B737-200 L3 Ata 70_80 Power Plant (jt8d).pdf 6vj6k

BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

GENERAL

For Training Purposes Only

Lufthansa LAN Technical Training

POWER PLANT GENERAL

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Lufthansa LAN Technical Training For Training Purposes Only

POWER PLANT GENERAL

BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

HISTORY General The JT8D is an axial flow, dual--spool, fully ducted, turbofan engine. Its design was initiated in April 1960 Some of the certification milestones are: February, 1963 JT8D--1 certified at 14,000 lb. thrust. April, 1963 JT8D--5 certified at 12,000 lb. thrust March, 1966 JT8D--7 certified at 14,000 lb. thrust May, 1967 JT8D--9 cerúfied at 14,500 lb. thrust September, 1968 JT8D--11 certified at 15,000 lb. thrust April, 1971 JT8D--15 certified at 15,500 lb. thrust February, 1974 JT8D--17 certified at 16,000 lb. thrust April, 1976 JT8D--17R certified at 17,400 lb. thrust By the end of 1989, production of JT8D--7 thru --17’s had ended with 11,882 units being produced. The numbers of JT8D engines produced, by dash model, are: JT8D--1/1A/1B/5 1,809 JT8D--7/7A/7B 2,612 JT8D--919A 2,832 JT8D--11 128 JT8D--15 2,525 JT8D--15A 338 JT8D--17 1,069 JT8D--17A 245 JT8D--17R 312 JT8D--17AR 12 By the end of 1988, almost all JT8D--1 and --5 models were converted to other dash models or were no longer in service. January, 1982 ”A” series was released for JT8D--15 and --17s to improve fuel burn. Differences of models can be found in Maintenance Manual, Section 72--00 Description.

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JT8D--7 thru --17 engines are used in the following commercial aircraft: -- Boeing 727 and 737 -- McDonnell Douglas DC--9 -- Aerospatiale Super Caravelle -- Dassault Mercure, and -- Saab Viggen Although this training guide deals only with the JT8D --7 to --17 models, we should mention that: The JT8D--209, which was the first of the --200 series, was certified in 1979 and entered service in 1980. As of 1989, only the --200 series JT8Ds were being manufactured.

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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00


POWER PLANT GENERAL

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JT8D Turbofan Cut - a - Way View Page: 3


POWER PLANT GENERAL

BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

DATA SHEET General The engine parameters presented on the data sheet, are for standard day, engine static condition. Operation Data sheet shows engine condition in the left hand column -- takeoff, maximum continuous, etc. Engine configuration, thrust, etc., are shown across the top. Note flag notes 1 through 15 for data pertaining to engine operation. Example Takeoff Thrust 15,500 lb. Part Power N2 RPM 92.1% Data Plate EGT 446 degrees °C Start EGT Limit Ground start 510 degrees flight start 620 degrees.


Flag Notes:

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Data Sheet P&W JT8D Page: 5


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

NACELLE Purpose: The nacelle is an aerodynamic structure in which the engine is mounted. It has the following purposes: -- It transmits the thrust force of the engine to the airframe -- It directs air to the engine inlet so that its fiow is smooth. -- It has connections between the engine and the airframe for electrical, air, fluids, etc. -- It smoothes the engine to reduce drag.

Miscellaneous Components: Noise reduction by added aconstical s. The fire detection system provides flight deck warning if high temperatures occur by means of electrical resistance, or gas pressure; single or dual zone. The engine mounts are located in two places as follows: -- Front at two points -- Rear at one point The vibration isolators assist in dampening airframe vibration from the engine.

Construction: The nose bullet attaches to the accessory drive housing on the front of the engine. It smoothes the airflow into the engine. The nose bullet is a QEC component. It is also referred to as: -- Inlet bullet -- Nose cone The nose cowl attaches to the front flange on the engine fan case (flange A). It catches and smoothes the incoming airflow. The upper cowl door and the lower cowl door, attach to the nacelle apron. They permit service access to the engine while it is mounted on the aircraft. The thrust reverser attaches to the rear flange on the exhaust case (flange M). It helps to decelerate the aircraft after landing by changing the direction of engine thrust. Quick Engine Change (QEC) Components: The inlet cowl is a divergent duct that sets up the air to the engine. It attaches to flange A. The inlet cowl has s and doors that: -- Are a set of s that aerodynamically smoothes the engine. -- Open to allow maintenance engine access. -- Have pins to allow removal. The exhaust nozzle is a convergent duct that imparts the final velocity to increase thrust. It attaches to flange M. The thrust reverser component provides for aircraft deceleration to reduce landing roll. It also consists of two positions operational only on the ground, stowed, or deployed. The following accessory components are found on the engine: SCL

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Nacelle Configuration Page: 7


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

ENGINE COWLINGS General The engine nacelle is composed of cowl s to provide a smooth airflow over the engines and to protect exterior engine components from damage. Components The engine cowlings consist of the following: A. Fixed fairing B. Forward, mid and aft f airings C. Nose cowl D. Engine cowling, left E. Engine cowling, right F. Tailpipe cowling, left G. Tailpipe cowling, right


Fixed Fairing The fixed fairing is a two--piece section of cowling which fairs with the side cowl s, nose cowl and strut mid faring. The two parts are bolted together at six points. The assembly is then secured to the engine flange mounted brackets by six bolts. Access s in the fairing are for the two engine anti--ice valvs, two forward cone bolts and compressor bleed system filter. Forward, Mid and Aft Fairings The engine to wing fairing, in three sections, acts as a cover between the engine and wing. The removable aft fairing is cantilevered aft from the wing rear spar, mid fairing is attached to the lower wing skin between spars, removable forward fairing is attached to mid fairing and top of nose cowl by latches. On the mid fairing are three access s for various system connections to engine. Below the forward fairing are located all ongine electrical disconnects. Nose Cowl The nose cowl is shaped to provide a smooth airflow over the nacelle and optimum airflow for the engine. It is attached to the forward engine flange by 23 bolts. An anti--icing air duct, is located on the rear face of nose cowl, which continues into the cowl leading edge as a spray tube.

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Engine Cowlings, Left and Right The engine left and right cowl s with the nose cowl, fixed fairing and tailpipe cowl s. Six hook latch fasteners attach the left and right cowl s to the fixed fairing and serve as cowl hinges. Two of these six latches on each cowl are safety latches which 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. 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. Six other hook latch fasteners the left and right cowl s together at the underside of the engine. A pin latch near the forward lower corner of the right cowl engages a fitting on the left cowl to provide a positive safety backup. Various ports and s are provided in the cowls. Left Cowl -- Fuel heater exhaust -- Gearbox breather port -- Accessory drain -- Precooler exhaust -- Oil tank service door -- Constant Speed Drive service door -- Start valve manual override hole -- Two pressure relief doors Right Cowl -- Generator cooling air exhaust Tailpipe Cowlings Left and Right The tailpipe left and right cowl s fair with the engine cowls fixed fairing and thrust reverser. The cowls are hinged and removed in the same manner as engine cowls. Interchangeability The only sections of the nacelle that are intçrchangeable between the two engines are the side and tailpipe cowlings.

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Engine Cowlings Page: 9


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

Design Specifications: The JT8D is an axial--flow, dual--spool, low by ratio, fully ducted turbofan and its exact specifications depend upon the dash model. Take--off Thrust: -- Pounds of thrust, to 84 °F (28.9 °C): * JT8D--7,--7A,--7B: 14,000 * JT8D--9,--9A: 14,500 * JT8D--11: 15,000 * JT8D--15,--15A: 15,500 * JT8D--17,--17A: 16,000 -- Pounds of thrust, to 77 °F (25 °C): * JT8D--17R,--17AR: 16,400 17,400 reserve T.O. thrust By Ratio: Approximately 1 to 1 Total Airflow: 315 to 331 lb/sec Compressor Pressure Ratio: -- LPC:4.2 to 4.4 -- HPC: 3.8 to 4 -- Overall: 15.8 to 17.5 Dry Weight: 3,205 to 3,500 lb.

Illustrations: You rnay see some illustrations in this chapter that are different from the engines that you work on. That is because there are rnany different parts and assemblies in the (large number of) JT8D engines in service. So the illustrations usually show typical parts such as those that are in: -- The later JT8D engines or -- Most of the engines that are in service.


Dimensions: (See illustration)

Configuration: The JT8D has: -- A full length annular duct -- A six stage front (LPC) compressor, including two fan stages -- A seven stage rear (HPC) compressor -- A combustor with nine combustion chambers -- A single stage front compressor drive turbine -- A three stage rear cornpressor drive turbine -- Seven main bearings. Because the JT8D was in production for 25 years, and it is used in different aircraft models: -- A number of JT8D ”dash models” are in service. --There have been many changes to engine design features and cornponents. Note: Difrerences are often found in the configurations of the JT8D engines in service. Differences between the engines are listed in the Introduction of the Maintenance Manual, Section 72--00.

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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00


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Dimensions Page: 11


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

POSITION REFERENCES Purpose: The use of standard directional references makes it easy to say where components and parts of the engine are located. General: Look at the engine from the rear to identify: -- The left and right sides -- The clock (angular) positions.


Clock Positions: The top of the engine is referred to as the 12:00 (twelve o;clock) position. The bottom of the engine is the 6:00 (six o clock) position.

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Position Refferences Page: 13


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

OVERVIEW Gaspath: Primary air is the part of the intake air that flows through the compressor; combustor, and turbine. Its energy level is increased by the compressor and combustor, and much of its energy is used by the turbine. Secondary air (by air) is the part of the intake air that goes through the full length annular duct. It is accelerated (and compressed) by the fan. Both the primary and secondary air through the two fan stages. Primary air is the fan discharge air that flows through the engine core, and secondary air is the fan discharge air that flows around the core. The secondary air mixes with the primary air (exhaust gases) in the exhaust nozzle. Major Engine Sections: The JT8D has a full length annular fan duct. The front compressor (also called the LPC): -- Has six stages (two fan stages, four primary stages) -- Rotational speed = N1 speed -- Is driven by the front compressor drive turbine and is connected to it by a drive shaft (inner). The rear compressor (also called the HPC): -- Has seven stages -- Rotational speed = N2 speed -- Is driven by the rear compressor drive turbine and is connected to it by a drive shaft (outer). The combustor has: -- Nine fuel nozzles -- Nine can--annular combustion chambers Note: ”Can--annular” denotes the configurahon of the burner cans. They he within an annulus (ring). The rear compressor drive turbine (also called the HPT): -- Has a single stage -- Rotational speed = N2 -- Drives the HPC, to which it is connected by a outer shaft.

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The front compressor drive turbine (also called the LPT): -- Has three stages -- Rotational speed = N1 -- Drives the LPC, to which it is connected by a inner shaft. The exhaust section: -- Receives LPT discharge air -- Receives secondary air from outer duct -- Has an optional lobed mixer to mix primary and secondary air. Seven main bearings the two major rotating assemblies, which are: -- LPC/LPT (front compressor, drive turbine and N1shaft. -- HPC/HPT (rear compressor, drive turbine and N2 shaft.

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Gaspath and Major Design Features Page: 15


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

ENGINE STATIONS Description and Purpose: Certain locations in the gaspath are referred to as stations. Each station has a number for easy (and standardized) identification. Station numbers are used to identify the important points in the gaspath for definitions and descriptions of engine: -- Performance -- Configuration and design. Some of the engine station numbers are attached to pressure abbreviations to form short names for the pressures at those locations. Short names are formed for temperatures in the same way. For example:

Ps4 is static pressure at station 4. Is the HPC discharge static pressure. -- Tt7 is total temperature at station 7. Is the turbine discharge temperature.

--


Station Number Definitions: 1 Nacelle inlet 2 Engine (fan) inlet 2.5 Fan discharge (exit through the annular duct 3 LPC discharge; 6th stage air 4 HPC discharge; 13th stage air 5 HPT inlet 6 HPT discharge; inter--turbine inlet 7 LPT discharge 8 Engine discharge 9 Exhaust nozzle outlet.

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Engine Stations Page: 17


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

FLANGES Purpose: The flanges are raised edges or rims that are at right angles to the cases or ducts. They have three functions: -- To be mating surfaces for attaching engine sections together. -- To increase the stiffness of structural parts, such as cases. -- To be used as attachment points for engine components. Flange Names: Each flange is identifled by a letter. The letters are assigned in sequence, starting ”A” for the flange at the front of the engine. One flange has a number added to its identifying letter name to indicate that it is a nonmating flange. That is, it is not used to cases or ducts.Cases, Ducts, and Flanges: The front and rear flange letters are given for each case and duct: A -- B Fan inlet case (titanium) B -- C Fan front case (steel) C -- D Fan rear case (steel) D -- E Fan exit case (aluminum) E -- F Fan discharge front outer duct (aluminum) F -- G Compressor intermediate case (steel) G -- H Fan discharge rear outer duct (aluminurn) H -- J Fan discharge diff outer duct (aluminurn) J -- K Fan discharge combustion/turbine outer duct (aluminum) K -- M Fan--discharge exhaust outer duct (steel or titanium).


NOTE:

FLANGE J1 IS FOR STIFFNESS AND BRACKET MOUNTING. IT IS NOT A MATING SURFACE FOR TWO DUCT SECTIONS.

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Engine Flanges Page: 19

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LEFT SIDE VIEW Components and Ports

* * * * * * * * * * * * * * *

_______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________



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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

LEFT ENG


A/I VALVE

FUEL FLOWMETER (CSD OIL COOLER)

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Left Side View Page: 21

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RIGHT SIDE VIEW Components and Ports

* * * * * * * * * * * * * * *

_______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________ _______________



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Right Side View Page: 23

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FRONT VIEW General Air entering the engine es through the 19 inlet guide vanes. The guide vanes have a shroud at mid--span. The engine anti--icing air es through the hollow inlet guide vanes into the nose dome area and is discharged back to the engine. On the central hub, behind the nose dome, is the drive pad for N1 tachometer generator and connection for Pt2 pressure sense line for engine pressure ratio system. The wires from N1 drive pad and Pt2 pressure line together with 3 oil system lines (supply, scavenge and breather for No. 1 bearing) through the 6 o’clock position inlet guide vane. At the bottom of the inlet are located Tt2 temperature sensor, at 5 o’clock position, for fuel control unit and Pt2 pressure sensor, at 7 o’clock position, for pressure ratio bleed control. Under the engine is located the accessories drive case. Mounted on the front of the drive case are oil tank, fuel pump, fuel filter and fuel control unit. Forward of the drive case, in the center is the fuel heater. Above the oil tank is located the oil cooler.



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Front View Page: 25


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

ENGINE MOUNTS General The engines are attached to the mounts by three cone bolts, two forward and one aft. The forward mounting, with thrust links, takes all thrust loads, plus vertical and side loads, Whilst the rear mounting takes only the vertical and side loads. Thermal expansion is accomodated by all three mounts. Location The engine mounts are located at forward and aft section at the top of the engine.


Component The forward attachment fittings consist of a steel forging suspended from fittings attached to wing structure and the aft attachment fittings is suspended from the outboard track of the inboard flap. The vibration isolation, located in forward and aft mount fittings consist of a resilient material permanently enclosed in a metal case. As engine vibrates, the resilient material deforms slightly, thereby dampening the vibration created by the engine. Three cone bolts are used to attach the engine to the vibration isolators. The two forward cone bolts are identical. Thread protectors should be placed on cone bolts whenever the engine is being removed. Maintenance Aditional component are used as Secondary Assembly, they are a steel cable or a crushable spacer The Steel Cable acting only as aditional if the aft cone bolt is broken. The honeycomb energy absorber is bonded to the end plate at the lower end and to the top of the housing at the upper end. The sides of the honeycomb energy absorber are not bonded to the sides of the housing. There is a small clearance on both sides of the housing, and this permit detect the abnormal vibration and used as secondary . Both are showed on the nex figure.

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Crushable Secondary Detail

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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

CRUSHABLE SECONDARY ASSY


STEEL CABLE SECONDARY ASSY

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Engine Mounts Page: 27


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BOEING 737 - 200 JT8D -- 7 to 17 71 -- 00

ACCESSORY DRIVE General The engine accessory drive case s and drives various accessories required for the operation of the powerplant. The drive case is driven by a ”tower shaft” from the high pressure compressor and is located under the engine. Accessory Drive Components A. Engine controls cross shaft, concentric thrust crank (inner) and start crank (outer) directs input from left to right side of engine. B. N2 tachometer drive pad, right side, turns counterclockwise. C. Hydraulic pump drive pad, right aft face, turns clockwise below drive pad is seal drain connection. D. Constant Speed Drive and generator drive pad, center aft face, turns clockwise. The drive pad is in the cavity which is equipped with a drain plug and standpipe/overflow line with plug. E. Oil supply line for No. 2 and No. 3 bearings, above the CSD drive pad. F. Starter drive pad, left aft face, turns clockwise, below drive pad is seal drain connection. G. Oil filter, left side, with provision for a differential pressure switch. H. Oil pump, combination pressure and scavenge, bottom of drive. I. Oil pressure regulator and sense line from oil cooler outlet, left aft side. J. Breather connection, left side, connects bearing No. 4, 4 1/2, 5 and 6 cavities to enter the accessory drive case. K. Overboard breather port fitting, left side, mates with hole in left cowling. L. Engine oil tank, left front face, with three lines integral with the tank, oil supply to pump, oil return to tank and tank vent to drive case. M. Fuel pump pad, right front face, turns clockwise, below drive pad is seal drain connection. N. Gearbox drain plug and magnetic chip detector, right side.

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Accessory Drive Page: 29


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ENGINE ACCESSORY DRAINS General Various components are driven by shafts from the accessory drive case. On these shafts are seals preventing various fluids from leaking between the component and the accessory case. Should a seal deteriorate leakage will occur and the fluid discharged overboard through the drain system. Location The lines from the various shaft seals and oil tank scupper are connected to a common drain located on forward bottom side of accessory drive case. This drain mates with a drain mast in the left cowling.


Components Shaft seal drains are from the following: A. Constant Speed Drive and generator -- engine type oil. B. Hydraulic pump -- hydraulic fluid or engine type oil. C. Starter -- engine type oil. D. Fuel pump -- fuel. E. Oil tank scupper drain -- engine type oil. Any fluid present in the drain outlet, calls for investigation as to its source, it could indicate leakage past the seal.

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Engine Drains Page: 31

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NOTES :



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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

PERFORMANCE DATA



POWER PLANT PERFORMANCES

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

PERFORMANCE DATA Engine‘s Purpose: The primary purpose of the engine is to supply propulsive power to the aircraft. The engine also supplies pneumatic power, heat, and accessory drive for electrical and hydraulic power. Why Some Knowledge of Performance is Useful Maintenance and troubleshooting can be done more effectivty if you have some knowledge of: -- The operational characteristics of the engine -- The type of loads, temperatures and pressures that occur in an engine This knowledge can also be useful to people who: -- Operate JT8D engines -- Require a familiarization with the JT8D engine.

-- Only a fourth of the prirnary air is burned. Most of the prirnary air is used for engine cooling and some of it is used for other purposes. Thus, 40 lb. of air is burned every second. At takeoff power condition, the pressure ratio for the JT8D overall compressor it 16 to 17:1.

Performance Topics: Engine performance is a large subject which includes to much technical information. But, for the purposes stated above, it is sufficient to discuss just a few topics They are: -- How thrust is produced in the JT8D engine -- Typical JT8D data (such as pressures, temperatures, and RPM) at different engine operating conditions. -- How thrust performance is rated The charts that follow do not include every type of engine performance data. A knowledge of this data can help you to understand the types of stresses that occur in the engine. IN THE DISCUSSION BELOW, THE DATA ARE VERY APPROXIMATE BECAUSE THERE ARE DIFFERENT JT8D MODELS (WITH DIFFERENT MODIFICATIONS) AND OPERATING CONDITIONS. THE PURPOSE OF THIS DISCUSSION IS TO GIVE YOU A GENERAL IDEA OF A TYPICAL JT8D ENGINE. At takeoff power, the airflow through a JT8D engine is about 320 lb/sec. Half of the air goes through the core, and the other half goes through the fan duct. -- Almost all of the secondary air contributes to engine thrust. A very srnall amount of it is used for other purposes.


NOTE:

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Figure 1 SCL

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Performance Data Page: 3



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ENGINE THRUST Purpose The primary purpose of the engine is to supply the required thrust force to propel the aircraft. Overview: There are several ways to describe how a turbofan engine produces thrust: -- Explain the scientific, technical, and mathematical principles. -- Describe the forces that occur in the engine and are transmitted through its structure to the aircraft. We will give a short and simple summary of the principles. Basic Principles of thrust: A jet engine (gas turbine) does not create a thrust force by pushing exhaust gases against atmospheric air. A jet engine is a reaction engine that increases the energy of the gases that go through it to produce thrust. A jet engine produces a thrust force due to Sir Isaac Newton’s Second and Third Laws of Motion. -- Second Law:

-- Add together all positive forces on the internal engine surfaces; these are the forces that push those surfaces forward. -- Add together all negative forces on the internal engine surfaces; these are the forces that push those surfaces rearward. -- The net thrust is equal to: (Total forward forces) minus (Total rearward forces).


