to ensure that the output is free of radio noise. The solid state unit produces the same output to the fuel low transmitter by electronic means. 

Fuel Flow Transmitter

A. The fuel flow transmitter consists of a housing containing two identical cylinders placed end-to-end so that the axes of the cylinders coincide. The upstream cylinder is named the impeller and the downstream cylinder is named the turbine. The housing fits closely to the outer diameter of both the impeller and the turbine. Around the periphery of the circular cross section of the cylinders, at a fixed radius from the center, are a number of equally spaced holes which are accurately parallel to the axes of the cylinders. Upstream of the impeller is the impeller motor. Downstream of the turbine a second harmonic transmitter is attached to the turbine with two sets of springs of different torque gradients used to restrain it. This permits a two-slope scale on the indicator allowing greater sensitivity at low flow rates. NOTE: To prevent dry operation damage and to prolong the life of the transmitter, it is recommended that the “FUEL FLOW” circuit breakers be pulled (open) whenever the engine fuel feed line is drained or the airplane is out of service for maintenance. B. The impeller motor drives the impeller at a constant angular velocity, thus as the fuel es through the impeller it is given a velocity at right angles to the direction of flow. This angular velocity constitutes a change in the momentum of the fuel directly proportional to the mass of fuel flow. As the fuel es through the turbine, the angular component of the momentum is removed, this imparts to the turbine a torque directly proportional to the mass rate of fuel flow. This torque rotates the impeller against calibrated springs and positions the magnet in the second harmonic transmitter to a position corresponding to the fuel flow is named the turbine. The housing fits closely to the outer diameter of both the impeller and the turbine. Around the periphery of the circular cross section of the cylinders, at a fixed radius from the center, are a number of equally spaced holes which are accurately parallel to the axes of the cylinders. Upstream of the impeller is the impeller motor. Downstream of the turbine a second harmonic transmitter is attached to the turbine with two sets of springs of different torque gradients used to restrain it. (See figure 2.) This permits a two-slope scale on the indicato allowing greater sensitivity at low flow rates.

Ilustración 7

NOTE: To prevent dry operation damage and to prolong the life of the transmitter, it is recommended that the “FUEL FLOW” circuit breakers be pulled (open) whenever the engine fuel feed line is drained or the airplane is out of service for maintenance. B. The impeller motor drives the impeller at a constant angular velocity, thus as the fuel es through the impeller it is given a velocity at right angles to the direction of flow. This angular velocity constitutes a change in the momentum of the fuel directly proportional to the mass of fuel flow. As the fuel es through the turbine, the angular component of the momentum is removed, this imparts to the turbine a torque directly proportional to the mass rate of fuel flow. This torque rotates the impeller against calibrated springs and positions the magnet in the second harmonic transmitter to a position corresponding to the fuel flow.

Ilustración 8



IGNITION SYSTEM

A. The purpose of the engine ignition system is to provide a means of initiating or sustaining combustion of the fuel-air mixture in nine can-annular combustion chambers in the combustion section. WARNING: IGNITION VOLTAGE IS DEADLY. DO NOT TOUCH IGNITER PLUGS IF IGNITION IS ON. DO NOT TEST IGNITION SYSTEM WHEN PERSONNEL ARE IN WITH THE IGNITER PLUGS OR WHEN INFLAMMABLES ARE NEARBY.

B. There are three identical ignition systems, one for each engine. The ignition system for each engine consists of control switches, an ignition exciter, two high tension leads and two igniter plugs. 

Electrical Power Supply

A. Ignition power supply The 20-4 joule single pack continuous ignition system is powered by two input voltage at the ignition exciter, 28 VDC and 115 V, 400 Hz. The single unit housing incorporates one power input connection and two output connectors. A dual pack system is also available. The dual pack system has two exciters with a power input connection for each exciter. The two output connections supply the high tension voltage through the exciter cables to the igniter plugs. P&W SB\5880 installed the optional dual pack system.

(a) Ignition exciter 1) The 20-4 ignition exciter is a capacitor discharge system designed to provide ignition for the JT8D Turbofan Engine. This ignition exciter serves the dual purpose of providing intermittent duty starting ignition and continuous duty ignition which are used as required after starting. Two different input voltages are required for the exciter. Both circuits of the ignition exciter are contained in one compact housing with one input power connection and two output connections. The dual pack system has two input power connections and two output power connections. 2) See Fig. for ignition exciter. 3) See Fig. for ignition exciter table of leading particulars. 4) See Fig. for ignition exciter wiring schematic.