* Force = Mass x Acceleration (F=Ma). A force is produced when a mass of air is accelerated (has its velocity increased). This also means that a force is produced when the momentum of the air is increased. (the momentum of something is defined as its mass multiplied by its velocity). -- Third Law: * For every force or action, there is an equal and opposite reaction. A jet engine increases the momentuin of the air that goes through it. This can be stated by the following equation, which is another way to write Newton’s Second Law: -- Force = (Momentum of exhaust gases) minus (Momentum of the incoming air) -- The engine applies a rearward force to the air which is equal to the increase in momentum. Due to the Third Law, the engine is pushed forward by a equal reactive force. The explanation above can also be applied in a practical way, we can calculate the thrust produced by an engine as follows: SCL

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Engine Thrust Page: 5



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ENGINE POWER SETTING CONCEPT Purpose The rating of a gas turbine engine is: -- The manufacturer‘s guarantee of new engine performance, i e: takeoff rating, maximum continuous rating, etc. -- A concept for setting engine power. In general, there are two rating concepts which are utilized for setting engine power. One concept is that of rating an engine to a constant compressor speed, and the other is that of rating an engine to a constant exhaust gas temperature. The bassic Pratt & Whitney engine rating concept is that of rating an engine to a constant exhaust gas temperature (flat rating). In discussing the flat rating concept, the following principles should be kept in mind: -- Ambient air temperature and air density vary inversely. -- Airflow through the engine compressor is a function of compressor speed and air density. -- Compressor speed is a function of compressor airflow and the energy available to the compressor’s turbine. Therefore, when the exhaust gas temperature is maintained at a constant value (constant energy level to the turbine), compressor speed becomes a function of ambient air temperature for a given pressure altitude. The flat rating concept shown on the facing page, below the flat rating temperature (Tt2) point, the engine is producing maximum rated thrust at reduced values of fuel flow, EGT, and compressor speeds. The reduced EGT results in a decreased rate of turbine deterioration resulting in an increased hot section life with no penalty in thrust. Above the flat rating temperature point a constant EGT is maintained. This results in a prolonged hot section life and at the same time provides the thrust to satisfy aircraft certification requirements.

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COMPRESSOR PERFORMANCE Purpose: Compressor performance is utilized to define compressor operating characteristics. Compressor maps are principally the tools of those who design compressors, but they can also be used to show certain operational effects that are of interest to maintenance personnel. Compressor maps are graphs illustrating compressor performance, and are plotted on the parameter axes of compressor pressure ratio vs corrected airflow, for a given compressor speed. This plotting identifies the compressor surge line. The steady state operating line (established by diff, combustor and turbine geometry) is then determined and plotted on the map. The distance between the compressor surge line and the steady state operatmg line determines the surge margin. Explanation of Surge: The causes of compressor surge (which is a more correct term than “stall” ) are too complex to describe here. The effect of surge is as follows: Surge is a sudden decrease of airflow through the compressor. This will decrease of airflow is caused by inability of the compressor to produce the proper pressure rise to maintain the mass of airflow through it. As a result the airflow will break down or reverse its direction.

COMPRESSOR MAPS


Compressor performance, graphically illustrated by compressor maps, is used to define compressor operatmg charactenstics. Compressor maps, used to identify the compressor surge line, are plotted on the parameter axes of compressor pressure ratio Vs. corrected airflow (for a given compressor speed). The steady state operating line (established by diff, combustor, and turbine geometry) is then determined and plotted on the map. The distance between the compressor surge line and tile steady--state operating line determines the surge margin. Compressor maps are used by compressor designers and, because they indicate various operational effects, by maintenance personnel. SCL

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Figure 4 SCL

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Compressor Map Development Page: 9



BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

REAR COMPRESSOR MAP: The rear compressor axis consists of: -- Compressor pressure ratio (Pt4/Pt3) -- Corrected airflow The rear compressor plotted applications consists of: -- Constant compressor speed lines -- Surge line Operating lines: -- Steady state operating line . The steady operating line is determined from engine testing and is a function of the diff, combustor, and turbine geometry. -- ACCEL line . During rapid acceleration the operating line moves towards the surge line. This movement is caused by a rise in burner pressure due to the increase in fuel flow which causes a back pressure at tile HPC exit. This can be compared to the increase in back pressure when the outflow valve is closed in the compressor test rig. -- DECEL line


. During rapid deceleration, the operating line moves away from the surge line. This movement is caused by a drop in burner pressure due to the reduction in fuel flow which reduces the back pressure on the HPC. This can be compared to the reduction in back pressure when the outflow valve is opened in the compressor test.

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Figure 5 SCL

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Rear Compressor Map Page: 11



BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

FRONT COMPRESSOR MAP The front compressor axis consists of: -- Compressor pressure ratio (Pt3/Pt2) -- Corrected airflow The front compressor plotted applications consist of: -- Constant compressor speed lines -- Surge line -- Operating lines . Steady state operating line The steady state operating line determined from engine testing and function of the díff, combustor, turbine geometry. . ACCEL line During rapid acceleration, tile operating line first moves toward the surge line because of an increase in burner pressure. Then it moves away from the surge line as the HPC speeds up. The rotational speed (N2) of the HPC increases more rapidly than the HPC speed (N1) because the HPC has less inertia. When N2 increases, the airflow through the HPC increases. This reduces the back pressure on the HPC.


. DECEL line During rapid deceleration the operating line moves toward the surge line. This movement is caused by an increase in back pressure on the HPC. This occurs because the HPC slow down faster than the LPC, so it cannot accept all of the airflow from the LPC. This is a stronger effect than the reduction in burner pressure.

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Figure 6 SCL

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Front Compressor Map Page: 13

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NOTES :




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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

ENGINE



POWER PLANT ENGINE

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Page: 1


POWER PLANT ENGINE BUILD GROUPS Description and Purpose: A build group is the largest assembly of engine parts that can be conveniently removed or installed as a unit. This permits: (1) easy disassembly of the engine for repair; (2) easy assembly for build--up. NOTE:


BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

A BUILD GROUP IS SOMETIMES REFERRED TO AS A MODULE, MODULAR COMPONENT, OR ASSEMBLY.

Build Group Concept to the JT8D: The JT8D engine was designed before the build group concept was fully developed. Thus the build group concept does not apply as well to the JT8D as it does to the newer Pratt & Whitney engine models. The JT8D engine is made up of 12 build groups and: -- Two sections whose assembly is completed when they are attached to the engine. (These can be referred to as ”build group sections”.) -- Some assembly parts which are used to connect some of the build groups together. The word ”section” is used in two different ways: -- There are six major engine sections: air inlet, compressor, combustion, turbine and exhaust, accessory drives, and fan discharge. -- There are two build group sections. Because the JT8D has a full length fan duct: -- It is not practical for the engine to be fully modular. (Some parts of the duct must he removed to permit access to the cases.) -- The name ”fan discharge section” has two diiferent definitions: As a major engine section, it includes all parts of the fan duct downstream from the fan exit guide vanes. As a build group section, it is just one part of the fan duct.

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POWER PLANT ENGINE

Figure 1 SCL

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Build Groups Page: 3


POWER PLANT ENGINE

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MAJOR ENGINE BUILD GROUPS: The major JT8D build groups are usted, and the first two elements of their ATA numbers are given. Two of the build groups are build group sections. -- Front accessory drive group 72--21--00 -- Fan inlet group 72--23--00 -- Front compressor group 72--33--00 -- Compressor intermediate group 72--34--00 -- Rear compressor group 72--3~00 -- Diff group 72--37--00 -- Diff outer fan duct group 72--38--00 -- Combustion and No. 5 bearing section 72--41--00 -- Turbine nozzle group 72--51--00 -- Rear compressor drive turbine group 72--52--00 -- Front compressor drive turbine group 72--53--00 -- Exhaust case group 72--54--00 -- Accessory gearbox group 72--61--00 -- Fan--discharge section 72--71--00 The two build group sections are: -- Fan--discharge section -- Combustion and No. 5 bearing section. The exhaust case group is sometirnes referred to as a build group section. That is because it has some parts that can be installed after it is attached to the engine. We refer to it as a build group, not as a build group section. The exhaust case group can have different names in different rnanuals. For example, it can be referred to as: -- (Engine) exhaust case section (group) -- (Engine) exhaust section (group) -- (Engine) exhaust group. The accessory gearbox can have different names in different manuals. For example, it can be referred to as: -- Main accessory drive gearbox -- Accessory drive gearbox -- Main accessory gearbox. SCL

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POWER PLANT ENGINE

Figure 2 SCL

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Section Locations Page: 5


POWER PLANT ENGINE

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FRONT ACCESSORY DRIVE GROUP (72--21--00) Purpose: The front accessory drive group contains gears which are turned by the front compressor. Those gears turn: -- The N1 tachometer; which is a QEC component -- The scavenge oil pump for the No. 1 bearing.


Construction: Cast magnesium housing contains the: -- Scavenge oil pump for the No. 1 bearing -- N1 tachometer drive -- Oil nozzle.

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POWER PLANT ENGINE

Figure 3 SCL

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Fron Accessory Drive Page: 7


POWER PLANT ENGINE

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FAN INLET GROUP (72--23--00) Purpose: The fan inlet directs the inlet air to the fan blades at the correct angle and with a minimum pressure loss. It s the No. 1 bearing and the front accessory drive. Construction: The fan inlet case has two bosses for anti--icing air. They are at the 2:00 and 10:00 positions. There are 19 equally spaced inlet guide vanes which: . Are hollow . Are anti--iced the No. 1 bearing. -- Eighteen of the vanes have two internal ribs which go from one end to the other. -- The other vane is the master vane, which is at the 6:00 position. It is also referred te as the six o’clock vane. It has five internal tubes that go from one end to the other. Probes: There is a Pt2 probe at the 5:00 position on the inlet case: It senses the total pressure at the inlet. This Pt2 sense pressure goes to the pressure ratio bleed control (PRBC), which is part of the compressor bleed system. There is a Pt2 probe* at the front of the nose bullet*. -- It is part of the EPR indicating system. A tube* goes from the Pt2 probe to a fitting* inside of the nose bullet.


That fitting is attached to the Pt2 tube at the front of the No. 1 bearing

compartment. The Pt2 tube goes through the master vane to the bottom of the inlet case. From there, the Pt2 sense pressure goes to the aircraft.

* These are airframe supplied. -- There is a Tt2 sensor at the 7:00 position on the inlet case. It senses the total temperature at the inlet. This Tt2 sense temperature goes to the fuel control unit.

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POWER PLANT ENGINE

Figure 4 SCL

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Fan Inlet Page: 9


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

FAN INLET GROUP Master Vane The master vane has five tubes that go from the bottom of the engine to the central hub as follows: -- The tachometer transfer tube is a conduit for the electrical lead from the N1 tachometer -- Breather air (with oil vapor from the No. 1 bearing compartment) flows down through the oil breather tube. -- Oil that has lubricated and cooled the No. 1 bearing flows down through the scavenge oil tube. -- Oil that is pumped to the No. 1 bearing goes up through the pressure oil tube to lubricate and cool the bearing --

Pt2 air pressure goes through the Pt2 inlet case tube.

The five tubes go into the central hub and turn towards the front of the engine (towards the front accessory drive area). The locations of the fittings at the end of these tubes are:

Pt2 inlet case tube

3:30 position

-----

Pressure oil tube Scavenge oil tube N1 tachometer tube Oil breather tube

4:00 position 5:00 position 6:00 position 8:00 position.


--

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POWER PLANT ENGINE

Figure 5 SCL

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Master VAne Page: 11


POWER PLANT ENGINE

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FRONT COMPRESSOR GROUP (72-- 33-- 00) Purpose: The front compressor group increases the pressure of the primary airflow. The fan stages also accelerate the secondary air. Construction: The cases are designed for fan blade containment. The cases are: -- Fan front case -- Fan rear case. There are six stages in the front compressor. They are: -- Two stage fan (compressor stages 1 and 2) (primary and secondary) -- Four stage low pressure compressor (compressor stages 3, 4, 5, and 6) (primary). The front compressor rotor assembly consists of: -- Integral front hub and fan disk -- Integral rear hub and 4th stage disk -- Disks secured by tie bolts and nuts -- Interstage spacers with knife--edge seals -- ed by the No. 1 and 2 bearings -- Driven by turbine stages 2, 3, and 4. The front compressor blades consist of: -- 1st stage (fan): . 30 blades in JT8D--7 models . 27 blades in JT8D--9/11/15/17 models

-- 3rd stage: . 64blades . Dovetail slots and retaining plate -- 4th stage: . 62 blades . Dovetail slots and tablocks -- 5th stage: . 64blades . Dovetail slots and tablocks -- 6th stage: . 62 blades . Dovetail slots and tablocks. The front compressor stators consist of: -- Five stator assemblies (compressor stages 1 through 5) -- Continuous ring construction -- Bolted together -- Interstage and blade tip seal lands. NOTE:

THE FAN EXIT CASE IS AN ASSEMBLY PART, NOT PART OF A BUILD GROUPS. IT IS A TWO PIECE (SPLIT) CASE.


. Dovetail slots and retaining plate -- 2nd stage (fan): . 42 blades in JT8D--7 models . 40 blades in JT8D--9/11/15/17 models . Pin--t attachment with rivets

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POWER PLANT ENGINE

Figure 6 SCL

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Front Compressor Page: 13


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

COMPRESSOR INTERMEDIATE GROUP (72--24--00) Purpose: The cornpressor intermediate group provides the following: -- s the No. 2 and 3 bearings -- s the accessory gearbox -- Has attachrnent points for the engine front mount brackets.

Borescoping The cornpressor intermediate group has borescope ports that give a view of the 6th and 7th stage area.


Construction: The ducts and cases consist of: -- Fan--discharge front compressor inner duct * -- Fan--discharge front compressor outer duct * -- Cornpressor intermediate case* (which includes the tower shaft housing) -- Fan--discharge rear cornpressor inner duct * -- Fan--discharge rear cornpressor outer duct. * * These three items are a single assembly. Thus the term ”compressor intermediate case” is frequently used to refer to all three of them as a unit. The compressor intermediate group includes: -- The 6th stage stator -- Vanes that are in the secondary gaspath -- Some fairings in the secondary gaspath for 6th and 8th stage bleed air. The fan--discharge rear cornpressor inner and outer ducts have bosses for 6th and 8th stage bleed air. One or three antisurge bleed valves are attached to the exterior of the fan--discharge rear cornpressor inner duct. The number of antisurge bleed valves is different number as a different JT8D configurations. Some of the earlier JT8D engines do not have any 8th stage bleed valves. Bleed Air 6th stage air is supplied for: -- The Ps3 sense signal to the Pressure Ratio Bleed Control (PRBC) -- Aircraft use 8th stage air is supplied for: -- Engine anti--icing -- Engine antisurge bleed (SB 4597) -- Aircraft use. SCL

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


POWER PLANT ENGINE

Figure 7 SCL

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Compressor Intermediate Page: 15


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

REAR COMPRESSOR GROUP (72-- 36-- 00) Purpose: The rear compressor group increases the pressure of the primary air. NOTE:

THE REAR (HIGH PRESSURE) COMPRESSOR RECEIVES PRESSURIZED PRIMARY AIR FROM THE LOW PRESSURE COMPRESSOR. IT SENDS THAT AIR TO THE DIFF.

Construction: The rear compressor stages consist of: -- Seven stages of high pressure compressor blades (compressor stages 7 through 13) -- Six stages of high pressure compressor vanes (compressor stages 7 through 12) The rear compressor rotor assembly consists of: -- Interstage spacers with integral knife--edge air seals -- Front hub, disks, spacers, and rear hub held together by 12 tie rods. -- Is a thrust balance chamber for the No. 4 bearing. -- ed by the No. 3 and 4 bearings. -- Driven by the turbine 1st stage. -- Drives the accessory gearbox. The rear compressor blades consist of: -- 7th stage

Dovetail slots and tablocks -- 11th stage 70 blades Dovetail slots and tablocks -- 12th stage 80 blades Dovetail slots and tablocks -- 13th stage 74 blades Dovetail slots and tablocks The rear compressor stators consist of: -- Six stator assemblies (compressor stages 7 through 12) -- Continuous ring construction -- Interstage blade tip seal lands.

60 blades Pin--t attachinent with rivets -- 8th stage For Training Purposes Only

58 blades Dovetail slots and tablocks -- 9th stage 60 blades Dovetail slots and tablocks -- 10th stage 64 blades SCL

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POWER PLANT ENGINE

Figure 8 SCL

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Rear Compressor Page: 17


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

DIFF GROUP (72--37--00) Purpose: The diff group primary function is: -- Receives high pressure primary air from the rear compressor -- Reduces air velocity suitable for combustion. Construction: The diff case functions are: -- s the No. 4 bearing -- Holds the nine fuel nozzles -- Has mounting structures for the nine combustion chambers. The diff case has nine hollow struts with holes in their sides. They are between the fuel nozzle s and slightly forward of them. 13th stage bleed air enters these struts through their side holes, and it flows outward to the 360 degree manifold on the exterior of the diff case. The following tubes go through these struts: -- Two (oil system> breather tubes (all models) -- One pressure oil tube (all models) -- One scavenge oil tube (all models) -- One seal air pressure tube (8th stage air) (labyrinth seal only) -- Three seal air bleed tubes (only if the No. 4 bearing seal is a labyrinth type) (JT8D--11 thru --17 only) -- Two seal air bleed tubes (JT8D--1 thru --9 only) -- At the 6:00 position: One of the three seal air bleeds (if the No. 4 bearing has labyrinth seal) or a seal drain (if the No. 4 bearing has a carbon seal) The diff group also includes: -- The compressor exit stator assembly (13th stage) -- Two fuel manifold fairings for the fuel inlet tubes.

P air has a turbine cooling air pressure test port, forward of the manifold at the 10:30 position. THE FAN--DISCHARGE DIFF INNER DUCT IS ATTACHED TO THE DIFF GROUP BUT IS NOT PART OF IT. THE FAN DISCHARGE DIFF INNER DUCT IS AN ASSEMBLY PART. The primary and secondary fuel rnanifolds on the right side of the diff case carry fuel to four fuel nozzles. Fuel enters these manifolds through fuel inlet tubes that are in a fan duct fairing at the 5:00 position The prirnary and secondary fuel manifolds on the left side of the diff case carry fuel to five fuel nozzles. Fuel enters these manifolds through fuel inlet tubes that are in a fuel rnanifold fairing at the 7:00 position. At the base of each fuel nozzle , there is: -- A connection to the prirnary fuel manifoid. -- A connection to the secondary fuel rnanifold. NOTE:

Bleed Air: 13th stage air has: -- Bosses on the manifold at the 5:00 and 7:00 positions for the antisurge bleed valves -- Bosses on the manifold at the 10:00 and 2:00 positions for fuel heat and aircraft use. SCL

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POWER PLANT ENGINE

Figure 9 SCL

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Diff Page: 19


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

The different JT8D models have several different types of: -- Fuel manifolds -- Fuel nozzle s -- Connections between the manifolds and fuel nozzle s. Three types of connections are shown on the facing page. They are typical of what you may see on JT8D engines, but sorne variations are not shown. For example, the conical seat designs have different types of gaskets. In the early JT8D engines (prior to SB 2722): -- On both sides of the engine, the prirnary and secondary fuel manifolds are a single assembly. Each assembly includes (four or five) flat plates for the fuelnozzle s. -- The base of the fuel nozzle has a flat surface that mates with a flat plate on the fuel manifold assembly. -- The two holes through which primary and secondary fuel flow (from the manifolds into the fuel nozzle ) have chevron ”pinch” seals. -- Some designs have a spacer plate between the two mating surfaces In the later JT8D engines (SB 2722): -- The primary and secondary manifolds are separate, and there are no ”flat plates”. -- For each fuel nozzle , both of the fuel manifolds have a fitting with a retaining nut. -- The fuel nozzle has two threaded fittings for those retaining nuts. Those fittings have conical seats on which there is a tapered metal Voishan seal. -- The newer design (SB 4484) has a transfer tube inside of the threaded fitting. The different JT8D models have several different types of fuel nozzles. The fuel nozzles have been redesigned to decrease emissions. The two major types of fuel nozzles are shown on the facing page. They are: -- The low emission type which is newer -- The type of fuel nozzle used prior to the low emission type.

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Tipical Fuel Nozzles

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POWER PLANT ENGINE

Figure 10 SCL

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Nozzle Seal - SB 4484 Page: 21


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

DIFF OUTER FAN DUCT GROUP (72-- 38-- 00) Purpose: The fan discharge diff outer duct is part of the outer fan duct, which contains the secondary airflow. Construction: The fan discharge diff outer duct is a one--piece component. -- This duct is attached to: The fan discharge rear outer duct (at flange H) The fan discharge combustion outer duct (at flange J). -- It has bosses that are used to attach: The Pressurizing and Dump (P&D) valve The Pressure Ratio Bleed Control (PRBC)


Lines for 13th stage bleed air, oil (pressure, scavenge, breather), and fuel. -- Some other parts are attached to it, such as elbows, tubes, and brackets.

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POWER PLANT ENGINE

Figure 11 SCL

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Diff Outer Fan Duct Page: 23

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FAN DISCHARGE SECTION (72-- 71-- 00) Purpose: The fan discharge combustion outer duct is part of the outer fan duct, which contains the secondary airflow. Construction: The fan discharge combustion outer duct is a two--piece case with top and bottom halves. This duct is attached to: -- The fan discharge diff outer duct (at flange J) -- The fan discharge exhaust outer duct (at flange K). The fan discharge combustion outer duct: -- Holds the ignition exciter(s) -- Has two ports for igniter plugs -- Has a fuel drain boss -- Has three service bleed ports for secondary (fan) air. The fan discharge combustion outer duct has: -- Bosses and flanges to attach the components identified above. In later JT8D models, the fan discharge combustion outer duct has a liner that absorbs sound. It is referred to as ” noise suppression insulation”.



POWER PLANT ENGINE

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Figure 12 SCL

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Fan Discharge Section Page: 25


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

COMBUSTION AND NO. 5 BEARING SECTION (72-- 41-- 00) Purpose: In the nine combustion chambers (burner cans), the fuel and air are mixed and burned to add energy in the prirnary gas path. The heatshields insulate the following components from the high temperatures in the combustion section: -- The No. 4, 4 1/2, and 5 bearings -- The N1 and N2 driveshafts -- The No. 4 bearing oil scavenge pump. -- This build group section s the No. 5 bearing. Construction: The combustion and No. 5 bearing section is attached to three other build groups: -- Diff group -- Turbine nozzle group -- Rear compressor drive turbine group. The combustion and No. 5 bearing section is attached to some assembly parts: -- The No. 4 bearing -- The oil scavenge pump for the No. 4 and No. 5 bearing -- Part of the No. 5 bearing The combustion chamber outer case attaches to two ducts to continue the inner surface of the fan duct. They are: -- The fan--discharge diff inner duct -- The fan--discharge turbine inner duct. The combustion chamber outer case has: -- Two combustion chamber drain bosses -- Two igniter plug bosses

The fuel drain manifold attaches to the cornbushon chamber outer case at the two drain bosses. At each boss there is a “reed~type” fuel drain valve. -- The combustion chamber inner case attaches to: . The inner rear flange of the diff case . The outer flange of the No. 5 bearing housing. Heatshield: There are some heatshields in the combustion chamber inner case. The illustration below shows three ”No. 4 and 5 bearing” heatshields, but: -- Some cylindrical shields that slide inside of them are not shown. -- The ”No. 4 and 5 bearing” heatshields are also referred to by other names such as ”turbine shafts bearing” heatshields. The inner heatshield has: -- A tube for the pressure oil tube -- A tube for the scavenge oil tube -- A bellows design at one end to perrnit axial movement due to thermal expansion.

-- A Ps4 boss.

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POWER PLANT ENGINE

Figure 13 SCL

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Combustion and No. 5 Bearing Section Page: 27


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

COMBUSTION CHAMBERS: Each combustion chamber (burner Cans) has: -- A set of welded liners -- A front mounting lug -- Low emission design (in the later JT8D models). The later JT8D engines, and many reworked engines, have a magnesium zirconate coating on the inside surfaces of their combustion chambers. A combustion chamber has three types of holes that permit primary air to enter it. -- The combustion holes are located near the front of the combustion chamber. The air that goes through these holes is used primarily for combustion. -- The dilution holes are located near the rear of the combustion chamber. The air that goes through these holes is used primarily to cool the very hot gases before they go out of the combustion chamber. -- There are small holes around the front of each segment (liner). The air that goes through these holes flows against the inner wall of the combustion chamber. It helps to insulate the combustion chamber from the very hot combustion gases. Other Features: -- There are crossover tubes between adjacent combustion chambers. They propagate the flame from one combustion chamber to the next one. -- Combustion chambers 4 and 7 have short guides that the igniter plugs go into. DIFFERENCES BETWEEN THE ENGINES ARE LISTED IN THE INTRODUCTION OF THE MAINTENANCE MANUAL, SECTION 72--00.


NOTE:

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POWER PLANT ENGINE

Figure 14 SCL

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Combustion Chamber Page: 29


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

TURBINE NOZZLE GROUP (72--51--00) Purpose: The turbine nozzle group s the rear (outlet end) of the nine burner cans. The 1st stage turbine vanes guide the flow of the primary gas from the burner cans through a precise area so it arrives at the 1st stage turbine blades at the correct angle, velocity, and pressure. NOTE:

THE 1ST STAGE TURBINE VANES ARE ALSO REFERRED TO AS THE TURBINE NOZZLE GUIDE VANES (NGV).

Construction: The turbine nozzle case is bolted to the rear of the combustion chamber outer case, it also: -- Is made of steel that resists heat and corrosion -- s the combustion chamber guides -- Is also referred to as: . Turbine outer front case


. Turbine nozzle case. -- Is cooled by 13th stage air. -- The turbine nozzle group s the combustion chamber rear guides

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POWER PLANT ENGINE

Figure 15 SCL

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Turbine Nozzle Group Page: 31

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TOBI DUCT The Tangential On Board Injection (TOBI) duct: -- Sends cooling air to the 1st stage turbine rotor -- Is also referred to as the ”turbine cooling air duct assembly” or the ”cooling air duct” -- Is used in the later JT8D models, --15 through --17AR. -- The combustion chamber outer duct and the combustion chamber inner outlet duct are: -- Non louvered (solid) in the earlier models -- Louvered in the later JT8D models.



POWER PLANT ENGINE

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POWER PLANT ENGINE

Non -- Louvered Figure 16 SCL

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Louvered Tangencial On Board Injection Duct Page: 33

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

TURBINE NOZZLE GUIDE VANES -- TNGV: The 1st stage turbine vanes are: -- Internally cooled by 13th stage air -- Coated for protection -- Individually replaceable -- Bolted at the inner and outer vane roots -- Classified by size to accurately control the size (area) of the primary gaspath from the combustion section to the turbine section. In the later JT8D models (--15 through --17AR), the 1st stage turbine vanes have inner and outer platform seals which are sometimes referred to as ”feather” seals. The three types of vanes shown below were developed for different JT8D production models. -- Partial Inner Cooling Flow -- Full Inner Cooling Flow -- Full Inner and Convex Cooling Flow But the newer types can be found in models other than the ones that they were originally developed for.



POWER PLANT ENGINE

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Figure 17 SCL

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Turbine Nozzle Guide Vanes Page: 35


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

REAR COMPRESSOR DRIVE TURBINE GROUP (72-- 52-- 00) Purpose: The rear compressor drive turbine group changes thermal and kinetic energy into rotational energy (shaft horsepower) to turn the N2 assembly. The rear compressor drive turbine: -- Is turned by the quickly expanding gases that come through the turbine nozzle guide vanes -- Drives the rear (high pressure) compressor -- Is also referred to as the High Pressure Turbine (HPT)

The 1st stage turbine disk: -- Has 80 blades in all JT8D models other than the JT8D--15A, 17A, and 17AR -- Has 64 blades in the JT8D--15A, --17A, and --17AR -- Has rivets for axial blade retention -- Has side plates on the JT8D--15 through --17AR -- Is bolted to the shaft.

Construction: The rear compressor drive turbine group rotor assembly: -- Includes the rear compressor drive turbine shatt which is the outer (N2) drive shaft -- Includes the 1st stage turbine disk and blades -- Is ed by the No. 5 bearing. The Rear Compressor Drive Turbine Shaft: -- Is attached to the rear hub of the rear compressor by a spline and coupling bolt -- Holds the inner ring for the No. 5 bearing -- Holds the following assembly parts . Outer ring for the No. 4 1/2 bearing . Carbon seal spacer (seal seat)


. Retaining nuts, lock, snap ring and spacer. -- Is referred to as the: . Outer (main) driveshaft . N2 driveshaft . HPC/HPT (rotor) driveshaft. -- A classified turbine shaft spacer is used to make sure that the position of the front turbine will be correct when the engine operates.

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


POWER PLANT ENGINE

Figure 18 SCL

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Rear Compressor Drive Turbine Page: 37

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

Blades: The 1st stage turbine blades have: --Shrouded tips -- ”Fir tree” roots for radial blade retention. The 1st stage turbine blades are: -- Solid construction on the JT8D--1 through --9 -- Internally cooled by P air on the JT8D--11 -- Internally cooled by 13th stage air through the TOBI duct on the JT8D--15 through --17AR. Different seal designs are used in the 1st stage turbine blade tip area in different JT8D models. -- There are several diiferent designs that use knife edge seals -- There are also several different designs that use honeycomb seals. The illustration below is for a typical performance improved JT8D model (--15A, --17A, and --17AR).



POWER PLANT ENGINE

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


POWER PLANT ENGINE

Figure 19 SCL

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Blade Turbine Configurations Page: 39


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

FRONT COMPRESSOR DRIVE TURBINE GROUP (72--53--00) Purpose: The front compressor drive turbine group changes thermal and kinetic energy into rotational energy (shaft horsepower) to turn the N1 assembly. The front compressor drive turbine: -- Is turned by the primary gas flow from the turbine 1st stage -- Drives the front (low pressure) compressor -- Is also referred to as the Low Pressure Turbine (LPT) Construction: This build group is the ”unit turbine”. It includes both the rotor and stator assembly. The rotor assembly is ed by the No. 2, 4 1/2, and 6 bearings. A classified turbine shaft spacer is used to make sure that the position of the rear turbine will be correct when the engine operates. The non--rotating part of the front compressor drive turbine group includes: -- The turbine rear case -- The second, third, and 4th stage turbine vanes, which are triple clustered in the JT8D--15A/17A/ and 17AR models -- The honeycomb seals (in the later JT8D models) and spacers between the stators -- The inner stator shrouds. The rotating part of the front compressor drive turbine group includes: -- The front compressor drive turbine shaft -- Three turbine disks and blades

The blades in the rear turbine section: -- Have shrouded tips with knife--edge airseals -- Have ”fir tree” roots for radial retention -- Are retained axially by rivets. The front compressor drive turbine shaft is: -- Attached to the front cornpressor rotor rear hub by a spline and coupling bolt -- Attached to the rear turbine between the second and 3rd stage disks -- Referred to as the . Inner main driveshaft . N1 driveshaft . LPC/LPT (rotor) driveshaft. -- Contains a trumpet shaped assembly which is referred to as the “No. 4 1/2 and 6 oil transfer tube” or ”bearings pressure and scavenge oil tubes assembly”.


. 2nd stage 88 Blades . 3rd stage -- 92 blades . 4th stage -- 74 blades -- Spacers and airseals between the disks -- Twelve tierods and 12 bolts which hold the disks, hubs, shaft, and spacers together -- 4th stage turbine rotor hub -- The No. 4 1/2 and 6 bearings (but not their outer rlngs).

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


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Figure 20 SCL

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Front Compressor Drive Turbine Group Page: 41


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

EXHAUST CASE GROUP (72-- 54-- 00) Purpose: The exhaust case group has the following functions: -- Controls the flow of the primary and secondary exhaust gases and sends them out of the engine -- s the No 6 bearing -- Contains probes for the exhaust temperature and pressure -- Provides an attachment point for a mixer to decrease exhaust noise. Construction: The ducts and cases consist of: -- Fan discharge exhaust outer duct or fan exhaust outer duct -- Fan discharge turbine exhaust inner rear duct or fan exhaust inner duct. Fan exhaust outer duct: -- Has eight struts to hold the exhaust case -- Has attachment points for the rear engine mounts. Exhaust case: Is an integral assembly -- Holds the No. 6 bearing -- Holds the exhaust mixer.


The Tt7 probes are line relaceable.

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


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Figure 21 SCL

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Exhaust Case Group Page: 43

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

EXHAUST MIXER (72-- 55-- 00) Purpose: The exhaust mixer reduces engine noise by uniformly mixing primary and secondary airflow. Construction: The mixer is composed of a one piece convoluted sheet metal duct that can be added to the exhaust case with a guide ring. The mixer is part of a noise reduction kit (hush kit) to comply with stage 3 noise. The exhaust mixer is presently an option item as per SB 5947.



POWER PLANT ENGINE

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


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Figure 22 SCL

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Exhaust Mixer Page: 45


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

ACCESSORY GEARBOX GROUP (72--61-- 00) Purpose: Transrnits mechanical power frorn the N2 rotor to drive some of the engine cornponents. Contains, and provides attachments points for, some engine components and accessones. The accessory gearbox group is usually referred to as the ”accessory gear box”, without the word ”group”. Also, the accessory gearbox is sornetirnes referred to as: -- Accessory drive gearbox -- Main accessory gearbox.


Construction: The accessory gearbox is a cast two--piece assembly. It attaches to the engine at three places. The accessory gearbox gets mechanical power from the gearbox driveshaft (also referred to as the ”tower shaft”, which: -- Is inside of a ”strut” in the intermediate case -- Goes into the accessory gearbox. The gearbox driveshaft is dnven by the front compressor drive turbine as follows: -- A bevel gear is splined to the front hub of the rear compressor (N2) rotor, in front of the No. 3 bearing. -- A mating bevel gear is splined to the top of the gearbox drives aft. Design Features: The accessory gearbox external components and features (front) are: -- A mount pad to attach the oil tank (using bolts) -- A mount pad to attach the fuel pump (using a QAD ring).

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


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Figure 23 SCL

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AGB Front View Page: 47


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

ACCESSORY GEARBOX GROUP Components The accessory gearbox external components and features (left side) are: Three ports that are used for aircraft instrumentation and for system checks. They are: . The LV3 -- Lubrication Vent 3 (breather) . The LP5 -- Lubrication Pressure 5 (oil filter inlet) . And the LP6 -- Lubrication Pressure 5 (oil filter outlet). NOTE: ON MANY AIRCRAFT, LV3 IS NOT USED; IT IS CAPPED. The accessory gearbox external components and features (right side) are: -- An overboard breather port to which a QEC breather duct is attached -- A N2 tachometer drive pad to which a QEC tachometer is attached -- And a engine identification plate. The accessory gearbox external components and features (rear) are: -- A mount pad to attach the QEC starter -- A mount pad to attach Constant Speed Drive (CSD) generator -- And a mount pad to attach the QEC hydraulic pump. The oil system components include: -- An oil pressure regulating valve -- An oil pump (pressure and scavenge oil) -- An oil filter -- A chip detectors (optional). -- And a deoiler (inside of the accessory gearbox). THESE OIL SYSTEM COMPONENTS ARE LRU‘S EXCEPT FOR THE DEOILER.


NOTE:

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


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Figure 24 SCL

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AGB Rear View Page: 49


POWER PLANT ENGINE

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

AIRFLOWS Purpose: Provides cooling for the hot section, thrust balance for front and rear compressors, and pressurizes the bearlng compartments. 6th Stage Air ( Ps3): Bleeds N1 compressor discharge pressure into compressor case annulus and through the 4th stage disk into the interior of the front compressor. Two airframe service bleed ports. Pressurization: -- No. 1, 2 and 6 bearing seals Cooling: -- No. 6 bearing housing -- Tailcone -- Exhaust case struts -- Rear face 3rd stage turbine disk -- Front and rear face of the 4th stage turbine disk Sensing: -- Pressure Ratio Bleed Control Buffer air to the No. 3 bearing labyrinth seal


8th Stage Air: 8th stage bleeds through openings in the 8th stage stator into an annulus in the intermediate case. Two airframe service bleed ports. Anti--ice: -- Inlet case -- Nose cone Anti--surge bleed valves up to 3 Buffer air to the No. 4 bearing (labyrinth seal only) 9th Stage Mr: Bleeds through holes in 9th and 10th stage spacer into rear compressor bore. Thrust balance: -- Reduces thrust load on No. 4 bearing Pressurization: -- No. 3 bearing seal

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13th Stage Mr (Ps4): Bleeds N2 compressor discharge pressure through diff case struts into circumferential annulus. 2 airframe service bleed ports, and 1 or 2 anti--surge bleed ports. De--icing air: -- Fuel heat Sensing air: -- Fuel control unit Actuation pressure: -- Muscle pressure for pressure ratio bleed control -- Muscle pressure to bleed control valve Cooling: -- Fuel nozzles -- Combustion chamber liners -- Inner and outer outlet ducts -- 1st stage turbine nozzle guide vanes -- 1st stage turbine outer air seal -- 1st stage turbine blades (--15 thru --17AR) Pressure Coollng Pressure (P): Bleeds across 13th stage disk rear knife edge airseal Cooling: -- No. 4 & 5 compartment heatshields -- Front and rear faces of the 1st and 2nd stage turbirie disks -- Front face of 3th stage turbine disk -- 1st stage turbine blades (--11 only) -- Interior of the diff case Pressurization: -- No. 4 bearing carbon seal -- No. 4 1/2 carbon seal -- No. 5 carbon seal

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BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00


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Figure 25 SCL

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Engine Airflows Page: 51

BOEING 737 - 200 JT8D -- 7 to 17 72 -- 00

NOTES:



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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL SYSTEM



POWER PLANT OIL SYSTEM

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL SYSTEM General The oil system: -- Supplies pressurized oil to cool, lubricate and clean the engine bearings and ry drives -- Collects oil from the bearing compartment and accessory drives and sends that oil back to the oil tank. -- Controls the air breather pressure in the bearing compartments.


Description and Operation: The oil tank supplies oil to the oil pump, and the pump sends pressurized oil through the oil filter. The filtered oil flows to the oil pressure regulating valve which: -- Byes part of the oil back to the oil pump if it is necessary to maintain the oil flow to the bearings -- Sends the non--byed oil to the fuel/oil cooler. The fuel/oil cooler transfers heat from the oil to the fuel (which flows through tubes in the fuel/oil cooler). Then the cooled, pressurized oil goes to the bearing compartments and to the gearbox. Scavenge pumps return the oil from the bearing compartments and gearbox to the oil tank. Breather air flows from the bearing compartments to the accessory gearbox. Oil particles are removed from it, and then the oil--free air is vented overboard. (This is not shown below.) Oil System Limits: Normal oil pressure: 40 to 55 psi. Temperature 120 °C JT8D--1, --7, --9 (continuous) 157 °C JT8D--1, --7, --9 (transient) 130 °C JT8D--11, --15, --17 (continuous) 165 °C JT8D--11, --15, --17 (Transient) Breather 8.0 inch Hg, maximum JT8D--1, --7, --9 8.5 inch Hg, maximum JT8D--11, --15, --17 7.13 inch Hg, maximum all models with No. 4 carbon seal SCL

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Components: * The external components are as follows: -- Oil tank (attached to the left front face of the AGB) -- Oil pump (at the bottom of the AGB) -- Oil filter (at the lower left side of the AGB) Three micron sizes are available. -- Oil pressure regulating valve (at the rear of the AGB) -- Fuel/oil cooler (between flanges B and D at the 7:00 oclock) -- Magnetic chip detectors (optional) * The internal components are as follows (not shown): -- Deaerator (in the oil tank) -- Deoiler (in the accessory gearbox) -- Oil filter by valve (in the oil filter) -- Four scavenge pumps (gearbox at bearing 1, 4and 5, and 6) -- Strainers and nozzles (at different places in the oil flowpath) Other Data: * The oil specifications are as follows: -- Specification PWA 521 tells the type of synthetic oil to be used in the JT8D. -- Service Bulletin 238 lists the approved oils by brand name. * The oil consumption guidelines are as follows: -- If the No. 4 bearing seal is labyrinth oil consumption is .2 to 2 quarts / hour. -- If the No. 4 bearing seal is carbon oil consumption is .1 to 1 quart / hour. Maintenance: The following components are LRUs: -- Oil tank --- Oil filter --- Fuel/oil cooler and by valve -- Scavenge pump in gearbox -- Scavenge pump for No. 1 bearing

Oil pump Oil pressure regulating valve -- Magnetic chip detectors. -- Scavenge pump for No. 6 bearing

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 1 SCL

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Oil System Overview Page: 3



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

ENGINE LUBRICATION SYSTEM General The engine lubrication system provides oil for the bearings and accessory drives. It consists of a: -- pressure system, -- scavenge system and -- breather system. Components The oil tank is located on forward left side of accessory drive case. Pressure and scavenge pump, filter, filter by valve and pressure regulator are inside the accessory drive case. Oil cooler with by valve on left side of engine, above oil tank. Seven bearings, last chance filters and scavenge pumps located inside the engine.


Bearings There are 7 bearing assemblies located through the engine. These are numbered 1 through 6.