Ilustración 9

Distribution 

High tension distribution

(1) The high tension distribution system delivers high voltage from the ignition exciter to the combustion chambers by means of high tension leads and igniter plugs. The high tension voltage from the exciter ionizes the gap at the plug and the result is a spark of very high energy capable of igniting fuel. (a) High tension leads 1) The igniter plug lead assemblies are installed between the ignition exciter and the igniter plugs. The lead assemblies transfer high frequency, high voltage from the ignition exciter to the igniter plugs. (b) Igniter plugs

1) There are two igniter plugs which are mounted on the lower front of the combustion chamber outer case. One projects into the number four combustion chamber and the other projects into the number seven combustion chamber. 

Operation

A. All ground starts or inflight airstarts should be made with the use of the 20-joule DC exciter (firing both igniter plugs). For optimum life of ignition system components, the operating duty cycle is 2 minutes ON, 3 minutes OFF, 2 minutes ON, and 23 minutes OFF. NOTE: If the 20-4 Joule single pack exciter unit has been operated beyond the recommended duty cycle, the unit must be removed and checked per overhaul instructions. The integrity of components within the exciter may have been compromised due to overheating (Ref 74-11-1 R/I). The dual pack exciter installed by SB 5880 has continuous duty. This exciter has no duty cycle limits. B. For continuous operation, the 4-Joule system should be used in lieu of the 20Joule DC system for protection against flameout during takeoff and prior to activating the engine inlet anti-icing system. The 4-Joule system, which can be operated continuously may also be used for protection against flameout if at any other time deemed advisable, such as during periods of moderate or severe turbulence. To conserve the life of the ignition system components, the 4-Joule ignition system should be turned OFF during normal flight conditions whenever engine operation is stable.

Ilustración 10

C. The intermittent duty starting circuit requires an input of 28 VDC nominal while the continuous duty circuit requires an input of 115 VAC 400 Hertz. The intermittent duty starting circuit discharges through both outlets, firing two igniter plugs. The continuous duty circuit discharges only through the outlet marked CONTINUOUS DUTY OUTLET, firing one igniter plug. Spark gaps prevent current from following in one circuit when the other circuit is in operation. NOTE: The exciter is repairable using test equipment available in most airline overhaul shops. Positive hermetic seal is assured through use of a stainless steel case weldment. The exciter capability for continuous operation will assure conformance with potential airline ignition duty cycles. Maximum exciter service life can, however, be attained by ignition utilization per recommended engine/or airplane operational procedures. D. The igniter plug provides a gap across which an electrical spark es to ignite the fuel-air mixture. The igniter plug gap becomes ionized by very high voltage (2226 KV) provided by a high tension transformer. Once the air gap becomes ionized, current stored in a storage capacitor discharges across the gap. This discharge results in a high temperature plasma arc which is capable of igniting the fuel-air mixture. NOTE: Pratt and Whitney recommends the use of any igniter plug regardless of type, that meets the overhaul manual examination and testing requirements. However, each operator must concern himself with determining how long the igniter plug can be expected to continue meeting these requirements while exposed to engine use within his operation. The appropriate inspection/check, cleaning/painting interval can only be established by the individual operator based on his experience.



DESCRIPTION

A. The following engine air systems are installed by Pratt and Whitney and comprise part of the basic engine; anti-icing provisions for the engine inlet guide vanes and forward compressor case, deicing system for the engine fuel supply, and cooling and sealing air for the engine bearings. Refer to figure 1. An air bleed system to balance performance between the engine high and low pressure compressors is also included with the basic engine. B. The following systems are installed by Boeing and require an air bleed from the engine at the appropriate temperature and pressure; airplane air conditioning pack, cowling and wing antiicing air system, generator cooling air, and CSD oil cooling air supply. An air pressure supply for the reverser actuating system is also extracted from the engine high pressure compressor. For engine cowl and inlet duct anti-icing, see Chapter 30, Ice and Rain Protection. C. Further information on the systems which use engine bleed air, but do not form part of the power plant assembly will be found in Chapter 36, Pneumatic.

Ilustración 11

ENGINE ANTI-ICING - DESCRIPTION AND OPERATION A. The engine compressor inlet case and inlet guide vanes are provided with internal ages which allow the circulation of hot air through the assembly. The resulting temperature increase will prevent ice from accumulating on the surface of the inlet guide vanes.

B. The hot air supply is extracted from the eighth stage of the engine compressor section and is controlled by thermostatic regulators and shutoff valves. The switches which control the shutoff valves also energize the nose cowl anti-icing air valves so that both the engine and the cowl anti-ice systems operate simultaneously. Refer to Chapter 30, Ice and Rain Protection. A valve position indication light and selector switch are provided to allow checking the position of any anti-icing valve. The switches and indicator lights are located on the captain’s overhead . C. The engine anti-icing system plumbing lines, control valves and temperature regulators are furnished by Pratt and Whitney. The switches and indicating lights are installed by Boeing. 

Engine Anti-Icing Air Valve

A. An engine anti-icing air shutoff valve is provided in both of the plumbing lines which supply hot air to the inlet guide vanes and inlet guide vane case. The valves are located at the 1 o’clock and 11 o’clock positions aft of the inlet guide vane case. B. Each anti-icing air valve is motor operated and when energized drives to the fully open or closed position as determined by the control switch selection. 