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Bearing Designation Table Name Location Number Front Compressor Front Inlet case; Front 1 compressor front hub Front Compressor Rear Compressor case 2 (intermediate) front; Front compressor rear hub Rear Compressor Front Compressor case 3 (intermediate) rear; Main accessory drive gear Rear Compressor Rear Diff case; Rear 4 compressor, rear hub Turbine Intershaft In line with midpoint 4 1/2 of combustion chamber case; outer race within rear compressor drive turbine shaft; Inner race race and rollers on front compressor drive turbine shaft Turbine Front In line with combustion 5 chamber case rear flange; Inner race on rear compressor drive turbine shaft Turbine Rear Turbine exhaust case; 6 Front compressor drive turbine rear hub

Type Roller

Duplex Ball

Ball

Duplex Ball Roller

Roller

Roller

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

Nr 2

Nr 3

Nr 4

Nr 4 1/2

Nr 6 Nr 5


Nr 1

Figure 2 SCL

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Bearing Locations Page: 5



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

ENGINE LUBRICATION SYSTEM Pressure System Oil is gravity fed from oil tank into the main oil pump within the accessory drive case. The pressure section of the pump forces oil through the filter located downstream of the pump. If the filter becomes clogged, a filter by valve opens allowing oil to flow to the oil cooler. Fuel is used to cool the engine oil. If the cooler is blocked, a by valve across the cooler opens. Oil leaves the cooler and flows through tubing and last change filters to the bearing compartments. An oil pressure sense line from the cooler is connected to the pressure regulator which maintains a constant oil pressure at the bearing jets, regardless of the pressure drop of the oil at the fuel oil cooler. Scavenge System Oil is scavenged from the bearing compartments by three scavenge pumps which return the oil to the drive case. The scavenge pump in the drive case sends the oil back to the tank.


Breather System To ensure proper oil flow and to maintain proper scavenge pump performance, the pressure in the bearing cavities is controlled by the breather system. Breather lines are provided from bearing cavities to the accessory drive case. In the accessory case vapor--laden atmosphere es through a rotary breather impellers mounted on the starter drive shaft. Oil is removed and oil free air is discharged overboard through a hole in left cowling.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 3 SCL

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Oil Circuits Page: 7



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

ENGINE EXTERNAL OIL DISTRIBUTION General The engine oil distribution system consists of a pressure system which supplies lubrication to the engine bearings, accessory drives and a scavenge system. A breather system interconnects the individual bearing compartments, the accessory drive gearbox, and the oil tank. Oil Pressure System Oil flows by gravity from the oil tank to the engine driven pump, located inside the accessory drive case. Pressure oil from the pump flows through an oil filter to a fuel cooled oil cooler and es to the various engine bearings. By valves are provided in the filter and in the oil cooler. An adjustable pressure regulating valve, installed in the accessory drive case maintains system pressure and flow bying oil back to the pump inlet. Scavenge Oil System Scavenge pump return oil from the bearing cavities to a sump in the accessory drive case. The scavenge stage of the engine driven pump then returns the oil to the tank. Oil Breather System An oil breather system interconnect the engine bearing cavities, the accessory drive case, and the oil storage tank.

EXTERNAL COMPONENTS


Detection Optional magnetic chip detectors can be installed in the scavenge lines, oil tank and accessory case.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 4 SCL

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External Components Page: 9



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

ENGINE OIL TANK General The engine is provide with a single oil tank used to store and supply oil to the oil pressure subsystem for engine lubrication. The oil tank is located on the left forward face of the accessory drive case. It is held at the drive case by three bolts and is secured at the front by a strap.

Maintenance: The oil tank is an LRU. The oil drain valve is cleanable and replaceable.


Components The tank consists of oil outlet tube to pressure pump and oil return tube from scavenge pump. The returning oil is sent to a deaerator where oil and air/fumes are separated by centrifugal action of swirling oil. The air/fumes mixture is vented to the accessory drive case and hence overboard through the breather system. The tank vent system consists of a climb and dive/normal ports, thus providing an all attitude venting system. Quantity indication is provided by a capacitance probe and by a dipstick which is integral with the filler cap. On the left side of the tank are provisions for optional sight glasses and pressure servicing ports. Description and Operation: The tank is a stainless steel weidment. It contains three transfer tubes which accommodate oil supply, scavenge and breather flow between the tank and main accessory gearbox. A self--locking filler cap and attached dipstick are located in the main filler port. A scupper sunip and drain system provides overboard drainage. Remote fill ports are provided for a remote filling option. An internal baffle minimizes oil sloshing in tank. The tank contains provision for an oil quantity transmitter. An internal deaerator removes air from scavenge return oil. Sight gage ports provide for optional visual method of checking oil level. The tank incorporates a drain valve. Tank capacities: -- Total: 6.3 U.S. gallons -- Serviceable: 5.0 U.S. gallons (20 quarts, 18.9 liters) -- Usable: 4.0 U.S. gallons (16 quarts, 15.1 liters) The oil should be checked and added as soon as possible after engine shutdown but no later than two hours after engine shutdown. Oil servicing can be accomplished through the filler port or the remote fill ports.

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OIL TANK

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 5 SCL

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Oil Tank Components Page: 11



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

FUEL OIL COOLER General Automatic engine oil cooling is provided by a fuel--oil cooler on the high pressure side of the oil system. The oil cooler is of the full flow type, there is full flow of fuel through the cooler whenever the engine is running and no temperature regulation of engine oil is accomplished within the cooler. The oil cooler is located on forward left side of engine above the oil tank.


Components The oil cooler consists of an outer case which houses a tube assembly through which fuel es. Oil es back and forth across the tubes to provide maximum heat transfer. A pressure by valve is provided only on the oil side of the cooler. Indication of opening of the by valve is by the increasing oil temperature. ”0” ring seals are located on the forward end of the case and on the ”0” ring holder to prevent fuel and oil leakage. A small leakage hole is located on the forward bottom part of the case and connects into the cavity between the two ”0” rings. No leakage of fuel is permissible from the hole. At the outlet from the cooler, pressure sense line supplies oil pressure signal to the regulator. Description and ()peration: The fuel/oil cooler is of the full flow type with a pressure by valve: -- It consists of a housing containing a removable core composed of more than 200 straw--like tubes through which fuel es. -- A series of baffles within the core directs the flow of oil around the tubes. The by valve opens at 55 psid in the event that the core becomes clogged and to ensure that oil flow to the bearing compartments is not restricted. Typical pressure drop across the cooler is 30 psi. The fuel/oil cooler has a 040 inch diameter weep hole, the leakage of fluid (fuel or oil) through this hole is an indication of an internal packing failure.

FUEL OIL COOLER

Maintenance: The fuel/oil cooler is a line replaceable uint. The by vaive is a line replaceable unit, and it is cleanable.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 6 SCL

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Fuel Oil Cooler Page: 13



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL QUANTITY COMPONENTS Purpose The Oil Quantity indicating system is used to display on the flight deck the quantity of oil, in quarts, gallons or liter units remaining in the oil tank. Components The oil quantity system components consist of: -- Oil tank -- Oil quantity transmitter -- Oil quantity indicator Description and Operation The oil quantity transmitter is located in the oil tank. There are two types: -- Reed switch potentiometer and magnetic--float assembly -- Capacitance probe The capacitor type probe has a capacitor in a housing. The capacitance is in proportion to the oil level in the tank. The capacitance is transmitted to the flight deck oil quantity indicator. A reed type has a set of switches that open and close by a magnetic float. The float moves with the oil inside a housing. The reed switches change the resistance in proportion to the oil level and transmit it to the flight deck oil quantity indicator. The oil quantity system is a QEC component.


Maintenance The oil quantity indicating system components are LRU‘s.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 7 SCL

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Oil Quantity Components Page: 15



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL QUANTITY INDICATING General The oil quantity gaging system measures the oil quantity in the engine oil tank and provides a visual indication in the control cabin for monitoring oil consumption.

Test Operating the test switch, connects the active area of tank unit to ground, thus simulating empty condition. This results in the gage pointer moving downscale. Upon release of the test switch, the gage returns to previous r

Components The gaging system consists of a capacitance type combination tank unit and compensator located in the oil tank, an oil quantity gage and a test switch on the P2 center instrument . The gage receives 115 Vac power from P6--2 circuit breaker . Tank Unit and Compensator The tank unit consists of two conductive tubular separated by an air gap. The active gaging area on these tubes is profiled, or varied, to provide an accurate reading of quantity at various oil levels and attitudes. A compensator section, composed of two short tubular separated by an air gap, is mounted at the lower end of the unit. It is completely immersed in oil and minimizes the effect of change in oil dielectric due to temperature or variation in composition.


Indicator The indicator consists of an amplifier--bridge assembly, bridge circuit, power supply transformer and two phase motor, which through gearing drives the pointer and rehajancing part of the bridge. On the rear of the indicator are the FULL and EMPTY adjustment screws. The indicator has integral lighting. Operation The tank unit acts as a variable capacitor, the capacitance of which varies with a change in the dielectric value of the material between its plates. In the oil tank, the dielectric material will be all oil or partially oil. and partially air, the ratio being determined by the amount of oil in the tank. Thus the capacitance of the tank unit will vary with the amount of oil in the tank. The capacitance change Is transmitted to an amplifier bridge which converts the signal to an amplified voltage used to drive an indicator which displays the quantity of oil in the tank.The motor also drives the potentiometer wiper to rebalance the bridge network.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 8 SCL

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Oil Quantity Indicating Page: 17



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL TEMPERATURE COMPONENTS Purpose The Oil Temperature indicating system is used to display, on the flight deck, the engine oil temperature. Components The oil temperature system consists of: -- Oil temperature sensor -- Oil temperature indicator Description and Operation The oil temperature sensor, a resistance type bulb located in the fuelloil cooler, senses oil temperature at the oil cooler outlet. The oil temperature indicator, located on the flight deck, indicates temperature in °C. The maximum continuous oil temperature is: -- 120 °C for the JT8D--1, --7, and --9 model engines and -- 130 °C for the JT8D--11, --15, and --17 model engines. Indicating system are QEC components.


Maintenance Indicator and sensor are LRU‘s.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 9 SCL

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Oil Temperature Components Page: 19



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL TEMPERATURE INDICATING General The oil temperature system monitors the oil temperature in the distribution system and provides a visual indication in the control cabin. Components The system consists of a resistance type temperature bulb located at the outlet from the oil cooler and oil temperature indicator on P2 center instrument . The indicator receives 28 Vac power from P6--2 circuit breaker . Temperature Bulb The oil temperature bulb comprises an enclosed resistance unit whose resistance varies directly with changes in temperature. It is in direct with oil flow leaving the cooler. Indicator The oil temperature indicator is a resistance ratiometer type unit. It consists of no--power sweep off coil, deflection coil, cabin temperature compensator and a dial calibrated in °C. The indicator has integral lighting.


Operation The indicator rectifies the AC power to DC and applies it to the metering circuits through an electro--magnetic sweep--off mechanism. This forces the meter to respond to the coil torques. The indicator applies a voltage to the temperature bulb and the resulting current is a measure of the oil temperature. The bulb resistance as a function of the oil temperature, determines the voltage applied to sensing coil of the ratiometer bridge. The quick release of the bulb provides for easy trouble shooting with an ice bath or by ambient temperature comparison.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 10 SCL

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Oil Temperature Indicating Page: 21



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL PRESSURE COMPONENTS Purpose The Oil Pressure indicating system is used to display on the flight deck the engine oil pressure in pounds per square inch (psi). Components The oil pressure system consists of: -- Oil pressure transmitter -- Oil pressure indicator -- Oil pressure transmitter restricter


Description and Operation The oil pressure transmitter is a synchro--magnetic type transmitter located on the fan exit case at the 10:00 position. The oil pressure indicator is located on the flight deck’s engine indicator and displays engine oil system pressure in psi (normal oil pressure range is 40--55 psi). The oil pressure transmitter restricter prevents the transmitter from sensing oil pressure fluctuations. The transmitter can be vented to the engine’s breather pressure. A low oil pressure switch sends a signal to the flight deck low oil pressure light when the oil pressure decreases to approximately 35 psi. Aditionally the pressure system has an advisory light from the oil filter when it becomes to block by oil contamination aproximately at 34 -- 38 PSID. Is posible during cold start the advisory light is comes on for a few seconds. The indicating system are QEC components. Maintenance The oil pressure Tx, pressure indicator and low oil pressure sw are LRU‘s.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 11 SCL

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Oil Pressure Components Page: 23



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL PRESSURE INDICATING General The oil pressure indicating circuit provides visual display in the control cabin of engine oil pressure, low oil pressure warning and oil filter by warning. Transmitter The oil pressure transmitter consists of a diaphragm connected to a synchro transmitter. The diaphragm senses oil pump output pressure on one side and local ambient on the other side. Indicator The indicator consists of a synchro receiver, a gear train, pointer and a dial calibrated in psig. The indicator has integral lighting.


Low Oil Pressure Warning The low oil pressure switch senses oil supply pressure. With low supply pressure, LOW OIL PRESSURE amber light has ground through the oil pressure switch. Oil Filter By Warning The oil filter differential pressure switch senses filter inlet and outlet pressures. When the filter begins to block, the differential pressure increases, OIL FILTER BY amber light has ground through switch. Operation Engine oil pressure and vent pressure are applied to opposite sides of the diaphragm in the oil pressure transmitter. The pressure differential positions the diaphragm and in turn, the rotor position of a synchro transmitter. The rotor of the oil pressure indicator’s synchro receiver assures a position corresponding to the transmitter rotor, since the stators of both are connected in parallel. The synchro receiver rotor drives the oil pressure indicator pointer. If power is interrupted, the indicator will remain at the last pressure indication. Oil pressure indicator, LOW OIL PRESSURE and OIL FILTER BY amber lights are on P2 center instrument .

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Figure 12 SCL

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Oil Pressure Indicating Page: 25

BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL PRESSURE SYSTEM General The oil pressure system using an oil pump increase the oil pressure, considering the engine drives the pump and the different engine power setting, an oil regulator pressure valve maintains the oil pressure values not more than 55 psi, an oil pressure transmitter on forward left side of engine monitors the pressure rates. Filtered oil goes to fuel oil cooler to reduce high temperature, then the oil flows through the last chance filters before sprayed the differents components as bearings and gears. An oil filter differential switch on left side of accessory drive case, becomes active if debris or oil contamination through the filter is sensed and oil filter by light comes on . As LOP warning, if the pressure decreasing lower than 35 psi, a low oil pressure switch, very near to the pressure transmitter, is active and Low Oil Pressure light comes on.




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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 13 SCL

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Pressure System Schematic Page: 27



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

MAIN OIL PUMP Purpose: The main oil pump supplies oil under pressure to the engine bearings and main accessory gearbox. The main oil pump also returns scavenge oil to the oil tank. Location: The main oil pump is bolted into the bottom of the main accessory gearbox. Description and Operation: The main oil pump is a single stage positive--displacement gear--type: -- Dual--element pump * Pressure * Scavenge. Pump discharge pressure varies with N2 speed. The upper element in the oil pump is the pressure stage. It pressurizes the oil to send it through the pressure subsystem. The lower element in the oil pump is the scavenge stage. It sends scavenge oil to the oil tank, and it is part of the scavenge subsystem. The pump discharge (output) pressure varies with the N2 speed. Maintenance: The oil pump is an LRU. THE DRIVE GEAR CAN FAIL OUT IF IT IS NOT HELD WHEN THE OIL PUMP IS TURNED OVER.


NOTE:

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Figure 14 SCL

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Oil Pump Page: 29



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

MAIN OIL FILTER General: The main oil filter traps solid contaminants thereby removing them from the engine lubricating oil. Location: The main oil filter is located in the lower left side of the accessory drive gearbox. Description and Operation: The following three micron filters can be used. -- A 15 micron, disposable filter -- A 40 micron, surface type filter that can be ultrasonically cleaned in a solvent. -- A 175 micron, stacked--screen type filter that can be dissassembled and washed in a solvent. THE FILTER BY VALVE IS INTERNAL WITH 40 MICRON AND 175 MICRON FILTERS AND IS ATTACHED TO THE COVER WHEN USING A 15 MICRON FILTER. A filter by valve opens when the pressure differential between the inlet and the outlet exceeds 70 psid. It is recommended that oil filters be removed for inspection between 250 and 400 hour intervals. The differential pressure switch sensing ports are located on the filter/gearbox housing. -- Inlet LP5 -- Outlet LP6 Filter cover is secured by 4 sleeve nuts and have a drain provision.


NOTE:

Maintenance: The oil filter is a line replaceable unit, A filter kit is available and a By valve is cleanable. THE 40 MICRON FILTER CANNOT BE INSTALLED IN SOME OF THE EARLIER JT8D ACCESSORY GEARBOXES. Test ports: -- LV3 Lubrication Vent 3 (breather pressure in accessory gearbox) -- LP5 Lubrication Pressure 5 (oil pressure before filter) -- LP6 Lubrication Pressure 6 (oil pressure after filter). NOTE:

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Figure 15 SCL

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Main Oil Filter Page: 31



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

OIL PRESSURE REGULATING VALVE Purpose: The oil pressure regulating valve regulates oil pressure to ensure proper flow through the oil nozzles. Location: The oil pressure regulating valve is threaded vertically into the rear of the main accessory gearbox. Description and Operation: The oil pressure regulating valve keeps the oil pressure at 40 to 55 psi at the oil nozzles. To do this, it has an internal piston that moves up or down to change the oil pressure. When the piston is all the way up, it blocks the by exit, so no oil is byed. Thus all of the oil goes to the fuel/oil cooler without a reduction in pressure. As the piston moves down, it permits some of the oil to go out through the by exit. The byed oil goes back to the oil pump inlet. This decreases the pressure of the oil that goes to the fuel/oil cooler.

Maintenance: The oil pressure regulating valve is an LRU. It can also be adjusted and cleaned. To set the adjustment screw at the bottom of the oil pressure regulating valve: -- Turn it clockwise to increase the spring force and increase the oil pressure. -- Turn it counterclockwise to decrease the spring force and decrease the oil pressure. NOTE:

ONE COMPLETE TURN OF THE ADJUSTMENT SCREW CHANGES THE PRESSURE BY APPROXIMATELY 2 PSI.

DURING ENGINE OPERATION, THE PISTON IS NEVER FULLY CLOSED. THUS, SOME OIL IS ALWAYS BYED. The up--and--down movement of the piston is caused as follows: -- The spring pushes the piston up. The adjustment screw can be turned to change the spring force. -- The sense pressure from the outlet side of the fuel/oil cooler pushes the piston down. If this pressure is high, more oil is byed, and the downstream oil pressure decreases. -- The breather pressure from the accessory gearbox pushes the piston up. If this pressure is high, less oil is byed, and the downstream oil pressure increases. -- The space above the piston is filled with oil that goes through a filter. This balances the force of the oil on the piston, and it dampens the movement of the piston.


NOTE:

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Figure 16 SCL

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Oil Pressure Regulating Valve Page: 33



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

SCAVANGE SYSTEM Purpose: The oil scavenge subsystem collects oil from the bearing compartments and accessory gearbox and sends it back to the engine oil tank. Description and Operation: The flow sequence for the oil scavenge subsystem is as follows: -- A gear--type scavenge pump sends the oil from the number 1 bearing compartment to the accessory gearbox. -- Oil from the number 2 and 3 bearing compartment drains (by gravity) through the tower shaft housing to the accessory gearbox. -- A gear--type scavenge pump sends the oil from the number 6 bearing compartment through an internal tube to the sump for the number 4, 4 1/2, and 5 bearing compartments. -- The scavenge oil from the number 4, 4 1/2, 5, and 6 bearing compartments sump is sent by a dual--element, gear--type scavenge pump to the oil tank through an internal age in the gearbox. -- The scavenge pump in the accessory gearbox sends the gearbox sump oil to the oil tank. -- When the scavenge oil enters the oil tank, it goes through a tube to a deaerator that removes air from the oil. -- A JT8D engine can have provisions for four magnetic chip detectors. The chip detectors are inserted into self--closing valves so they can be removed without oil leakage. The chip detector locations are: -- In the scavenge line from the No. 1 bearing; on the accessory gearbox. -- In the scavenge line from the No. 4, 4 1/2, 5, and 6 bearings; where that scavenge line comes out of the diff outer fan duct. -- In the accessory gearbox sump. -- In the scavenge oil age to the oil tank; on the accessory gearbox. NOTE:

THE JT8D OIL SYSTEM IS REFERRED TO AS A ”HOT TANK” SYSTEM BECAUSE HOT SCAVENGE OIL IS RETURNED TO THE OIL TANK. THUS THE OIL THAT IS SUPPLIED TO THE OIL PUMP IS HOT.

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Figure 17 SCL

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Scavange System Schematic Page: 35



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

SCAVENGE PUMPS Purpose: The scavenge pumps return oil and entrapped air from the bearing compartments and main accessory drive sump to the oil tank. Location: The scavenge pumps can be found in the following locations: -- Inside the front accessory drive cover (No. 1 bearing) -- Accessory drive gearbox (No. 2 and 3 bearing and gearbox) -- No. 4 and 5 bearing compartment (dual element) -- No. 6 bearing compartment Description and Operation Gear type pumps scavenge the oil and air from bearing compartments and main accessory drive gearbox at all operating conditions and return the oil and air mixture to the oil tank.