Engine Anti-Icing Air Regulator

A. An air temperature regulator or a fixed orifice plate is located on the upstream side of each antiicing air valve. The regulators or orifice plates limit the volume of anti-icing air supplied to the inlet guide vane assembly. This keeps the guide vane assembly from becoming too hot preventing excessive temperature of the engine inlet air.

Ilustración 12



Operation

A. The engine anti-icing system uses hot air bleed from the engine compressor section to heat the engine inlet guide vane assembly. The air supply is controlled by shutoff and temperature regulating orifices. The shutoff valves are operated by switches located on the pilots’ overhead . B. When the anti-icing air valves are opened hot air circulates around the engine inlet guide vane assembly, the air es through each inlet guide vane and into the inner shroud which forms part of the engine front bearing hub. The air es from the front hub into the engine nose dome, a slot in the nose dome exhausts the air into the engine inlet air stream. C. The valve position indicating lights are illuminated whenever a valve is in a position corresponding to the anti-icing control switch setting. During valve transit, the valve position indicating lights are not illuminated. ENGINE EXAUSTING SYSTEM The engine exhaust system controls the direction of the engine exhaust gases. The engine exhaust system has these sub-systems: * Turbine exhaust * Thrust reverser (T/R).

Ilustración 13

The engine exhaust system controls the direction of the turbine exhaust gases and the fan air exhaust gases. 

Turbine Exhaust System

The turbine exhaust system supplies an exit for the engine exhaust gases. This exit increases the velocity of the exhaust gases. This increases engine thrust. The major components of the turbine exhaust system are the exhaust nozzle and the exhaust plug. 

Thrust Reverser System

The thrust reverser (T/R) system changes the direction of the fan air exhaust to help create reverse thrust. The flight crew uses reverse thrust to slow the airplane after landing or during a rejected takeoff (RTO). The turbine exhaust airflow direction does not change during reverse thrust. The T/R system has a electrohydraulic control system and an indicating system. The T/R system has two thrust reversers. T/R 1 is the thrust reverser for engine Each T/R has a left and right half. Each half has a translating sleeve which moves aft (deploy position) for reverse thrust. The two sleeves work independently from each other. Fan air exhaust goes out radially and forward when the translating sleeves are in the deploy position. Four hinges attach each T/R half to the strut. You must deactivate the thrust reverser before you open a T/R half. Latches are at the bottom of the two halves. The latches keep the two halves together.

Ilustración 14



Thrust Reverser System

The thrust reverser (T/R) system has these subsystems:  Thrust reverser  Control  Indicating. The T/R system controls the direction of engine fan air exhaust for forward and reverse thrust. Reverse thrust helps decrease the speed of the airplane after landing or during a rejected take off (RTO). T/R Control System The T/R control system controls electrical and hydraulic power to the T/R system. T/R Indicating System The T/R indicating system supplies T/R system and T/R control system indication in the flight compartment. The T/R system changes the direction of the fan air exhaust to help decrease the speed of the airplane after landing or during a rejected takeoff (RTO). The T/R system has two thrust reversers. T/R 1 is the thrust reverser for engine 1 (left). T/R 2 is the thrust reverser for engine 2 (right). Each T/R has a left and right half. Each half has a translating sleeve which moves aft for reverse thrust. The sleeves work at the same time, but are independent from each other. Three hydraulic actuators move each sleeve. Rotary flex shafts make sure hydraulic actuators extend and retract at the same rate.



T/R Control System

The T/R control system lets you deploy the T/R when the airplane is less than 10 feet (3 meters) from the ground. You give a deploy signal to the control system when you raise a reverse thrust lever. You supply a stow signal when you return the reverse thrust lever to the stow position. The T/R control valve module controls hydraulic power to the hydraulic actuators. The reverse thrust lever operates the switches necessary to send a deploy or stow signal to the T/ R control valve module. The sync locks prevent the operation of the hydraulic actuators when there is no deploy signal. The primary purpose of the engine accessory unit (EAU) is to control the T/R stow operation. The EAU supplies front built-in-test equipment (BITE) to help you do troubleshooting of the control system. The EAU uses two T/R proximity sensors for each translating sleeve for control. The EAU also interfaces with the T/R indicating system to control the REVERSER light. 

T/R Indicating System

The T/R indicating system supplies these indications in the flight compartment:  REV message on common display system (CDS)  REVERSER light on the P5 aft overhead  Linear variable differential transformer (LVDT) data on the control display unit (CDU). The common display system (CDS) shows the REV message. This message refers to the positions of a T/R’s translating sleeves. Each T/R has LVDTs which supply translating sleeve position data to the electronic engine control (EEC). When on, the REVERSER light shows that there is a failure in one of these areas:  T/R control system  Mechanical failure (which prevents the control system from correct operation). The REVERSER light comes on for 10 seconds during a T/R stow operation. The light will stay on if the T/R does not stow in 10 seconds. The EAU controls this light.