Maintenance: * Three of the scavenge pumps are LRUs. They are: -- No. 1 bearing scavenge pump -- No. 6 bearing scavenge pump -- Oil pump (which includes the accessory gearbox scavenge pump).

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Figure 18 SCL

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Scavange Pump Page: 37



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

SCAVENGE OIL DEAERATOR General: The scavenge oil deaerator separates the air from scavenge oil that is returned to the oil tank. Location: The scavenge oil deaerator is attached to the outlet of the scavenge oil return line inside the oil tank. Description and Operation: The deaerator is a stationary canister with an air vent at the top and an oil outlet at the bottom. A mixture of oil and air is returned to the oil tank by the scavenge oil pumps. The scavenged oil/air mixture is discharged tangentially onto the inner wall of the deaerator thereby reducing its velocity and pressure. This permits the entrapped air to separate from the oil and escape through the top of the deaerator into the oil tank air space. The deaerator oil drains out the bottom of the deaerator into the tank to be recirculated by the main oil pump.


Maintenance: Not a line replaceable unit.

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Figure 19 SCL

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Oil Tank Deaerator Page: 39



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

BREATHER SYSTEM Purpose: The breather subsystem prevents excessive air pressure in the bearing compartments so that the flow of oil to the bearings, and the operation of the scavenge systern is not impaired. The breather pressure also affects the operation of the oil pressure regulating valve which controls the oil pressure. Description and Operation: A controlled amount of air from the primary gaspath leaks past the bearing compartment airseals into the bearing compartments. This air: -- Prevents the leakage of oil into the gaspath -- Pressurizes the bearing compartments -- Mixes with the very small oil particles in the bearing compartments. The ”air/oil mist” is referred to as breather air. The word ”breather” is used because the air goes into the bearing compartments and is then expelled, which is similar to breathing. Breather pressure is required for proper operation of the scavenge pumps. The breather air flows from the bearing compartments to the accessory gearbox as follows: -- The breather air from the number 1 bearing compartment goes through an external breather tube to the accessory gearbox. -- The breather air from the number 6 bearing compartment goes through an internal tube to the No. 4, 4 1/2, and 5 bearing area. Then the breather air from the No. 4, 4 1/2, 5, and 6 bearing compartments goes through an external breather tube to the accessory gearbox. -- The breather air from the No. 2 and 3 bearing compartments goes through the tower shaft to the accessory gearbox. The accessory gearbox also gets air through an internal age from the oil tank. (This is the air that is removed from the scavenge oil by the deaerator). In the accessory gearbox, this air mixes with the breather air from the bearing compartments.

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The air/oil mist in the gearbox es through a deoiler that separates the air from the oil. Then: -- The oil (with the air removed from it) drains from the deoiler to the gearbox sump. It is pumped to the oil tank by the scavenge stage of the oil pump. -- The air (with the oil removed from it) is sent from the deoiler through the gearbox breather port. Then it flows through the airframe ducts and is vented overboard.

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Figure 20 SCL

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Breather System Schematic Page: 41



BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

MAIN ACCESSORY GEARBOX DEOILER General: The deoiler separates oil particles from the breather air. Another name for the deoiler is the ”gearbox breather impeller”. Location: Two deoiler impellers are located on the starter gear shaft in the main accessory drive gearbox. The breather vent is located on the left side of the main accessory gearbox case. Description and Operation: The deoiler consists of two impellers with side wall entry for the breather air/oil mixture. Breather air enters the side wall of the de--oiler. The centrifugal force created by the impeller forces oil particles to the outer rim of the impeller where it drains into the gearbox. Breather air pressure is lighter than the oil and escapes through the hollow shaft in the middle of the deoiler and on to the overboard vent port.


Maintenance: The front of the accessory gearbox can have a port which gives access to the deoiler carbon seal assembly, without removing the gearbox. See SB 4539.

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BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 21 SCL

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Main Accessory Gearbox Deoiler Page: 43


BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 22 SCL

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Oil System Review Page: 44


BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00



Figure 23 SCL

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Oil System Troubleshooting Chart Page: 45

BOEING 737 - 200 JT8D -- 7 to 17 79 -- 00

NOTES:




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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL SYSTEM



POWER PLANT FUEL SYSTEM

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

ENGINE FUEL SYSTEM DIAGRAM General The engine fuel distribution and control system consists of an engine--driven pump, control, anti--icing system, pressurizing and dump valve and a split manifold, delivering fuel to nine individual fuel nozzles.

* Pump Fuel is supplied from the tanks through the engine fuel shutoff valves to the engine--driven fuel pump assembly. From here it is pumped to the fuel control where it is metered in the proper quantity. Excess fuel is returned to the pump. A fuel filter is integral with the pump.

* Heat System The fuel heat system is located between the boost and main stages of the engine--driven fuel pump and consists of an air--fuel heater, air shutoff valve and differential pressure switch. The differential pressure switch provides a means of indicating, in the control cabin, icing conditions or a clogged filter.


* Control The fuel control is provided with two control levers; one to control engine starting and shutdown the engine speed during all forward and reverse thrust operations, and the other to control the fuel control accurately governs the steady state selected speed, acceleration and deceleration, and it indirectly governs the maximum turbine temperature of the engine during both forward and reverse thrust operation. * Pressurizing and Dump Valve From the fuel control, the fuel flows through the fuel flowmeter and the fuel oil cooler to the pressurizing valve. The pressurizing valve schedules flow to the secondary fuel nozzles as a function of pressure drop across the primary nozzles. * Manifolds and Fuel Nozzles The divided fuel flow from the pressurizing valve is delivered through the annular duct to the dual--tube fuel manifolds mounted on the diff case. The fuel then enters the nine fuel nozzles.

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Figure 1 SCL

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Engine Fuel System Diagram Page: 3



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL SYSTEM Purpose The fuel system supplies clean, metered fuel to the combustion chambers during all conditions of engine operation. Description and Operation: All airframe boost pump sends the fuel from the airplane fuel tank to the engine pump inlet. The fuel flows through an impeller boost stage in the fuel pump, which increases its pressure. Then it goes to the fuel deicing heater. Thirteenth stage compressor air can be sent to the fuel heater by the fuel deicing air shutoff valve and actuator. That air contain the heat that is tranferred to the fuel in the fuel heater to melt ice in the fuel. The fuel flows from the fuel heater back to the fuel pump, where it goes through the (main) fuel filter. The filtered fuel goes through the main gear stage in the fuel pump. Then the fuel goes from the pump into the fuel control which: -- Schedules the correct fuel flow rate and pressure that is necessary for combustion, and sends the metered fuel to the fuel flow tranmitter -- Sends the unneeded fuel back to the fuel pump. The metered fuel flows through the fuel flow transmitter which sends flow rate data to the aircraft flight deck. The fuel goes from the fuel flow transmitter to the fuel/oil cooler, where it cool the oil. The fuel goes from the fuel/oil cooler to the Pressurizing and Dump (P&D) valve. This valve divides the fuel into primary and secondary fuel and sends it to the fuel manifolds. The fuel flows through the primary and secondary fuel manifolds and goes into the nine fuel nozzles. It is atomized as it goes into the combustion chambers. Fuel Specification: The type of fuel specilied for use in the JT8D is identified by: -- Specifcation PWA 522 -- PWA Service Bulletin 2016.

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 2 SCL

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Fuel System Overview Page: 5



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

ENGINE LEFT SIDE: Components Fuel/Oil Cooler -- Located between flanges B and D at the 8:00 o‘clock position, attached to flange B. Fuel Flow Transmitter -- Located under the fuel/oil cooler, attached to the fuel/oil cooler.

Maintenance: The following components are LRU‘s: -- Fuel/oil cooler -- Fuel flow transmitter -- Fuel heater -- P&D valve -- External fuel manifolds.

Fuel Heater -- Located behind between flanges C and E at the 6:00 o‘clock position; attached to the left side of the fuel pump. (The fuel heater is shown in the left side view, but it is under the engine.) NOTE:

THE TUBING WITH 13TH STAGE AIR GOES TO THE FUEL HEATER, WHICH IS ON THE RIGHT SIDE OF FLIE ENGINE.

Tt2 Probe -- Located at the 7:00 position in the inlet case -- Connected to the fuel control, which is on the right side of the engine.


Pressurizing and Dump (P&D) Valve -- Located between flanges II and J at the 7:30 position; attached to the fan discharge diff outer duct. Fuel Manifolds -- Two short fuel manifold tubes come out of the P&D valve and are divided into two pairs of tubes. -- One pair of tubes goes into the dif outer fan duct at the 7:00 position. The other pair of tubes goes into the diff outer fan duct at the 5:00 position. -- The two pairs of fuel manifold tubes go from the diff fan duct to the diff case. They are connected to the fuel manifolds which are installed on the diff case. Fuel Nozzles and s -- Five fuel nozzle and assemblies; located internally.

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 3 SCL

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Left Side Fuel System Component Locations Page: 7



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

ENGINE RIGHT SIDE: Components Fuel Pump -- Located between flanges C and F at the 5:30 o‘clock position, attached to the front of the accessory gearbox. Fuel Control Unit (FCU) -- Located between flanges B and D at the 5:30 o‘clock position, attached to the front of the fuel pump.

Maintenance: The following components are LRU‘s: -- Fuel pump -- Fuel control Unit -- Fuel deicing air shutoff valve and actuator -- Moisture trap -- External fuel manifolds.

Moisture Trap (Condensation Trap) -- Located on the front of the FCU. Fuel Deicing Air Shutoff Valve and Actuator -- Located between flanges D and F at the 1:30 o‘clock position, attached to flange E. External Tubing -- Two tubes with 13th stage air come out of the fan discharge diff outer duct at the 1:00 and 11:00 o‘clock positions. They connect to one tube at the top of the engine, and it goes to the fuel heater.


-- A tube with Ps4 air comes out of the fan discharge combustion outer duct at the 12:30 o‘clock position. It goes to the FCU. ( This tube is routed differently on different JT8D models.) Fuel Manifolds -- Two short fuel manifold tubes come out of the P&D valve and are divided into two pairs of tubes. -- One pair of tubes goes into the diff outer fan duct at the 5:00 o‘clock position. The other pair of tubes goes into the diff outer fan duct at the 7:00 o‘clock position. -- The two pairs of fuel manifold tubes go from the diff outer fan duct to the diff case. They are connected to fuel manifolds which are installed on the diff case. Fuel Nozzles and s -- Four fuel nozzle and assemblies, located internally.

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Figure 4 SCL

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Right Side Fuel System Components Page: 9



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL PUMP The fuel pump provides pressurized fuel to the fuel control unit. Location: The fuel pump is mounted on the front, right side of the main accessory gearbox.


Description and Operation: The fuel pump consists of a single element gear stage with a centrifugal boost stage. A relief valve is incorporated to limit the pressure rise across the gear stage. The unit provides a rigid mounting pad for the fuel control unit. An integral fuel filter containing a replaceable filter element is located between the discharge of the centrifugal stage and the inlet of the gear stage. Should the pressure drop across the filter exceed a predetermined limit, a by valve directs flow into the gear stage. A mounting pad is provided on the filter housing to permit the use of a differential pressure warning device. An accessible and removable cover forms the lower portion of the sump area of the filter housing. In the event of a malfunction of the boost stage, a by valve opens into the inlet age of the pump to direct flow into the gear stage. This is held normally closed by a light spring force and remains closed due to boost stage pressure. Outlet and return ports are provided between the boost stage discharge and the filter inlet for installation of an external fuel heater. A driveshaft seal drain is located in the lower extremity of the mounting flange. Maintenance: Line replaceable unit Fuel pump -- Filter 40 micron, disposable -- Filter Delta P (AP) switch.

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Figure 5 SCL

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Fuel Pump Page: 11



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

ENGINE FUEL DE-- ICING SYSTEM General With engine operating, the two stage fuel pump is driven by the engine through the accessory drive case. Fuel enters primary-- stage, es through the heater, which normally does not operate, enters the filter, then secondary stage of the pump to fuel control unit.


Components The engine fuel de--icing system consists of: A. Filter icing switch (differential pressure Switch), located at top of filter. B. Fuel heater valve, motor operated, located on right side of engine. C. Fuel heater, located at forward bottom center of the engine. D. FILTER ICING amber light, VALVE OPEN blue light , HEAT two position switch and associated electrical circuits located in Fuel Module P5 forward overhead .

If the fuel heater is blocked on the fuel side, by valve opens, fuel byes heater. This failure is evidenced by no rise in oil temperature when fuel heat is used. If the filter remains blocked, by valve opens, fuel byes filter. This failure is evidenced by the FILTER ICING amber light on P5 remaining illuminated when fuel heat is used. If the output from the pump assembly is blocked, by valve opens, fuel circulates around secondary stage.

Operation Filter Icing Indication. At low fuel temperature, water in fuel will be in the form of ice collecting on the filter causing-- its blockage. With high fuel differential pressure across the filter, switch closes and the FILTER ICING amber light, two MASTER CAUTION lights and FUEL annunciator light are illuminated. Heater Operation. Wtth the filter blocked due to ice, opening of the fuel heat valve, allows 13th stage engine bleed air to through the heater. The air exhausts overboard through a hole in left cowling. Warm fuel will melt ice on filter. Duty cycle for fuel heat if fuel tank temperature is below 0 °C is 1 minute on, 30 minutes off, or as required. Fuel heat system should not be used for prolonged time on the ground, since hot fuel will not provide the necessary oil cooling. Fuel By System There are 4 by valves in the fuel system to ensure fuel flow to the engine in case of primary stage failure, heater blockage, filter blockage or output blockage. If the pump primary stage fails, by valve opens, allowing fuel to proceed directly to the secondary stage. This failure is evidenced by no rise in oil temperature when fuel heat is used.

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Figure 6 SCL

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Fuel De-- icing Schematic Page: 13



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

ENGINE FUEL DE--ICING CIRCUIT GeneraI The fuel de-- Icing ( heating) system is monitored and controlled from P5--2 Fuel Module on P5 . This consists of two lights and a switch, receiving 28 Vdc Batt Bus 1 and 2 power from P6--3 circuit breaker . Filter Icing Indication The filter icing Indicatjon circuit consists of 28 Vdc power supply from P6--3 , FILTER ICING amber Iight on P5 and differont ial pressure switch located on top of filter. When fuel differential pressure across the filter rises, at 4 . 4 to 5. 8 psid , switch closes. FILTER, ICING, FUEL annunciator and two MASTER CAUTION amber lights are illuminate. At decreasjng differential pressure of 4.0 to 2.0 psid, switch opens and lights are extinguished.

The VALVE OPEN light operates as follows: A. Valve closed -- no light B. Valve open -- dim blue C. Valve in transit -- bright blue D. With power on, electrical connector to valve removed -- bright blue Dim blue light when ground is through zener diode and transistor. Bright blue light when ground is through both transistors.


Valve and Timer Operation The valve is motor operated powered by 28 Vdc from P6--3 and is located on right side of engine. Open/closed position indicator is located on valve body. The timer consists of a solid state 60 second dual time delay circuit and a holding solenoid in P5--2 fuel module. Placing the fuel heat switch to ON, 28 Vdc is supplied to valve motor, valve opens allowing compressor discharge air to fuel heater. 28 Vdc energizes the 60 second timer and solenoid, thus hoIding switch in ON posit ion. after 60 seconds, solenoid is de--energized by timer, switch returns to OFF position, 28 Vdc is supplied to close valve. If necessary, switch can be manually placed to OFF prior to 60 seconds. Valve Position Light Indication The VALVE OPEN blue light circuit consists of 28 Vdc power supply from P6--3 and printed circuit card in P5--2 module. The circuit senses 28 Vdc supp]y to the valve motor, if 28 Vdc is present, transistor is biased, if 28 Vdc is absent transistor is not biased.

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 7 SCL

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Fuel De-- icing Control Page: 15



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL DE--ICING AIR SHUTOFF VALVE Purpose: The fuel deicing air shutoff valve and actuator controls flow of 13th stage compressor air to the fuel heater. Location: The fuel deicing air shutoff valve and actuator is mounted at the 2:00 o‘clock position on the fan duct on the engine right side. Description and Operation: The fuel deicing system consiste primarily of an air--fuel heater, air shutoff valve differential fluid pressure switch and the necessary tubing. The pressure drop warning switch mounted on the fuel filter indicates when the filter is iced. When the cockpit fuel heater switch is activated, an electrically actuated air shutoff valve located in the bleed air line at the inlet of the heater opens to permit 13th stage air to flow through the heater. The fuel heater and filter are installed in the fuel system between the boost and main stages of the engine drive fuel pump. AII of the engine fuel flow es through heater at all times. The fuel is heated however, only when the air shutoff valve is opened, allowing high compressor discharge air to flow through the air side of the heater. Actuator has mechanical position indicator.


Maintenance: The fuel deicing air shutoff valve and actuator assembly is an LRU.

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 8 SCL

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Fuel De-- Icing SOV Page: 17



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL DEICING HEATER The fuel heater heats fuel pump interstage fuel to protect the fuel system from ice crystals. Location: The fuel heater is mounted on the fuel pump housing at the 6:00 o‘clock position.


Description and Operation: The fuel heater is located between the boost and main stages of the engine driven fuel pump. The fuel heater functions as an air--fuel heat exchanger to protect the engine fuel system from ice. It consists of a housing containing a core composed of over 200 straw--like tubes through which compressor bleed air es and around which the entire engine fuel flow is circulated; a series of baffles within the core which direct the flow of fuel around the tubes so that the fuel is uniformly heated; and a by valve which permits the fuel flow in the event of clogging. The fuel heater uses 13th stage air as a source of heat, and functions only when the air shutoff valve is open, allowing high compressor discharge air to flow through the air side of the heater. Operation of the fuel heater is controlled manually. A dífferential pressure switch on the fuel deicing filter activates a warning light in the cockpit when there is a pressure drop across the filter caused by ice or clogging. A fuel heater switch can then be actuated, to open the fuel heater air shutoff valve. Engine bleed air es through the tubes of the heater, warming the fuel which is baffled around these tubes. The resulting warm fuel will melt any ice within the filter and the WARNING light will go out as the pressure drop across the filter is decreased. The heater is used intermittently. Maintenance: The fuel heater is an LRU. Be sure the restrictor is installed before you attach the exhaust duct (a QEC item) to the fuel heater.

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 9 SCL

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Fuel De-- Icing Page: 19



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL DEICING CONTROL AND INDICATION Purpose: The fuel deicing controls are used to energize the fuel deicing system. The fuel deicing indication lights: -- Warn the flight crew that the differential pressure switch on the fuel filter has closed. -- Tell the status (condition) of the fuel deicing air shutoff valve and actuator. NOTE:

THE FUEL DEICING SYSTEM IS ALSO REFERRED TO AS THE FUEL HEAT SYSTEM IN AIRCRAFT MANUALS.

Location: Flight deck P5 overhead . Description and Operation: Activation of the fuel heat selector switch provides power to the actuator motor to open the 13th stage air valve illuminating the fuel heat valve light (blue). Some models have an optional 60 second timer. Recommended cycle: -- “ON” is 60 seconds. -- “0FF” is 30 minutes.

Indication: The fuel deicing indication lights operate differently in different aircraft. There are usually two lights: -- The caution light goes on when the fuel filter differential pressure increases. -- On different aircraft, the fuel deicing status light can: . Come on while the deicing valve is open and go off when it closes, or . Come on only while the deicing valve is ”in transition” (when it goes from closed to open, or from open to closed), or . Comes on brightly while the deicing valve goes from closed to open; stay dim while the valve is open; the go off when the valve closes.

Control Switches: There is one fuel heat switch for each engine. When the flight crew turns on the fuel heat switch for an engine: -- The fuel deicing actuator is energized to open the air shutoff valve, so 13th stage air can flow to the fuel heater. The fuel heat switch has an optional timer that turns it off after 60 seconds. THE FLIGHT CREW CAN TURN THE FUEL HEAT SWITCH OFF BEFORE 60 SECONDS. When the switch is turned off, the fuel deicing valve closes, so 13th stage air does not go to the fuel heater.


NOTE:

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 10 SCL

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De-- Icing Control and Indications Page: 21



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL CONTROL Purpose: The fuel control (FCU) supplies the correctly metered fuel flow that is necessary for all conditions of engine operation, including: -- Start--up -- Steady--state (idle, cruise ,etc.) -- Acceleration -- Deceleration. Description: The FCU “schedules” the fuel flow as a result of the throttle postion (power lever angle) and engine sense inputs. It has a hydro--mechanical computing section and a metering section. The computing section controls the operation of the metering section. To do so, the computing section uses: -- Inputs from the flight deck and from some engine sensors -- Hydraulic power from the metering section. The rnetering section of the fuel control gets fuel from the fuel pump gear stage, and it sends some of that fuel (metered pressure) to the fuel flow transmitter. The metering section also sends: -- The unwanted fuel back to the gear stage inlet -- A small amount of fuel to the computing section.

Only two line maintenance adjustments are permitted on the fuel control unit. -- The idle trim adjusting screw is used to set the N2 idle speed to a value that is specified in a trim table for the ambient temperature. -- The part power trim adjustment screw is used to set the EPR to a value that is speclfied in a trim table for the ambient temperature and pressure. The part power trim adjustment is also referred to as the mil (military) or max trim. NOTE:

THE PART POWER TRIM STOP IS USED AS A FIXED REFERENCE (STOP POSITION) FOR THE THROTTLE LEVER WHILE THIS ADJUSTMENT IS PERFORMED. IT MUST BE IN THE CORRECT POSITION, AS SPECIFIED IN THE TRIM TABLE.


Maintenance: The fuel control is an LRU. The fuel control main filter group includes the main (149 micron) and servo (40 micron) fiiters. This assembly can be removed and cleaned. The Ps4 bellows is serviced with silicone oil. The oil prevents condensation from entering the bellows. The drain plug in the bottom of the moisture trap can be removed to: -- Drain accumulated moisture -- Inspect the .020 inch bleed hole in the plug to make sure it does not have a blockage.

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Figure 11 SCL

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Fuel Control Unit Page: 23



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

Operation: The metering section consists of the following: -- The fuel from the pump gear stage comes into the FCU and flows through the coarse filter. (The coarse filter has a by valve.) -- A small quantity of this fuel goes through the fine filter. It supplies hydraulic power for the servo mechanisms in the computer section. -- The pump discharge pressure fuel goes from the coarse filter to the throttle (metering) valve. But before it goes into that valve, some of it may be sent back to the pump interstage by the pressure regulating valve. -- The metered fuel goes from the throttle valve to the minimum pressure and shutoff valve. The position of this valve (open or closed) is controlled by the fuel shutoff lever. The valve is moved by a spring and by fuel pressure. -- The metered fuel goes out of the FCU and flows to the down stream components in the fuel system. The computing section consists of the following: -- Fuel shutoff lever on the flight deck is connected by linkage to the control This input opens or closes a servo to control pressure behind the shutoff valve to hydraulically lock the shutoff valve closed or permit it to open.

-- PLA, (throttle and reverse throttles) are inputs from the flight deck requesting speed (power level). -- RTT (Reserve Takeoff Thrust) only on JT8D--17R or AR This system operates automatically during takeoff; if an engine loses power all electrical signal is sent to the FCU solenoid to increase the thrust level in the engines.

-- Ps4 (high compressor discharge pressure) air goes from the end of the compressor through external tubing to the FCU. It goes through the moisture (condensation) trap bolted on the FCU to a bellows assembly in the FCU. It is used to maintain constant ratios of Wf/Ps4. It is a primary signal with high authority, going directly to the throttle valve.


--

Tt2 sensor sends inlet temperature through an external tube that divides into two capillary tubes (covered with a flexible conduit) these tubes are connected to a bellows chamber in the FCU. The signal biases the schedule at: . Idle to allow for different air density

. Snap acceleration to maintain a constant response time. -- N2 signal comes from a splined driveshaft of the fuel pump which turns a speed governor with flyweights to maintain the requested speed. If this input is lost due to malfunction, the FCU schedule goes to the zero speed protective circuit which is a high cruise power.

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 12 SCL

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Fuel Control Unit Schematic Page: 25



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL FLOWMETER General The fuel flow indicating system provides a visual indication in the control cabin of the fuel consumption of each engine. Components The system consists of a single Power Supply Module (M316), located on the E3--3 Electronic Equipment rack, flow transmitter located on forward left side of engine and indicator calibrated in lb/hr or kg/hr on P2 center instrument . Power supply is 115 Vac and 28 Vdc from P6 circuit breaker .

Indicator The fuel flow indicator consists of a permanent magnet rotor surrounded by a coil which is supplied with 115 Vac from P6--3 circuit breaker , a pointer and a dial face calibrated in lb/hr or kg/hr. The coil picks up the output signal from the transmitter, rotates the permanent magnet in indicator by the same angular displacement as the magnet in transmitter. The pointer gives a visual indication of fue l flow during engine operation. The indicator has integral illumination.


Operation The power supply converts 28 Vdc to 3 ∅, 4Hz, 17 Vac which is supplied to the flowmeter transmitters. The conversion is accomplished by timer switches and oven controlled oscillator. Transmitter The transmitter consists of a motor driven impeller and a fuel flow driven turbine. An impeller motor is located upstream of the impeller. A signal transmitting unit is located downstream of the turbine and is attached to the turbine with two restraining springs of different load characteristics. This permits the use of an indicator with tine index marks at the low side and coarser index marks at the high side resulting in greater sensitivity at low flow rates. The impeller motor receives 3 phase, 4 Hertz, 17 volts AC from PSM and is driven at constant 240 RPM, the impeller is driven through reduction gears at constant 60 RPM regardless of fuel flow. The fuel ing through the impeller is given a swirling motion and then es through the turbine. The act.ion of the fuel on the turbine blades produces a torque, against the restraining springs, which is directly proportional to the fuel flow rate. Rotation ol the restraining springs rotates a permanent magnet rotor surrounded by a coil which is supplied with 115 Vac from P6--3 circuit breaker . It is recoimnended that the FUEL FLOW INDICATOR circuit breaker be pulled (open) whenever the engine fuel feed line is drained or the airplane is out of service for maintenance. This will prnvent damage to the transmitter from dry operation and will prolong its life.

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Figure 13 SCL

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Fuel Flow Meter Page: 27



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL/OIL COOLER Purpose The fuel/oil cooler maintains oil temperature within operating limits by transferrig heat from the oil to the fuel. Location: The cooler is mounted on B and C flanges at the 8:00 o‘clock position on the front compressor case. Description and Operation: THIS DESCRIPTION OF THE FUEL/OIL COOLER IS SHORT. A MORE DETAILED DESCRIPTION IS GIVEN IN THE ”OIL SYSTEM” SECTION. The fuel enters the fuel/oil cooler and flows through thin parallel tubes. Then it goes out of the fuel/oil cooler to the P&D valve. The oil flows around the tubes and through a set of baffles. This causes good heat transfer from the oil to the fuel. Then the cooled oil goes out of the fuel/oil cooler. The fuel/oil cooler has a .060 inch diameter weep hole. The leakage of fluid (fuel or oil) through this hole is an indication of an internal packing failure. NOTE:

Maintenance: -- Line replaceable unit -- By valve is removable and cleanable


. Byes at 55 psid -- .060 inch weep hole, inspect for leakage -- Test port for measuring inlet oil pressure.

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Figure 14 SCL

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Fuel Oil Cooler Page: 29



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

PRESSURIZING AND DUMP (P&D) VALVE Purpose The P&D valve is a flow divider between primary and secondary fuel flows. Location: The P&D valve is mounted on the fan diff case at the 8:00 o‘clock position. Description: Fuel enters inlet, es through a 74 micron screen, then opens a dump valve, which is acting as a check valve, proceeding to the prirnary manifoid outlet. Above idle, approaching the power range, fuel acting on a pressurizing valve opens to allow fuel through the secondary manifold.


Maintenance: The P&D valve is a LRU. The fuel inlet strainer is an LRU, and it can be cleaned. There are two test ports: -- FP4 (primary fuel pressure) -- FP5 (secondary fuel pressure).

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 15 SCL

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P&D Valve Page: 31



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

Operation: Fuel enters the P&D valve and goes through an inlet screen. If that screen is clogged, the pressure differential across it increases; then it unseats so the fuel can by it. The fuel collects in the filter chamber, and its pressure increases. This makes the fuel push against the dump valve, which opens it. The fuel flows through the dump valve and goes into the primary chamber. Then it goes out of the P&D valve’s primary outlet, into the primary manifolds. If the pressure increases in the prirnary chamber due to an increased fuel flow, the pressurizing valve starts to open. This permits fuel to flow into the secondary chamber. Then it goes out of the P&D valve’s secondary outlet, into the secondary manifolds. NOTE:

ON MOST AIRCRAFT, THE DUMP VALVE PORT IS SEALED WITH A PLUG FOR PROTECTION OF THE ENVIRONMENT.


Modes of Operation: The fuel pressurizing and dump valve has three modes of operation: -- When the engine is shut down, the dump valve and pressurizing valve are closed. The dump valve does not permit fuel to flow back to the fuel/oil cooler, flowmeter, and fuel control. Thus it is a check valve in this condition. -- From engine start--up to idle, the fuel flows to the primary chamber and goes to the primary manifolds. The pressurizing valve stays closed, so fuel does not go into the secondary chamber. -- Above idle, the pressurizing valve opens, and fuel does go into the secondary chamber and out to the secondary manifolds. Thus the fuel nozzles get both primary and secondary fuel.

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Figure 16 SCL

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P&D Valve Operating Modes Page: 33



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL MANIFOLDS Purpose: The fuel manifolds deliver fuel from the pressurizing and dump valve to the fuel nozzles at nine locations around the diff case. Location: The fuel manifolds are mounted on the exterior of the primary diff case.


Description and Operation: The divided fuel flow (primary and secondary) from the fuel pressurizing and dump valve is delivered through two short external manifolds to two points on the exterior of the diff outer fan duct at the 5:00 and 7:00 o‘clock positions. These tubes lead through the duct to the left and right fuel manifolds which are mounted on the exterior of the diff case. The left manifolds supply fuel to the nozzles for combustion chambers six, seven, eight, nine and one. The right manifolds supply fuel to the nozzles for chamber numbers two, three, four and five.

Left and Right Manifolds

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 17 SCL

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Fuel Manifolds Page: 35



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL NOZZLE Purpose: The fuel nozzles atomize the fuel and spray it, with 13th stage air, into the combustion chambers. Description and Operation: There are 9 fuel nozzle and assemblies. They are inside of the diff case, in the primary gaspath. Each is attached to the inside of the diff case. A nozzle has fittings that are attached to the primary and secondary fuel manifolds on the outside of the diíf case. The fuel nozzle nut is threaded onto the nozzle , and the combustion chamber is installed over it. Thus the fuel nozzle nut goes into the front of the burner can. The fuel nozzle nut functions as a nozzle. Thirteenth stage air ( from the diff) goes into it and is made to swirl (turn). Then that air goes into the combustion chamber. The fuel nozzle nut holds a fuel nozzle assembly which is put into the nut before the nut is attached to (threaded onto) the . This inner nozzle: -- Has strainers and a metering plug for the primary and secondary fuel. -- Makes the fuel swirl, atomizes, and sends it into the combustion chamber. In some conflgurations, the inner nozzle also gets thirteenth stage air, makes it swirl, and mixes it with the fuel when it is sprayed into the combustion chamber. Other Data: The different JT8D models have several different types of fuel nozzles. The fuel nozzles have been re--designed to decrease emissions. The two major types of fuel nozzles are shown on the facing page. They are: -- The low emissions type which is newer, went into production January 1, 1984, and can be retrofitted by SB 5650. -- The type of fuel nozzle used prior to the low emission type.

Primary and Secondary Flow

A Duplex Type. When both orifices are delivering fuel, their output is blended into a single spray. SCL

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Figure 18 SCL

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Typical Fuel Nozzles Page: 37



BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00

FUEL FLOW SEQUENCE: The aircraft boost pumps send fuel from the fuel tank to the inlet of the engine’s fuel pump. -- Engine fuel pump inlet pressure must be at least 5 psi to ensure satisfactory pump life. -- Typical pump inlet pressure is 15 -- 25 psi from the aircraft boost pumps. -- If the inlet pressure is as low as 5 psi, the low pressure switch sends an electrical signal which turns on the INLET FUEL PRESS LOW caution light on the flight deck. The fuel flows through the (impeller) boost stage which increases its pressure by approxirnately 10 to 60 psi from start to takeoff power. -- If there is a boost stage failure, the nP across the impeller by valve increases. -- The by valve opens, this permits the fuel to flow through it, directly to the pump’s gear stage. (If this occurs, the fuel does not go through the heater and main fuel filter) The fuel goes from the fuel pump boost stage to the fuel heater. The fuel is heated as it es through the fuel heater if the fuel deicing system is turned on. -- If the pressure drop across the fuel heater is large (due to an internal blockage), the fuel byes the fuel heater. The fuel flows from the fuel heater to the fuel filter in the fuel pump. If the filter becomes clogged by dirt or ice particles, the nP across the filter increases. -- If nP increases the difrerential pressure (caution) switch sends an electrical signal which turns on a caution light on the flight deck. This warns the flight crew to turn on the fuel de--icing system. To do so, they turn on the fuel heat switch to energize the actuator that opens the fuel de--icing air shutoff valve. That permits thirteenth stage compressor air to flow to the fuel heater and increase the temperature of the fuel. -- If nP increases to a higher value, the filter by-- valve opens. This permits the fuel from the fuel heater to go directly to the gear stage, without going through the filter. NOTE:

THE FUEL THAT GOES FROM THE BOOST STAGE, THROUGH THE FUEL HEATER AND FILTER, IS CALLED INTERSTAGE FUEL.

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The fuel goes from the fuel filter to the pump gear stage, which increases its pressure to: -- Approximately 150 psi when the engine is motored -- Approximately 900 psi (1,000 psi for JT8D--17/R/AR ) at TO power. If the output pressure of the gear stage increases to more than 950 psi (1,050 psi for JT8D--17/R/AR engines), the high pressure relief valve opens. That lets some of the fuel flow back to the gear stage inlet. The fuel goes from the gear stage, out of the pump, into the fuel control unit. The fuel control unit receives inputs that tell it: -- Some of the engine operatlng conditions -- The power level requested by the flight crew. . Schedules the correct fuel flow rate and pressure that is necessary for combustion, and sends the metered fuel to the fuel flow transmitter . Sends the unneeded fuel back to the gear stage in the fuel pump. THE FUEL THAT GOES FROM THE FUEL CONTROL UNIT (THROUGH THE F/F TRANSMITTER, FUEL/OIL COOLER, AND P&D VALVE) TO THE FUEL NOZZLES IS CALLED METERED FUEL. The metered fuel goes through the fuel flow transmitter which sends flow rate data to the aircraft flight deck. Then it es through the fuel/oil cooler which reduces the oil temperature. The fuel goes from the fuel/oil cooler to the P&D (Pressurizing and Dump) valve. When it enters the P&D valve, it goes through a filter. That filter can be byed if it is clogged. NOTE:

The P&D valve divides the fuel into primary and secondary fuel flows and sends it to the fuel manifolds. During engine start--up and idle, it only sends fuel to the primary manifolds. Above idle (at higher power levels), it also sends fuel to the secondary manifolds. The fuel goes from the primary and secondary fuel manifolds, through the nozzle s, into the nine fuel nozzles. It is atomized and sent into the combustion chambers.

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BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 19 SCL

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Fuel Flow Sequence Page: 39


BOEING 737 - 200 JT8D -- 7 to 17 73 -- 00



Figure 20 SCL

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Fuel System Troubleshooting Chart Page: 40

BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

AIR SYSTEM



POWER PLANT AIR SYSTEM

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BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

AIR SYSTEM General The Engine Air System purpose is divide in two targets: A. The Pneumatic System (Bleed Air). B. The Internal Airflows (Anti Surge).


Bleed Air Compressed air from the 8th-- and 13th--stage engine compressor are bleeds to supply the aircraft pneumatic system for diferent uses such as: -- Wing anti--ice ATA 30 -- Inlet cowl anti--ice ATA 30 -- Cabin air conditioning ATA 21 -- Cabin pressure ATA 21 -- Engine start ATA 80 -- Generator cooling ATA 24 -- CSD oil cooling ATA 24 -- Hydraulic tank pressurization ATA 29 -- Water tank pressurization ATA 38 This air bleed is obtained through 8th stage Check Valve and 13th stage Modulating and SOV, a precooler serve to reduce the temperature and a Engine Bleed SOV to control the bleed air extraction. This system will be completelly comment in the ATA 36 of this training.

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Figure 1 SCL

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Typical Engine Bleed Page: 3



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

ANTISURGE BLEED SYSTEM Purpose. The antisurge bleed system decreases the possibility of compressor surge. The engine is designed for maximum efficiency and stability during high power operation (ifight conditions). The system helps to prevent surge during low power operation (start, idle, and deceleration).


Components: The following components make up the antisurge bleed system: -- Pt2 probe -- Pressure Ratio Bleed Control (PRBC) -- Ps3 strainer -- Antisurge bleed valves -- Start bleed control valve (SB 5425) Pneumatic Operating Flows: The antisurge bleed system includes the following pneumatic operating flows: -- Pt2 -- compressor inlet total pressure -- Ps3 -- 6th stage air static pressure -- Ps4 -- 13th stage air static pressure to the PRBC -- Ps4 muscle -- Ps4 muscle pressure to the bleed valves.

Configurations JT8D engines have three different antisurge bleed system configurations. Two of them are prior to SB 5425, and the other one is as specified in SB 5425. -- The SB 5425 configuration uses a start bleed control valve (SBCV) and a pressure ratio bleed control (PRBC) to control the antisurge bleed valves. -- The two pre--SB 5425 configurations do not have a SBCV. They use a PRBC to control all of the bleed valves.

Description and Operation: The anti--surge bleed system senses N1 inlet (Pt2) pressure and N1 outlet (Ps3) pressure to schedule the opening and closing of the antisurge bleed valves. The system is automatic. It is not controlled from the flight deck. The antisurge bleed system permits some of the 8th and 13th stage high compressor air to go out (through bleed valves) into the fan duct. This reduces the ”back pressure” in the primary gaspath, so that air can more easily flow through the compressor.

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BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00



Figure 2 SCL

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Antisurge Operation Page: 5



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

Components Location: The 8th stage antisurge bleed valves are on the compressor case. The 13th stage bleed valves are on the diff case. The Pt2 (Inlet Pressure) Probe is in the inlet case at the 5:30 position. The Pressure Ratio Bleed Control (PRBC) is on the diff outer fan duct at the 4:30 position. The Start Bleed Control Valve (SBCV) is on the diff outer fan duct, at flange H. It is usually at (or near) the 3:00 to 3:30 position. The SBCV is only on SB 5425 engines. The Ps3 strainer (filter) is on the diff outer fan duct at the 11:00 position. NOTE:

THE ROUTING OF THE EXTERNAL TUBING, AND THE LOCATIONS OF SOME OF THE COMPONENTS, ARE DIFFERENT ON DIFFERENT AIRCRAFT MODELS.


Maintenance: The following components are line replaceable: -- Pt2 probe -- Pressure ratio bleed control -- Start bleed control valve -- Ps3 strainer -- External plumbing (tubes, fittings, etc.). The (8th and 13th stage) antisurge bleed valves are not line replaceable. All of the lines, tubing, connectors, and components must be attached tightly and leak--free. The bleed system components should be cleaned as specified in the maintenance manuals.

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BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00



Figure 3 SCL

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Antisurge Component Locations Page: 7



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

PRESSURE RATIO BLEED CONTROL Purpose: The PRBC senses the pressure ratio across the front compressor and schedules bleed valves to prevent surge. Location: The PRBC is mounted on the diff outer fan duct at the 4:30 position. Description and Operation The PRBC is: -- Pneumatically controlled and operated -- Three cleanable screens in the vent ports -- When the engine is static, the bleed valves are normally open.


Maintenance: * The PRBC is line replaceable. * The three vent screens are cleanable. * The venturi area is cleaned as specified in the manual.

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Figure 4 SCL

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PRBC Valve Page: 9



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

BLEED VALVE Purpose: The bleed valves allow high compressor air to bleed into secondary gaspath when opened. Location: The bleed valves are mounted internally on diff case and compressor case. Description and Operation: A bleed valve contains a piston that moves freely in the bleed valve housing. When the engine is static (shut down), the position of this piston is a result of: -- Its position when the engine was shut down -- The pull of gravity and the drag of the piston against the inner surface of the bleed valve. * During engine operation, the position of the piston in the bleed valve is controlled by the air pressures that push against it. -- When Ps4 actuating air (muscle pressure) is not sent to a bleed valve, the compressor air pushes the bleed valve’s piston open. This permits the HPC air to go (bleed) through the bleed valve, into the fan duct. -- When muscle pressure is sent to a bleed valve (by the PRBC or SBCV), it pushes the piston closed. This stops the flow of compressor bleed air to the secondary gaspath. A 13TH STAGE BLEED VALVE HAS PS4 AIR PUSHING ON BOTH SIDES OF ITS PISTON. BUT, THE MUSCLE PRESSURE HAS MORE FORCE BECAUSE IT PUSHES AGAINST A LARGER AREA THAN THE AIR FROM THE PRIMARY GASPATH.


NOTE:

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BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00


7 2


Figure 5 SCL

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Antisurge Bleed Operation Page: 11



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

OPERATION PRE SB 5425: * Engine start and idle is: -- During the engine start anti idle operation, compressor air pushes the bleed valves open when the muscle valve in the PRBC is (1) closed to the flow of Ps4, and (2) opened to atmospheric vent. * Above idle is: -- As the engine is accelerated above idle, the increasing Ps3 overcomes the combined force of Pt2 and spring. This forces the poppet valves to reverse position, allowing Ps4 to reposition the muscle valve. The repositioned muscle valve allows Ps4 muscle pressure to flow to the compressor bleed valves to force them closed. * Engine shutdown is: -- With the engine shut down there are no active forces on the compressor bleed valves. Since they are open at idle power they would remain in the open position after engine shutdown.