Ilustración 15

Ilustración 16

THRUST REVERSER - HYDRAULIC ACTUATORS AND SYNC SHAFTS The hydraulic actuators move the translating sleeves during T/R deploy and stow operations. The sync shafts make the hydraulic actuators extend and retract at the same speed. The sync shafts also let you manually operate the hydraulic actuators. Each T/R half has three hydraulic actuators. The actuators extend during a deploy operation and retract during a stow operation. Each T/R half has one locking actuator and two nonlocking actuators. The locking actuator must unlock for the other hydraulic actuators on that same half to operate.

The locking actuators have a position mechanism and a manual unlock lever. The position mechanism operates a linear variable differential transformer (LVDT). The manual unlock lever lets you unlock the locking actuator for a manual translation of the T/R sleeve. There are two sync shafts on each T/R half. The locking actuators are the top actuators on each T/R half. The two non-locking actuators are below the locking actuators. All actuators attach to the torque box and to the translating sleeve. You open the fan cowl and move the translating sleeve aft to get access to the hydraulic actuators. The upper sync shaft is inside the deploy hydraulic tube, between the upper and center actuators. The lower sync shaft is inside the deploy hydraulic tube, between the center and lower actuators. The deploy tubes are larger than the stow tubes. You open the fan cowl to get access to the tubing.

Ilustración 17

ENGINE STARTING The engine starting system uses pneumatic power to turn the engine’s N2 rotor during a start or motor procedure. Pneumatic power comes from one of these sources:  APU  Pneumatic ground equipment  Opposite engine. These components control the engine start system:

 Flight compartment switches  Display electronics unit (DEU)  Electronic engine control (EEC). The engine starting system operates on the ground and in flight.

Ilustración 18

The engine starting system uses these airplane and engine systems or components:      

Pneumatic power Electrical power Flight compartment switches Engine fuel control system Engine control system. Common display system (CDS).

 Engine Start Switch You put the engine start switch to the GRD position to turn the engine with the starter. The switch automatically moves to the OFF position at starter cutout. When electrical and pneumatic power is available, this happens when you put the switch to the GRD position:  Electronic engine control (EEC) receives a start signal  APU receives an engine start signal  Start valve opens and the pneumatic starter turns the engine.

The crew uses the FLT position to start the engine in flight when the starter is not necessary. The CONT position supplies continuous ignition 

Start Valve and Starter

The start valve opens to supply power to the starter. Usually, this valve opens when you put the engine start switch to the GRD position. The start valve position shows on the engine display. You can manually open the valve. The starter turns the engine N2 rotor through the engine accessory gearbox (AGB). 

EEC

The EEC protects the engine during start. The EEC shuts off fuel supply to the engine when it finds the engine parameters are out of limits during a start. 

Display Electronics Units (DEUs)

The DEUs are components of the common display system (CDS). The DEUs monitor N2 and let the engine start switch go back to the OFF position at starter cutout.

Ilustración 19

GENERATOR COOLING - DESCRIPTION AND OPERATION A. The generator is a high speed unit that requires cooling whenever it is operating. To provide this cooling the generator is supplied with engine bleed air. A port on the outer fan case of the engine is connected by a large duct, to a fitting on the end of the generator. An exhaust duct, attached to the generator shroud, mates with the cooling air exhaust port in the engine cowl .

B. When an engine is operating, low pressure fan air is forced through the large duct into the generator. The air circulates around the inside of the generator then leaves through the screened openings around the forward end of the generator casing. After leaving the generator, the air enters the generator shroud, es through the exhaust duct and is then exhausted overboard through the exhaust port in the engine cowl .

Ilustración 20



CSD OIL COOLING - DESCRIPTION AND OPERATION

A. Each CSD is cooled by the oil which acts as a lubricant for the unit and as a hydraulic media in transmission. This oil is cooled in a cooler that uses air as the cooling agent. During normal flight ram air is forced through the coolers to extract the heat from the oil. At low flying speeds this ram air may not be sufficient to cool the oil adequately, and on the ground no ram air is available. To provide sufficient cooling under these conditions, engine bleed air is used to induce air flow through the coolers. B. The CSD oil cooling system comprises an air inlet duct, an oil cooler, an exhaust duct, an ejector nozzle and an oil cooler shutoff valve. Each engine has a CSD air scoop which directs ram air into the CSD air inlet duct. The air scoops for engines No. 1 and 3 are located in the lower portion of the engine nose cowl. The engine No. 2 CSD air inlet is situated in the lower portion of the engine inlet duct. The air inlet duct connects to the CSD oil cooler, located on the lower forward portion of the engine, and the exhaust duct connects the aft side of the oil cooler to the air exhaust port in the engine cowl s. The air ejector nozzle is mounted in the exhaust duct and is connected to the thirteenth stage engine bleed air duct by

tubing into which the oil cooler shutoff valve is installed. The valve is located on the upper portion of the forward fan case on the right side of the engine. C. When the system is energized and the oil cooler control switch is in the correct position, the oil cooler shutoff valves will open and allow engine bleed air to flow through the ejector nozzles. Airflow will be induced in the CSD oil cooler to provide cooling for the oil flowing through the coolers.

Ilustración 21

ENGINE CONTROL SYSTEM - DESCRIPTION AND OPERATION A. A manually-operated control system for each engine provides separate control of engine starting and thrust. Starting of each engine is accomplished by use of a single lever to energize the ignition system and to initiate fuel flow to the engine. Another lever assembly controls both forward and reverse thrust by regulating fuel flow and actuating the thrust reverser. An interlock mechanism prevents simultaneous actuation of forward and reverse thrust levers for each engine. B. The engine control system consists of an engine start lever and a thrust lever assembly for each engine, connected by a series of throttle control cables and mechanical linkages to the fuel control units on the engines (Fig. 1). A thrust lever friction regulator applies a braking force to all thrust lever assemblies during forward thrust operation.

C. A drum-and-shaft assembly for each engine transmits engine control cable travel to the upper thrust and start rods. A stop lug on the start cable drum provides a mechanical stop against the bracket when the start lever is at CUTOFF. On engines Nos. 1 and 3, the engine control drum-and-shaft assembly is in the nacelle strut. On engine number 2, the assembly is mounted on a bracket attached to the aft side of the engine firewall. D. Each engine start lever is connected by cables to the drum-and-shaft assembly. Linkage rods connect a crank on the start shaft to a crank on the engine crossshaft and from a crank on the right end of the cross-shaft to a lever on the fuel control unit. E. The thrust lever assembly is connected by cables to the drum-and-shaft assembly. A rod connects a crank on the thrust shaft to a crank on the engine cross-shaft. A crank and linking rod on the right end of the engine cross-shaft connects to the power control shaft on the fuel control unit. F. Actuation of the thrust lever assembly regulates fuel flow in the fuel control unit. For reversen thrust, the lever assembly movement actuates the thrust reverser in addition to increasing fuel flow. It should be noted that the direction of travel of the thrust control cables and drums is the same for decreasing forward thrust as it is for increasing reverse thrust. Two distinct increases in lever loads are encountered as the lever is moved in the direction of increasing reverse thrust. The first increase occurs when a spring-loaded roller engages a detent (maximum power detent). The second increase occurs at the point of engagement of a spring (reverse thrust feel control, if installed).

Ilustración 22

Ilustración 23

ENGINE INDICATING - DESCRIPTION A. The engine indicating systems described in this chapter include the engine pressure ratio (EPR) indicating system, a tachometer system to measure the speed of the low pressure compressor (N1) and high pressure compressor (N2), an exhaust gas temperature (EGT) indicating system, and an airborne vibration monitoring system. B. Each system provides a reading of engine operating conditions on indicators located in the control cabin. This information enables the monitoring of engine output and maintaining a selected flight performance. C. 727-200 Series Airplanes have engine failure warning lights located on the lightshield. The lights are controlled by pressure switches in the engine bleed system and provide visual warning in the event of engine thrust loss (Ref 21-11-0). 

ENGINE PRESSURE RATIO INDICATING SYSTEM - DESCRIPTION AND OPERATION

A. The engine pressure ratio (EPR) indicating system shows the engine power output and is used for setting engine thrust and for monitoring engine performance. The EPR indicating system consists of one inlet pressure (Pt2) sensing probe, six exhaust pressure (Pt7) sensing probes, an engine pressure ratio transmitter and a pressure ratio indicator for each engine. B. The engine inlet and exhaust pressures, sensed by the pressure sensing probes, are transmitted to the pressure ratio transmitter. The transmitter converts the exhaust and inlet pressures into a ratio, provides output signals proportional to

the EPR and transmits the signals to the EPR indicator located in the flight compartment. The indicator transforms the electrical input signals into the indicator pointer shaft rotation and digital three-wheel counter to show the engine pressure ratio. A test receptacle, used to attach a master indicator, is included in the circuit to provide a means of adjusting and checking the system (Fig. 1). On airplanes incorporating EPR-activated takeoff warning system, refer to Chapter 31, Instruments, for a description. 

Inlet Pressure Sensing Probe

A. The engine inlet pressure (Pt2) is sensed by a probe similar to a pitot tube. This probe is mounted through the center of the nose dome with the open end of the tube facing the inlet air stream. The vent hole in the probe functions as the probe ice detector by decreasing engine inlet pressure (increasing EPR) when icing occurs. The probe is anti-iced by the engine anti-ice system. 