OPERATION POST SB 5425


As RPM increases during engine start, Ps4 pressure opens the start bleed control valve and closes the 13th stage bleed valve. The 13th stage bleed valve remains closed until the engine is shut down. When the engine is shutdown, Ps4 pressure decreases to close the start bleed control valve and open the 13th stage bleed valve.

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Figure 6 SCL

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Antisurge Bleed Schematic Post SB Page: 13



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

CONFIGURATIONS: The SB 5425 not incorporated engines are: * There are two antisurge bleed system configurations in these engines. In both of these configurations: -- All of the bleed valves are controlled by the PRBC. There is no start bleed control valve. -- There are two 13th stage bleed valves. * The two configurations differ in one way: -- In the earlier configuration, there is no 8th stage bleed valve. -- In the later configuration, there is one 8th stage bleed valve. (All of the JT8D engines built after January, 1976 had this 8th stage bleed valve.) SB 4597 WAS ISSUED IN MAY, 1976. IT OFFERS THE ADDITION OF THE 8TH STAGE BLEED VALVE TO THE EARLIER ENGINES THAT DID NOT HAVE IT. The SB 5425 incorporated engines are: -- The one 13th stage bleed valve is controlled by the SBCV. -- The three 8th stage antisurge bleed valves are controlled by the PRBC.


NOTE:

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Figure 7 SCL

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Antisurge Bleed Configurations Page: 15



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

START BLEED CONTROL VALVE Purpose: To close the 13th stage bleed valve before the engine reaches idle. Location: The start bleed control valve is mounted on the fan diff case at the 3:30 position. Description and Operation The pneumatic/spring poppet valve is: -- Static spring pressure closed In depressurized condition, the valve is closed and outlet port is vented to ambient. When pressure is applied to inlet port, the poppet moves off its seat and permits air to flow through the outlet port and close the bleed valve. Simultaneously, the ambient bleed holes are closed.


Maintenance: * Line replaceable unit * Cleanable

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Figure 8 SCL

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Start Bleed Control Valve Page: 17



BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00

ANTI SURGE BLEED General During an initial run the engine bleed valves are required to be checked for proper operation. The operation of the valves is a functionof engine inlet temperature (Tt2), engine inlet pressure (Pt2) and low pressure compressor speed (N1).


Procedure From the graphic, knowing Tt2 and Pt2, can be determine at wich N1 % the valves should close or open during engine acceleration and deceleration respectively. During acceleration, when valves close, a sudden increase in N1 % rpm of approximately 8 % and EPR of approximately 0.06 units will be observed. During deceleration, when valves open, a sudden decrease in N1 % rpm of approximately 8 % and EPR of approximately 0.07 units will be observed.

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BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00



Figure 9 SCL

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Antisurge Bleed Limits Page: 19


BOEING 737 - 200 JT8D -- 7 to -- 17 75 -- 00



Figure 10 SCL

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Air System Troubleshooting Chart Page: 20

BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

ANTI -- ICING



POWER PLANT ANTI-- ICING SYSTEM

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BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

ANTI--ICE SYSTEM General The 737 has three anti--icing air shutoff valves on each engine. Two of them control the flow of 8th stage air to the engine inlet, as described earlier. The third valve controls the flow of 13th stage air that prevents ice formation on the engine inlet cowl. The 13th stage air is mixed with cooler, lower pressure air before it goes to the cowl. There is one anti--icing toggle switch for each engine. In correct operation, all three anti--icing valves open when the pilot turns the switch ON. There is an indicator light for each of the three anti--icing valves on an engine. These indicator lights have the following names: -- COWL VALVE OPEN (blue) -- R VALVE OPEN (blue) -- L VALVE OPEN (blue). Description and Operation THIS DESCRIPTION APPLIES TO ENGINE NUMBER 1 SIMILAR DESCRIPTION CAN BE GIVEN FOR ENGINE NUMBER 2. When the pilot turns the antii--cing (ENG ANTI--ICE) switch ON: All three indicator lights for Engine No. 1 go on brightly to show that the valves are in transit. When all three valves are open, all three indicator lights for that engine become dim. But if a valve does not open, the corresponding indicator light stays bright.The condition of each of these lights can be summarized as follows: OFF if the switch is OFF and the valve is closed. ON dimly if the valve is open. ON brightly -- While the valve is in transit (opening or closing) -- If the valve position is different than the switch position.


NOTE:

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Figure 1 SCL

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Anti-- Ice System Page: 3



BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

ENGINE ANTI--ICE General During icing conditions, the Pt2 probe in the nose dome becomes blocked by ice. Air pressure trapped between nose dome and engine pressure ratio transmitter is now vented to the nose dome, thus reducing the Pt2 signal and causing engine pressure ratio (EPR) indicator, on P2 center Instrument , to rise. The engine anti--icing system prevents the formation of ice, or melts ice, on the: -- Fan inlet case -- Inlet guide vanes -- Inlet nose bullet (also referred to as the nose cone, or nose dome).

Maintenance All of the anti--icing system components are LRU‘s. Pratt & Whitney recommends that the sleeve strainer element be inspected and cleaned every 500 hours.


Components There are two anti--icing tube assemblies, one on each side of the engine. The routing of each assembly is from the fan--discharge rear--compressor outer duct, forward to the fan inlet case. There are two air shutoff valve and actuator assemblies, one on each side of the engine. They are located on the front compressor case at the 11:30 and 12:30 positions. Description and Operation When the anti--icing system for an engine is turned on from the flight deck: -- Both actuators are energized to open the air shutoff valves. -- Eighth stage bleed air flows through the tubes to the fan inlet case. -- The hot air flows into the outer annulus of the fan inlet case, through the hollow inlet guide vanes, and into the inner annulus. -- The air then flows into the nose bullet and is sent into the gaspath. The moisture that condenses in the outer annulus of the fan inlet case comes out through the sleeve strainer element at the bottom of that case. The screen prevents blockage of the weep hole. NOTE:

THE SLEEVE STRAINER ELEMENT IS ALSO REFERRED TO AS A WATER DRAIN PLUG OR CONDENSATION DRAIN.

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Figure 2 SCL

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Engine Anti -- Ice Page: 5



BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

CONTROL AND INDICATION Indication The engine anti--icing indicator lights give a visual indication of the operation and condition of the air shutoff valves to the flight crew. Location The engine anti--icing indicator lights are near the anti--icing control switches, on the forward overhead on the flight deck.

THE ANTI--ICING SWITCH ALSO OPENS (OR CLOSES) ONE OR MORE VALVES THAT CONTROL THE FLOW OF 13TH STAGE ANTI--ICING AIR TO THE NACELLE COWL. Those valves and tubing assemblies are QEC components. NOTE:

Description and Operation The anti--icing indicator lights tell the flight crew if: -- The anti--icing system is turned on or off -- The anti--icing valves are open or closed. The anti--icing indicator lights are different in different aircraft models. For example: -- For each engine, there can be one, two, or three indicator lights. -- The indicators can be rectangular or circular, and they can be amber, blue, or green. NOTE:

ACTUATOR HAS MECHANICAL POSITION INDICATOR.

Control The flight crew uses the engine anti--icing control switches to turn the anti--icing system ON or OFF.


Location The engine anti--icing switches are on the forward overhead on the flight deck. Description and Operation There is an anti--icing toggle switch for each engine. When the anti--icing switch is turned ON for an engine, the actuators are energized to open both of the anti--icing air shutoff valves on that engine. When the anti--icing switch is turned OFF, both of the anti--icing air shutoff valves close.

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Figure 3 SCL

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Control and Indication Page: 7



BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

ENGINE NOSE COWL ANTI--ICING General Downstream of nose cowl anti--ice valve is thermostatic valve, and in the nose cowl is an injector. Control and monitoring of the system is from P5, forward overhead .


Operation Placing engine anti--ice switch to the ON positions, allows three anti--icing valves to open. 13th stage engine bleed air es through the nose cowl anti--icing valve and a thermostatic regulator valve and discharges through the injector at the bottom of the nose cowl. Ambient air is drawn in by the injector, mixed with 13th stage bleed air and dissipated in the nose cowl leading edge through a perforated tube. The air is then discharged overboard through a port in the nose cowl. Engine and nose cowl anti--icing use should be limited to the minimum during systems checks on the ground.

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Figure 4 SCL

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Nose Cowl Anti-- Ice Page: 9



BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

NOSE COWL ANTI-- ICE REGULATOR VALVE General The nose cowl anti--ice system utilizes 13th stage engine bleed air. This air supply mass flow and temperature at high engine thrust operation is greater than required for the anti--icing of the nose cowl. Regulation of the air mass flow as a function of air temperature is provided by the regulator valve. Location The nose cowl anti--ice regulator valve is located on the forward right side of engine in the air duct downstream of the nose cowl anti--ice valve. Components The regulator valve consists of a sleeve enclosing a bi--metal temperature sensitive coil connected by a shaft to a rotary plate, a stationary plate, closed and open position stops.


Operation When the nose cowl anti--ice valve is open 13th stage engine bleed air es through the initially fully open regulator valve. The bi--metal temperature sensitive coil expands radially when subjected to the high air temperature causing the rotary plate to rotate against a stationary plate. This rotation reduces the relative openings between the plates, thus changing the flow area and hence air mass flow. Two stops are provided for the rotary plate, a closed stop and open stop, which is also used as a valve failure indicator. When the valve is cold, it should be fully open with the rotary plate resting on an open stop, thus preventing inward movement of the indicator pin. If the pin can be depressed, it indicates that the valve has failed in the partial or fully closed positions.

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BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00



Figure 5 SCL

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Nose Cowl Anti--Ice Regulator Valve Page: 11



BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

ENGINE AND NACELLE ANTI-- ICE CIRCUIT General The engine and nacelle anti--ice system is controlled from P5--11. Engine and Wing Anti--Ice Module on P5 . This consists of three lights and a switch receiving 28 Vdc and 115 Vac power from P18--3 circuit breaker . Valve Operation The cowl and two engine anti--ice valves are motor operated and are located on forward section of engine. Open/closed position indicator is located on each valve body. Placing the engine anti--ice switch to ON, 115 Vac is supplied to the three valve motors, valves open allowing compressor discharge air to anti--icing ducting in nose cowl and engine. Placing switch to OFF closes three valves.


Valve Position Light Indication The circuits for COWL VALVE OPEN, R. VALVE OPEN and L. VALVE OPEN blue lights are identical and consist of 28 Vdc power supply from P18--3 and printed circuit card in P5--11 module. The circuit senses 115 Vac supply to the valve motor, if 115 Vac is present, transistor is biased, if 115 Vac is absent, transistor is not biased. The lights operate as follows: A. Valve closed -- no light. B. Valve open -- dim blue. C. Valve in transit -- bright blue. D. With power on, electrical connector to valve removed -- bright blue. Dim blue light when ground is through zener and transistor. Bright blue light when ground is through both transistors.

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Figure 6 SCL

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Engine Anti -- Ice Control Circuit Page: 13

BOEING 737 - 200 JT8D -- 7 to -- 17 30 -- 00

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BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

ENGINE CONTROLS



POWER PLANT ENGINE CONTROLS

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BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

ENGINE CONTROLS General The engine control system transmitts control lever movement in the control cabin to the fuel control unit on the right side of the engine. The system consists of three sub--systems: A. Start system B. Forward thrust control C. Reverse thrust control Start System The start system consists of a start lever, connected to fuel control unit by control cables, push--pull cable and rod linkage. It energizes the ignition system and initiates fuel flow. Forward Thrust Control The forward thrust control system consists of a thrust lever, connected to fuel control unit by control cables, push--pull cable and rod linkage.

Thrust Reverser Follow--up Cable The thrust reverser follow--up cable connected between drum and shaft assembly and the reverser, prevents application of engine power whenever the thrust reverser is in transit. Control Cable Nomenclature A. ESS1A -- Engine start system, No. 1 engine, A cable. B. ESS1B -- Engine start system, No. 1 engine, B cable. C. ESS2A -- Engine start system, No. 2 engine, A cable. D. ESS2B -- Engine start system, No. 2 engine, B cable. E. T1A -- Thrust, No. 1 engine, A cable F. T1B -- Thrust, No. 1 engine, B cable G. T2A -- Thrust, No. 2 engine, A cable H. T2B -- Thrust, No. 2 engine, B cable

Reverse Thrust Control The reverse thrust control system consists of a reverse thrust lever connected to fuel control unit by the same control cables, push--pull cable and rod linkage as used by the forward thrust control.


Drum and Brake Assembly Drum and friction brake assemblies, located between floor beams at the aft end of the Electronic Equipment compartment, furnish braking for the forward and reverse thrust control cables. Drum and Shaft Assembly The drum and shaft assembly is mounted on the forward face of front spar. It consists of a start and thrust drums and two concentric shafts. It transmits control cable travel to push--pull cables on left side of engine. Cross Shaft The push--pull cables are attached to the left side of the concentric cross shaft which es under the engine to the right side. From the cross shaft, motion is transmitted to fuel control unit by push--pull rods.

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BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00



Figure 1 SCL

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Engine Controls Page: 3



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

THRUST AND START LEVERS General Engine start ing, thrust control and reverse thrust operation are performed by three levers located on the control pedestal in the control cabin. Start Lever The start lever provides ignition and fuel supply to the engine during starting. The lever is provided with a spring--loaded detent catch which is released by lilting the knob. The detent secures the lever in the CUTOFF and IDLE positions and is mechanically connected to a start drum on which are mounted cams for operation of ignition and fuel shutoff valve switches. ing over the drum is the engine start systun (ESS) cable which mechanically actuates the minimum pressure and shutoff valve in the fuel control unit.

The first movement, of reverse thrust lever opens a hydraulic valve which in turn provides hydraulic pressure to open the reverser. The lever is locked out by the thrust reverser interlock cable until the reverser is fully open. Further movement of the reverse thrust lever actuates the throttle valve in the fuel control unit. When reverse thrust is increased, a temperature indicating detent is encountered. The detent serves as a caution, it is possible to exceed allowable engine exhaust gas temperature limits if the lever is advanced beyond the detent. Movement of the lever forward, reduces engine thrust and closes the reverser.


Forward Thrust Lever The forward thrust lever provides variable fuel supply to thu engine, thus controlling the engine thrust. A lockout mechanisn prevents simultaneous actuation of the forward and reverse thrust levers to assure positive forward or reverse thrust control. The reverse thrust lever is attached to the forward thrust lever which in turn is concentrically mounted with a thrust drum. ing over the drum is the engine thrust (T) cable, which mechanically actuates the throttle valve in the fuel control unit when forward thrust lever is moved. The forward thrust lever idle position is against the idle stop on the control stand; and forward motion is terminated by a stop on the tuel control unit. The GO AROUND switches for the flight director are mounted on the thrust levers. Reverse Thrust Lever The reverse thrust lever actuates the thrust reverser and provides variable fuel supply to the engine, thus controlling the engine thrust during reverse thrust operation. A lockout mechanian prevents simultaneous actuation of the forward and reverse thrust levers to assure positive forward or reverse thrust operation. The reverse thrust lever is attached to the forward thrust lever. Actuation of the reverse thrust lever, rotates the thrust drum. (It should be noted that the direction of travel of thrust drum and cables is the same for decreasing forward thrust as it is for increasing reverse thrust.).

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Figure 2 SCL

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Thrust and Start Levers Page: 5



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

DRUM AND BRAKE INSTALLATION General Drum and friction brake assemblies are installed in the throttle control cable runs. Location The brake assenblies are located between the floor beams at the aft end of the E/E compartment. Components The assembly consists of two pulleys on the same shaft with a friction brake and spring. Cable T1A is wrapped around one pulley and T1B around the second pulley. Landing and takeoff warning switches are installed in the assembly.


Operation During manual operation the assemblies furnish system braking and act as idler pulleys. The friction brakes release when the thrust lever is in motion and reset when the thrust lever stops moving. Rig pins are installed in the assemblies for rigging of the control cables.

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Figure 3 SCL

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Drum and Brake Assy Page: 7



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

ENGINE MOUNTED CONTROLS General The engine mounted controls transmit the movement of the control cables for engine start system, thrust system and thrust reverser operation. Components The controls consist of a drum and shaft assembly mounted on the forward face of the front wing spar, extending through the firewall into the nacelle strut. Connected to the drum and shaft assembly are two push--pull cables, on the left side of the engine, which are connected to the cross shaft. On the right side of the engine are two rods connected between the cross shaft and fuel control unit. Part of the drum and shaft assembly consists of a thrust reverser selector valve, valve actuating cam, thrust reverser follow up crank and cam. Connected to the reverser follow up crank is a thrust reverser push--pull cable assembly with a quick disconnect.

When the reverser is fully open, the push--pull cable moves the interlock cam and releases the thrust drum. Further movement of the reverse thrust lever now actuates the fuel control unit through the thrust crank, push--pull cable, cross shaft and rod. For engine removal, on the left side of the engine the push--pull cables are disconnected from the cross shaft. Also the reverser push--pull cable is disconnected at a quick disconnect point in the strut.

Start System When the start lever, in the control cabin, is advanced from CUT--OFF to IDLE, the motion is transmitted to the start drum by ESS( )A cable. Rotation of the start drum moves the start crank and start push--pull cable which in turn rotates the cross shaft and by a start rod, the fuel control unit is actuated.


Thrust System When the thrust lever, in the control cabin, is advanced from IDLE, the motion is transmitted to the thrust drum by T( ) A cable. Rotation of the thrust drum moves the thrust crank and thrust push--pull cable, which in turn rotates the cross shaft and by a start rod, the fuel control unit is actuated. Reverse Thrust System When the reverse thrust lever, in the control cabin, is moved from stowed position, the motion is transmitted to the thrust drum by T B cable. Rotation of the thrust drum moves the selector valve actuating cam opening the selector valve. Hydraulic pressure is now supplied to the reverser system for opening the reverser. Further movement of thrust drum is restrained by the interlock and follow up cam. This cam is connected by crank and puqh--pull cable assembly to the reverser actuator.

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Figure 4 SCL

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Engine Mounted Controls Page: 9



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

THRUST REVERSER INTERLOCK LINKAGE Deploy Moving the reverse thrus lever in the control cabin rotates the thrust drum and control cam on the drum and shaft assembly clockwise. The control cam actuates a rocker arm to select the deploy position on the thrust reverser hydraulic selector valve. A roller on the control cam strikes the follow--up cam to prevent the addition of reverser thrust power until the thrust reverser is fully deployed. The follow--up cam is rotated by a push--puIl connected to one of the reverser hydraulic actuators. At full deployment the follow--up cam clears the roller, allowing reverser thrust power to be applied.


Stow Returning the reverse thrust lever to the stow position rotates the control cam, and actuates the rocker arm to select the stow position on the hydraulic selector valve. A roller on the control cam strikes the follow--up cam to prevent the application of forward tnrust until the reverser is stowed. When stowed, the push--pu I I cable rotates the follow--up cam to clear the roller.

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Figure 5 SCL

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Thrust Reverser Interlock Linkage Page: 11

BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

RIGGING Static Setting The thrust cables are rigged with the thrust system at idle using rig pins located on the drum and brake assembly, on the drum and shaft assembly, on the right side of tho cross--shaft, and on the fuel control unit. The start cables are rigged with the start system at cutoff, using the start system stop pin on the drum and shaft assembly, and a rig pin on the right side of the cross shaft. A spacer is used for start lever adjustments on the fuel control unit.The thrust reverser follow--up linkage is rigged with rig pins on the drum and shalt assembly and on the left side of the thrust reverser. If any of the rig pins can not insert on the above positions, must be perform adjust of the control cables in acordance with the M.M.




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BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

Autothrottle Clutch 3 Upper Quadrant

2

3

4 FCU

1

Rig Pin

3

6 LH RH 5 Cross Shaft

Thrust and Start Lever 1

6 5 Thrust 4 FCU

Push Pull Cables

Start Cross Shaft

6

4

FCU

Engine Left Side 5

Autothrottle Clutch

Engine Right Side


2

Figure 6 SCL

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Control Cable Rigging Page: 13



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

ENGINE DATA PLATE General During the Production Test Cell Operation, prior to delivery, the engine is checked for proper operation and certain adjustments are made to insure the performance matches a standard for that engine model. Once the engine has been fully adjusted and performance checked, a reference speed is established for use in service. The reference speed is called ”DATA PLATE SPEED”. Data Plate Speed is established from a curve which relates Engine Pressure Ratio (EPR) and High Pressure Compressor RPM (N2). This curve is slightly different for each engine. Prior to extracting the Data Plate Speed, the curve is corrected to Sea Level Standard conditions. The corrected curve is then entered at a given EPR of 1.65 (for all engines). The speed corresponding to an EPR of 1.65 is then stamped on the Data Plate in % N2 and actual RPM, together with the engine serial number. Data Plate Speed can now be used to determine engine condition in service. Location The engine Date Plate is located on the front face of the forward turbine flange at the 4 o’clock position.