Exhaust Pressure Sensing Probe

A. Each engine has six exhaust (discharge, Pt7) pressure sensing probes projected into the stream of turbine exhaust gases. The probes are connected to a common manifold for obtaining an average pressure of the exhaust gases. Exterior connection to the manifold is made at a single point through the fan discharge outer duct at approximately the 7 o’clock position.

Ilustración 24

THRUST REVERSER - DESCRIPTION AND OPERATION A. A thrust reverser unit, located on the aft end of each engine, is used to reduce the length of the landing roll. The thrust reverser is of the clamshell door type, providing thrust reversal by blocking the engine exhaust gas flow path with

clamshell doors and deflecting the gases through openings in the reverser frame onto two external deflector doors or through the cascade vane deflectors which further deflect the gas forward and overboard. Each thrust reverser operates independently. The major components of the thrust reverser are the thrust reverser frame assembly, two axially-mounted clamshell doors, two hinge installations for the clamshell doors, two deflector doors, or cascade vane deflectors, two thrust reverser actuators, a sequence valve on reversers with deflector doors, and a thrust reverser lockout actuator or lock mechanism as applicable. The tailpipe is attached to the thrust reverser unit and is considered as part of the thrust reverser assembly. A thrust reverser shroud assembly installed around the circumference of the thrust reverser forward mounting ring acts as a gas seal between the engine and the thrust reverser. B. The thrust reverser on the center engine (engine No. 2) is the same as the thrust reversers on the strut-mounted engines (engines 1 and 3) except for certain installation features. The center engine thrust reverser is mounted so that the exhaust gases are deflected out the sides of the thrust reverser or aft fuselage since the thrust reverser forms the aft end of the fuselage. The strut-mounted thrust reverser installations are oriented to deflect exhaust gases above and below the engine, resulting in the center engine reverser installation being rotated 90 degrees with respect to the strut-mounted reversers. Also, a double-flanged adapter ring or spacer tapered from bottom to top is installed between the aft end of the engine and the forward end of the thrust reverser on engines 1 and 3 to direct the exhaust gases for the proper thrust angle. Engine No. 3 thrust reverser installation is rotated 180 degrees from engine No. 1 installation to make the installations compatible with the thrust reverser pneumatic tubing installations on the respective engines. C. The thrust reverser is pneumatically actuated and control is provided by a reverse thrust lever (one for each engine) on the pilot’s control stand. Pressure is supplied by engine 13th-stage bleed air source (Ps4). The thrust reverser pneumatic plumbing taps into a 13th-stage air distribution line located on the engine (center engine) and in the strut (side engines). The pneumatic supply line routes air to a thrust reverser directional valve which is mounted to the engine control shaft bracket underneath the forward end of the engine on the center engine and in the strut on the side engines. Thrust lever input is transmitted to the engine controls shaft installation by direct cable connection. A thrust reverser control cam on the lower end of the shaft provides position input for the directional valve. Air is routed from the directional valve by the pneumatic plumbing to two thrust reverser actuators providing the input to the thrust reverser for "cruise" and reverse thrust actuation and operation. The thrust reverser lockout actuator, located upstream of the thrust reverser actuators in the reverse thrust pneumatic line, maintains the thrust reverser mechanically locked in forward thrust position until reverse thrust is selected. When reverse thrust is selected the actuator is pneumatically operated to unlock the reverser and route air to the thrust reverser actuators for reverse thrust operation. The sequence valve is installed only on

reversers with deflector doors and is located upstream of the actuators in the forward thrust pneumatic line to route air sequentially to the clamshell door piston and the deflector door piston of the actuators to retract the clamshell doors to the forward thrust position before retracting the deflector doors. A push-pull control follow-up system running between the thrust reverser clamshell hinge arm and a follow up cam in the thrust control mechanism limits thrust control motion while reverser is in transit or is not in commanded position. D. Reverse or "cruise" thrust operation is selected by use of the reverse thrust lever in the control cab. Actuation of the reverse thrust levers positions the directional valve to direct Ps4 pneumatic air to the appropriate ports of the thrust reverser actuators causing the reverser deflector doors, if installed, and clamshell doors to be actuated to the selected thrust position. The follow-up controls prevent full thrust command from being applied while the thrust reverser is in transit or is not in commanded position. 

Thrust Reverser Frame Assembly

A. The thrust reverser frame is a welded structure to which the clamshell doors, deflector doors or cascade vane deflectors, clamshell and deflector door actuating linkage, if installed, thrust reverser actuators, sequence valve, and tailpipe are attached. Hinge assemblies are installed through the upper and lower vertical centerlines of the frame providing attachment for the clamshell door hinge assemblies.

Ilustración 25

ENGINE OIL SYSTEM - DESCRIPTION AND OPERATION A. The engine oil system comprises an oil storage and an oil distribution system together with the necessary indicating systems which provide measurements of oil quantity, oil pressure and oil temperature. A low oil pressure and filter by indicating system is also provided. B. Each engine is provided with an independent oil system which provides cooling and lubrication of engine gears and bearings. An oil storage tank, mounted on the lower left side of the engine, furnishes a continuous supply of oil to the engine driven oil pressure pump in the accessory drive gearbox housing. An external line carries oil from the pump to a full flow type fuel/oil cooler. Cooled oil is then delivered to the engine bearings through a distribution manifold and galleries formed in the engine structure. C. An oil filter is provided downstream of the oil pump. The filter housing is made integral with the accessory drive gearbox casing. A removable cover is located on the outside of the gearbox to allow replacement or cleaning of the filter core. A by valve is arranged between the inlet and outlet of the filter. If the filter becomes clogged this valve will open and allow a flow of unfiltered oil to circulate in the engine. D. Oil is scavenged from the engine bearing cavities by three pumps and returned to the accessory drive gearbox. From there it is pumped back into the engine oil tank. 

ENGINE OIL TANK - DESCRIPTION AND OPERATION

A. Each engine is provided with a cylindrical shaped oil tank which mounts on the left front face of the accessory drive gearbox and is secured at the front by a strap. With the engines installed on the airplane, the tank holds approximately five U.S. gallons. The remaining tank volume accommodates any oil foaming and expansion. NOTE: Since the NO. 1 and NO. 3 engines are installed in a slightly nose high attitude, the oil tanks on these engines will hold less oil than the tank on NO. 2 engine. B. The tank is constructed of stainless steel and is capable of withstanding, without permanent deformation, the stresses imposed by pressure, vibration, and shock loads such as may occur during landing, rough flight conditions, etc. A baffle serves to minimize sloshing of the oil in the tank. A deaerator in the tank separates most of the air from the returning oil, thus minimizing foaming. C. Servicing of the oil tank is accomplished through a filler port, located in the oil tank sump cavity. Any oil that is spilled in the sump cavity is drained to a fluid drain tank on the underside of the engine. D. For ground check of oil quantity, a dipstick is attached to the self-locking filler cap. A capacitance sensing probe in the tank transmits an electrical signal for remote indication of oil quantity during flight. E. The tank is equipped with an inlet strainer at the filler port. An outlet strainer is located at the drain valve on the underside of the tank. 

OIL TANK DRAIN VALVE – DESCRIPTION AND OPERATION

A. The oil tank drain valve is located at the bottom of the oil tank. The handle is springloaded to the valve closed position. Manually rotating the handle 90 degrees in a clockwise direction opens the valve allowing the oil to drain from the tank. The handle is locked in the valve open position by a ball detent arrangement. Counterclockwise movement of the handle unseats the balls from the detents and allows the spring tension to return the handle to the valve closed position.

Ilustración 26



ENGINE OIL DISTRIBUTION SYSTEM - DESCRIPTION AND OPERATION

A. The engine oil distribution system is a self-contained high pressure design consisting of a pressure system which supplies lubrication to the main engine bearings and to the accessory drives, and a scavenge system by which oil is withdrawn from the bearing compartments, and from the accessories, and then returned to the oil tank. A breather system connecting the individual bearing compartments, the accessory drive gearbox, and the oil tank completes the oil distribution system. For a complete description of the internal engine oil system refer to 72-00, P&WA JT8D Maintenance Manual. 

Oil Pressure System

A. Oil flows by gravity from the oil tank to the engine driven pump, located inside the accessory drive gearbox housing. Pressure oil from the pump flows through an oil filter to a fuel cooled oil cooler. From this unit it es to the various engine bearings. By valves are provided in the filter and in the oil cooler. These valves open and allow oil to continue flowing through the system in the event that either unit becomes clogged. An adjustable pressure regulating valve, installed in the accessory drive gearbox on the pressure side of the pump, maintains system pressure and flow by bying oil back to the pump inlet. 

Scavenge Oil System

A. Four scavenge pumps return oil from the bearing cavities to a sump in the accessory drive gearbox. The scavenge stage of the engine driven pump then

returns the oil to the tank. A deaerator in the tank separates the air from the returning oil, thus minimizing foaming. 

Oil Breather System

A. An oil breather system connects the engine bearing cavities, the accessory drive gearbox, and the oil storage tank. This system controls the pressures in the accessory drive gearbox and main bearing cavities, thereby ensuring adequate oil flow and preventing scavenge pump cavitation during engine operation. B. Oil droplets and vapor are removed from the breather airstream by a centrifugal separator located in the accessory drive gearbox. After ing through the separator unit the breather air is exhausted overboard through a vent pipe. 

P&WAENGINE OIL DISTRIBUTION SYSTEM - DESCRIPTION AND OPERATION

A. The engine oil distribution system is a self-contained high pressure design consisting of a pressure system which supplies lubrication to the main engine bearings and to the accessory drives, and a scavenge system by which oil is withdrawn from the bearing compartments, and from the accessories, and then returned to the oil tank. A breather system connecting the individual bearing compartments, the accessory drive gearbox, and the oil tank completes the oil distribution system. For a complete description of the internal engine oil system refer to 72-00, P&WA JT8D Maintenance Manual. 

Oil Pressure System

A. Oil flows by gravity from the oil tank to the engine driven pump, located inside the accessory drive gearbox housing. Pressure oil from the pump flows through an oil filter to a fuel cooled oil cooler. From this unit it es to the various engine bearings. By valves are provided in the filter and in the oil cooler. These valves open and allow oil to continue flowing through the system in the event that either unit becomes clogged. An adjustable pressure regulating valve, installed in the accessory drive gearbox on the pressure side of the pump, maintains system pressure and flow by bying oil back to the pump inlet. 

Scavenge Oil System

A. Four scavenge pumps return oil from the bearing cavities to a sump in the accessory drive gearbox. The scavenge stage of the engine driven pump then returns the oil to the tank. A deaerator in the tank separates the air from the returning oil, thus minimizing foaming. 

Oil Breather System

A. An oil breather system connects the engine bearing cavities, the accessory drive gearbox, and the oil storage tank. This system controls the pressures in the accessory drive gearbox and main bearing cavities, thereby ensuring adequate oil flow and preventing scavenge pump cavitation during engine operation. B. Oil droplets and vapor are removed from the breather airstream by a centrifugal separator located in the accessory drive gearbox. After ing through the separator unit the breather air is exhausted overboard through a vent pipe. PNEUMATIC STARTING SYSTEM - DESCRIPTION AND OPERATION A. The pneumatic starting system provides means for rotating the engines to the rpm range at which starting can be accomplished when fuel and ignition are supplied. B. The system consists of three pneumatic starters, three starter valves, and the associated hardware. For component location. The pneumatic starting system is controlled by: (1) Start switches, located on the pilots’ overhead . (2) Air conditioning switches, located on the third crewman’s upper . (3) Start levers on the pilots’ control stand. C. With air pressure in the pneumatic manifold, actuating the engine start switch will supply electrical power to open the starter valve. Low pressure air acts on the turbine blades of the pneumatic starter causing it to rotate. Rotation of the starter is transmitted to the N2 compressor through the accessory drive gear system. When the engine has accelerated to starting speed, application of fuel and ignition, by advancing the start lever, should result in the engine starting. For engine start procedure, refer to Chapter 71, Power Plant - General. At starter cutout speed, electrical power is interrupted mechanically, the starter valve closes, and the starting cycle is ended. 

The pneumatic starting system can utilize low pressure air from three separate sources

. Normally, the engines are started with bleed air from the auxiliary power unit (APU). Secondly, the low pressure air can be obtained from a ground source through the pneumatic ground service connection. The cross-bleed air from an operating engine is the third source of low pressure air for starting the remaining engines. However, the cross-bleed starting is not desirable because the operating engine must operate at approximately 80% power setting to develop adequate air pressure for starting another engine.

Ilustración 27

Pneumatic Starter A. The pneumatic starter is a lightweight turbine-type air motor which converts the kinetic energy of compressed air into starting torque sufficient to accelerate the engine to starting speed. Low pressure air and electrical power are required for starter operation. The starter will continue to assist the engine until starter cutout speed is attained. B. The starter consists of a scroll assembly, turbine wheel, reduction gear assembly, engaging mechanism, and an output shaft. The starter is fitted with a shutoff valve to control the inlet air flow. When the valve is open it its air to the inlet connection on the starter scroll assembly; the air then es through the starter vanes of the scroll assembly and is directed radially inward through the turbine wheel imparting high-speed rotation. Exhaust air from the turbine wheel then es through the vendor air outlet screen or through the turbine wheel containment assembly. C. The reduction gear train translates the high speed, low torque of the turbine wheel into low speed, high torque. This output is transmitted through a pawl and ratchet engagement mechanism to the output shaft. From the starter output shaft the cranking torque is transmitted to the N2 compressor by way of the accessory drive gears. A clutch mechanism provides engagement of the reduction gear train with the output shaft for engine starting; when the speed of the output shaft exceeds the speed of the internal gear hub, the clutch mechanism overruns, thus providing automatic disengagement. Turbine overspeed control is maintained by the cutout switch mechanism, which automatically closes the shutoff valve when the output shaft reaches the predetermined cutout speed. This safety switch is mounted on the output shaft and is actuated by centrifugal force. The pneumatic starter is mounted on the accessory drive gear case beneath the engine.

Ilustración 28

Bibliographic:

Aircraft Maintenance Manual Manual JT8D Aviacsa Information