Condition Check Parameters The Data Plate Speed must be corrected for ambient temperatures different from standard (6O °F), the corrections are obtained from trim tables in the Maintenance Manual. A tolerance band is specified for the corrected Data Plate Speed, this band being +1.2 % N2 and --0.8% N2. If during engine Data Plate Speed check (engine operating at EPR of 1.65), N2 RPM is within tolerance, the engine shows no deterioration and can be trimmed. If N2 RPM is above the tolerance, engine has compressor distress, if N2 RPM is below the tolerance, engine has turbine distress.

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Figure 7 SCL

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Engine Data Plate Page: 15



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

ENGINE TRIM General Trimming of the power plant is required to assure maximum operating efficiency and to maintain the power plant within certain operating limitations to prolong engine life. Engine trimming will be required as follows: A. After installation of new engine. B. After a fuel control change. C. If it becomes necessary to restore thrust lever travel between takeoff and full throttle position at the control stand. Engine trimming is performed at part power setting with the engine operating under ” no bleed and no load” conditions. Procedure Prior to performing an engine trim, certain parameters or target values must be obtained from the trim tables, i n the Maintenance Manual. , These values are a function of ambient temperature and pressure. As an example consider the following. Outside Air Temperature (OAT) 20 °F Barometric Pressure 29.5 in Hg. Data Plate % RPM 88.80 % N2. From the trim tables: A-- At OAT of 20 °F, Data Plate Correction is --3.42% N2. Hence Adjusted Data Plate Speed is 88.80 --3.42 = 85.38 % N2 Allowable tolerance is 85.38 + 1.20 = 86.58 % N2 and 85.38 --0.80 = 84.58 % N2. During Data Plate Speed Check, N2 tachometer read 85..5 % N2 hence engine condition is still within tolerance. B-- At OAT or 20 °F, idle RPM is 55.4 % N2. C-- At OAT or 20 °F and barometric presure of 29.5 in Hg, Part Power Pt7 is 60.1 in Hg. Part Power EPR is 2.04 Takeoff EPR is 2.06

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D-- Position of part power trim spacer is with C side showing. The instrumentation required for engine trimming is: A. N2 Master Tachometter connected to a test receptacle on the right side of the control cabin. B. Pt7 Pressure Gage connected at the front right side of engine. After activation of the trim spacer on the fuel control unit, the two trim screws, IDLE and MIL are adjusted until the target values are obtained. The final adjustment is always made in the increasing direction. After stowing the trim spacer, takeoff conditions are checked. Considering the staggered line, for the targets at right side, use “C” and for the targets at the left side, use “S” . Use the next figure to complete the following excercise: Excercise: For OAT 14 °F and 29.8 in.Hg. Data Plate % RPM . . . . . . . . . . . . . . . . . . . . . . __________% N2 Data Plate Correction . . . . . . . . . . . . . . . . . . . . __________% N2 Adjusted Data Plate . . . . . . . . . . . . . . . . . . . . __________% N2 Allowable N2 % tolerance +1.2 --0.8 At data check N2 tachometer read . . . . . . . . . __________% N2 The engine still within tolerance? . . . . . . . . . . . __________ Idle RPM . . . . . . . . . . . . . . . . . . . . . . . . . . . . . __________% N2 Part Power Pt7 . . . . . . . . . . . . . . . . . . . . . . . . __________in.Hg. Part Power EPR . . . . . . . . . . . . . . . . . . . . . . . . __________ Take Off EPR . . . . . . . . . . . . . . . . . . . . . . . . . . __________ Position of part power trim spacer is showing side. . __________

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Figure 8 SCL

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Trim Table Excercise Page: 17



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

PART POWER TRIM SPACER General Engine trimming is performed at part throttle or part power setting. On the fuel control unit, is installed a spacer having two positions, S for standard and C for cold. The S or C side is used in accordance with data from the trim tables. Location The trim spacer is located on the fuel control unit under the thrust crank arm stop.


Procedure Depending on the ambient conditions (tenperature and pressure), from the trim tables the position of the trim spacer is obtained. If S position is required, break satety wire, loosen screw and rotate the spacer to between the stop and thrust crank. If C position is required, break safety wire, remove the spacer and install it with C side showing. Upon completion of trim procedure, the spacer should be stowed. The spacer should not be removed from the fuel control unit and used on a different fuel control unit.

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Figure 9 SCL

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Part Power Trim Stop Page: 19



BOEING 737 - 200 JT8D -- 7 to -- 17 76 -- 00

FUEL CONTROL General Engine trinning is performed by adjustment of two screws on the fuel control unit in accordance with target values obtained from the trim tables. Location The two adjustment points are located on the fuel control unit next to the thrust crank on the forward right side of the engine.


Operation The two trim screws are called IDLE and MIL. The IDLE screw is rotated counterclockwise for increasing the idle RPM and the MIL screw is rotated counterclockwise for increasing part power Pt7. One complete revolution of each trim screw generates 32 ”clicks”. On IDLE screw, 16 ”clicks” is equivalent to idle RPM change of 1% N2 and on the MlL screw, 10 ”clicks” is equivalent to Pt7 change of 1 in Hg. During trim procedure, both screws are adjusted until the target values are obtained.

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Figure 10 SCL

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Fuel Control Page: 21

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BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

INDICATION SYSTEM



POWER PLANT INDICATION SYSTEM

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BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

INDICATIONS SYSTEM Purpose The indicating systems monitor (sense) some of the important engine conditions and show them on the flight deck. Types of Engine Data There are indicating systems for these types of engine performance (or power) data: -- Engine pressure ratio (EPR) -- Low rotor (N1) speed ---- percent of RPM -- High rotor (N2) speed ---- percent of RPM -- Exhaust gas temperature (EGT) ----°C -- Fuel flow rate ---- thousands of lb/hr or kg/hr. Also, fuel used (optional). There are indicating systems for these types of oil system data: -- Oil temperature ---- °C -- Oil pressure ---- psi. Also, low oil pressure caution -- Oil quantity ---- quarts, gallons, or liters. There is an indicating system for the amplitude of engine vibration, but is optional, because in the most of the 737 it‘s not used.

Oil Indications The oil indications are comments in the Oil System ATA Chapter 79. -- Oil pressure -- Oil temperature -- Oil quantity

Gages The flight deck indicators are usually circular analog gages (except for the low oil pressure caution light). These gages have pointers, and some of them also have digital indicators. There is one gage for each engine in the aircraft. On 737 aircraft, the gages are located on the center instrument : -- EPR -- N1 % RPM -- EGT -- N2 RPM -- Fuel Flow.

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Figure 1 SCL

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Indications System Page: 3



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

ENGINE PRESSURE RATIO SYSTEM General The engine pressure ratio (EPR) system provides a visual indication in the control cabin of power or thrust for any throttle setting, except idle. Same as, must be consider the EPR indication is the primary or master indication, therefore the engine must not accelerate if this indication is wrong or inoperative. Components The system consists of one engine inlet pressure, Pt2, sensor located in the nose dome, six engine exhaust pressure, Pt7 sensors, manifolded together at engine exhaust, pressure ratio transmitter located in forward section of air conditioning bay and indicator on P2 center instrument . The transmitter receives 115 Vac power from P6--2 circuit breaker .


Inlet Pressure Sensing Probe The engine inlet pressure (Pt2) is sensed by a pitot probe mounted through the center of the nose dome. The pressure signal es through the moisture trap and exits the engine through the 6 o’clock inlet guide vane. The vent hole in the moisture trap functions as an ice detector. With formation of ice, Pt2 pressure in line decreases by venting to nose dome, thus causing EPR to increase. The probe is anti--iced by the engine anti--icing air present in the nose dome. Pt2 test connection is provided on front right side of engine and is used for leakage check.

Pressure Ratio Transmitter The transmitter consists of a bellows actuated capacitance type sense unit, a cam and gear train, an amplifier, two phase motor and gear train and a transmitting synchro enclosed in a case. The transmitter is mounted on an anti--vibration rack. The Pt2 and Pt7 lines entering the transmitter are provided with test connections for leakage checks. Pressure Ratio Indicator The indicator consists of a receiver synchro, dial. and pointer, an adjustable index and counter and an external set knob. The set knob is used to position index and counter to a desired engine pressure ratio for reference only. Different EPR indicator are installed on 737 aircrafts, some of these indicators are provided with one window to digital readout of actual value, and the other indicators with two windows, one for actual value and the other window to command or refference EPR value. Some airplanes incorporating EPR actived takeoff warning aural. The indicator has integral lighting. Maintenance Both Indicator and EPR Tx are LRU‘s.

Exhaust Pressure Sensing Probes The engine exhaust preesure (Pt7) is sensed by six probes projecting into the exhaust stream. The probes are connected to a common manifold, thus providing an average exhaust pressure. Exterior connection to the manifold is made at a single point through the fan discharge outer duct at 7 o’clock position. Pt7 test connection is provided on front right side of engine and is used for leakage check or Pt7 pressure gage for engine condition check and engine trimming.

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Figure 2 SCL

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EPR System Page: 5



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

EPR OPERATION Operation When the system is not at null, as would occur when either Pt2 or Pt7 was changed, the pressure transmitter is also at off--null position. In this condition, the center plate of the capacitance pickoff is not centered between the outer plates and therefore a voltage is produced which is amplified by a phase sensitive amplifier. The voltage is then supplied to a two--phase motor, whose direction of operation depends on whether the voltage is in phase or out of phase with respect to line voltage. The motor operates the output synchro and the outer plates in a direction to null the center plate of the capacitance pickoff. The signal from the output synchro is transmitted to the indicator.

DIGITAL READOUT EPR ONLY

General The engine pressure ratio transmitter converts the pneumatic signals from engine inlet and exhaust to an electrical signal which is supplied to the EPR indicator. Location The transmitter for each engine is located in the respective air conditioning bay, outboard wall, opposite the waer separator. Located on the transmitter are test connections for the engine inlet (Pt2) and exhaust (Pt7) pressures.

DIGITAL READOUT COMMAND AND


ACTUAL EPR

EPR INDICATOR TYPES

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Figure 3 SCL

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EPR Operation Page: 7



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

ENGINE TACHOMETER SYSTEM General The engine tachometer system measures the rotary speed of the engine low (N1) and high (N2) pressure compressors, and provides a visual indication in the control cabin for monitoring engine performance. Components The tachometer system consists of a Low pressure compressor tachometer generator located on the trout accessory drive pad behind the nose dome, a high pressure compressor tachometer generator located on right side of the accessory drive case and two indicators on the P2 center instrument . Tachometer Generator The tachometer generator consists of a three phase stator winding, end shields and a permanent magnet rotor assembly all contained in a sealed case. The frequency of the generato output signals is a function of the compressor RPM. Tachometer indicator The tachometer indicator consists of a three phase synchro motor, a rotating drag assembly, a calibrated clock spring, all hermetically seated in a case, and a round dial with a subdial. The indicator dial is graduated for readings between zero and 110 percent RPM, while the small subdial is graduated in ten divisions for each 10 percent change in speed.


Maintenance Both indicator and tachometer are LRU‘s. The generator tachometer are interchangeable.

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Figure 4 SCL

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Engine Tachometer System Page: 9



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

GENERATOR TACHOMETER Operation The alternating electrical signals from the respective tachometer generator are fed into a tachometer indicator which cause the drive shaft of the indicator synchrro motor to rotate. The synchro motor drives the flux coupling which, in turn, restrained by a clock--type spring, moves the indicator pointers to a dial position corresponding to the compressor rotor speed. The engine tachometer system operates on self--generated electrical power. The airplane’s electrical power is required only for the integral lighting of the tachometer indicators.


Engine Trim Connection A N2 test connection is provided on the right side of the control cabin.This connection can be used for checking N2 RPM on a precision tachometer during the engine data plate speed check and engine trim.

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Figure 5 SCL

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Generator Tachometer Operation Page: 11



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

EXHAUST GAS TEMPERATURE CIRCUIT General The operating limits of the engine, monitoring of the mechanical integrity of the turbines and engine condition during operation are performed by measurement of the engine exhaust gas temperature which is displayed in the control cabin. Components The exhaust gas temperature (EGT) system consists of eight thermocouples arranged radially in the engine exhaust, thermocouple junction box located on left side of turbine case and an indicator on the P2 center instrument . The indicator receives 115 Vac power from P6--2 circuit breaker . Chromel and alumel wires are used between the thermocouples and indicator.

Operation The difference in temperature, between the hot junction at the thermocouple probes and the cold junction at the indicator, causes electrical signals to be generated in the thermocouple circuit. This electrical signals, with voltage proportional to temperature, is applied to the amplifier. The amplifier receives regulated power supply from transformer within the indicator. Output of the amplifier is applied to a torquer which in turn drives the indicator pointer. With power interrupted, pointer goes to the 0 position. The indicator incorporates a cabin temperature compensator to provide exhaust gas temperature indication irrespective of prevailing ambient temperature.

Thermocouple Probe The thermocouple consists of a junction of a chromel and alumel wires enclosed in a tube which has a series of inlet holes facing the gas stream. The probe has two stud terminals, the alumel (--) terminal is longer than the chromel (+) terminal. The correct orientation of the probe with respect to gas stream is obtained with an index slot in the probe. The eight thermocouples are connected electrically in parallel to obtain the average gas temperature.


Thermocouple Junction Box The two wired (chromel and alumel) from each thermocouple terminate at the junction box. The chromel wires are connected together to form a parallel circuit, the alumel wire is cornnon to all thermocouples. The jjunction box is used to check thermocouple continuity. From the junction box, the chromel and alumel wires are routed along the left side of engine diff case to terminal strips. Exhaust Gas Temperature Indicator The indicator is an independently powered servoed instrument with a dial graduated from 0 to 850 °C. The normal and dangerous operating tempera ures are marked in a color code The range between 500 °C and 700 °C is expanded to give a more accurate indication. The indicator has integral lighting.

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Figure 6 SCL

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EGT Circuit Page: 13



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

EGT THERMOCOUPLE General The exhaust gas thermocouples are located within the primary exhaust flow and transmit electrical signals to a flight deck indicator. Installation The eight thermocouples are mounted in the inner wall of the primary exhaust duct. An index guide assures the proper positioning of the probe and a lock nut holds the probe in position. A two segment duct cover ring covers the terminal end of the probe and creates a age for the interconnecting wiring.


Accessibility Access to the thermocouples on a built up engine is through the tail pipe and thrust reverser. Deactivation of the thrust reverser is required prior to entering the engine.

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Figure 7 SCL

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EGT Termocouple Page: 15



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

FUEL FLOWMETER SYSTEM General The fuel, flow indicating system provides a visual indication in the control cabin of the fuel consumption of each engine. Components The system consists of a single Power Supply Module (M316), located on the E3--3 Electronic Equipment rack, flow transmitter located on forward left side of engine and indicator calibrated in lb/hr or kg/hr on P2 center instrument . Power supply is 115 Vac and 28 Vdc from P6 circuit breaker .


Operation The power supply converts 28 Vdc to 3 ∅, 4Hz, l7 Vac which is supplied to the flowmeter transmitters. The conversion is accomp]ished by timer switches and oven controlled oscillator.

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Figure 8 SCL

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Fuel Flow System Page: 17



BOEING 737 - 200 JT8D -- 7 to --17 77 -- 00

FUEL FLOW CIRCUIT Transmitter The transmitter consists of a motor driven impeller and a fuel flow driven turbine. An impeller motor is located upstream of the impeller. A signal transmitting unit is located downstream of the turbine and is attached to the turbine with two restraining springs of different load characteristics. This permits the use of an indicator with fine index marks at the low side and coarser index marks at the high side resulting in greater sensitivity at low flow rates. The impeller motor receives 3 phase, 4 Hertz, 17 Vac from Power Supply Module and is driven at constant 240 RPM, the impeller is driven through reduction gears at constant 60 RPM regardless of fuel flow. The fuel ing through the impeller is given a swirling motion and then es through the turbine. The action of the fuel on the turbine blades produces a torque, against the restraining springs, which is directly proportional to the fuel flow rate. Rotation ol the restraining springs rotates a permanent magnet rotor surrounded by a coil which is supplied with 115 Vac from P6--3 circuit breaker . It is reconnended that the FUEL FLOW INDICATOR circuit breaker be pulled (open) whenever the engine fuel feed line is drained or the airplane is out of service for maintenance. This will prevent damage to the transmitter from dry operation and will prolong its life. Indicator The fuel flow indicator consists of a permanent magnet rotor surrounded by a coil which is supplied with 115 Vac from P6--3 circuit breaker , a pointer and a dial face calibrated in lb/hr or kg/hr. The coil picks up the output signal from the transmitter, rotates the permanent magnet in indicator by the same angular displacement as the magnet in transmitter. The pointer gives a visual indication of fuel flow during engine operation. The indicator has integral illumination. Maintenance The indicator and transmiter are LRU‘s.

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Figure 9 SCL

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Fuel Flow Circuit Page: 19


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Figure 10 SCL

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Summary Engine Indications Page: 20

BOEING 737 - 200 JT8D -- 7 to -- 17 74 -- 00

IGNITION SYSTEM



POWER PLANT IGNITION SYSTEM

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BOEING 737 - 200 JT8D -- 7 to -- 17 74 -- 00

IGNITION SYSTEM Purpose The igrntion system supplies a fast sequence of electrical sparks to ignite the fuel/air mixture in combustion chambers 4 and 7. On the left side of the engine there is: -- The exciter (box) is mounted betweefl. flanges J and J1 at the 1:00 position. -- A high tension lead goes from the exciter to the igniter plug which is threaded into the combustion chamber 7 at the 8:00 position. On the right side of the engine there is: -- A high tension lead goes from the exciter to the igniter plug which is threaded into the combustion chamber case at the 4:00 position. -- The high tension lead goes into combustion chamber 4. Description and Operation The ignition system is used when the engine is started (on the ground) or restarted (in flight). It is also used (as a precaution) to prevent flameout during takeoff, landing, and bad weather. The ignition system has an exciter that, gets electrical power from the aircraft, transforms it into a high voltage and sends that voltage through high tension leads to (one or both of) the igniter plugs. The igniters make very hot sparks which ignite the air/fuel mixture in the combustion chamber. There are two kinds of exciters: A-- The standard 20--4 joule system has a single exciter box. The aircraft supplies 115 Vac and 28 Vdc electricity to it. The 20--4 joule exciter can operate in the intermittent or continuous mode. In the intermittent mode, it fires both igniters in the continuous mode, it only fires the igniter in combustion chamber 7. B-- The optional dual 20 joule system has two exciter boxes bolted together. ( They are referred to as a ”twin pack” ). Both exciters get 115 Vac electricity from the aircraft. One exciter fires the igniter in combustion chamber 4, and the other exciter fires the igniter te combustion chamber 7. The two exciters can be operated individually or at the same time. Earlier models were for intermittent use only. Current models are available for continuous duty.

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The ignition system is controlled from the flight deck. When it is turned on: -- Electrical current flows from the aircraft to the exciter. -- The exciter produces a fast sequence of high energy electrical pulses and sends them through the high tension leads to the igniters. -- Each electrical pulse causes the igniter (plug) to make a strong spark at its tip, which is in a combustion chamber. The following steps occur to produce each spark: -- The exciter increases the voltage of the input electricity and builds a charge on a storage capacitor. -- After the charge has built up sufficiently on the capacitor, a part of that charge is output (sent) to the igniter. -- The igniter makes a ”trigger spark” which ionizes the air at its tip. -- Then the capacitor discharges fully, and the remaining charge is sent to the igniter. -- The igniter makes a high--energy spark which is conducted by the ionized air to make extended with the fuel/air mixture. The heat that is transmitted from the very hot spark to the aerated fuel particles causes combustion. NOTE:

EACH ELECTRICAL OUTPUT PULSE FROM THE EXCITER HAS TWO PARTS WHICH OCURR IN A VERY FAST SEQUENCE. THE FIRST PART CAUSES THE TRIGGER SPARK, AND THE SECOND PART CAUSES THE HOT SPARK.

WARNING:

TO PREVENT A DANGEROUS ( FATAL) ELECTRICAL SHOCK, DO NOT WORK ON THE IGNITION SYSTEM: -- WHILE IT IS OPERATING. -- UNTIL THE EXCITER HAS BEEN DISCHARGED BY GROUNDING THE DISCONNECTED HIGH TENSION LEAD(S) TO THE EXCITER MOUNTING BRACKET, NOT TO THE ENGINE CASE.

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Figure 1 SCL

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Ignition System Components Page: 3



BOEING 737 - 200 JT8D -- 7 to -- 17 74 -- 00

ELECTRICAL FLOW Description and Operation A high tension voltage developed in the ignition exciter is sent through the ignition leads across the gap of the igniter plug. This voltage causes the air between the igniter plug electrodes to ionize. When ionization occurs, the current discharges between the electrodes result in a high energy spark which ignites the fuel/air mixture in the combustion chamber. Pratt & Whitney recommends that both igniters be used for ground starts and air starts to ensure rapid, dependable starting.


* 20--4 Joule System: -- One power cable supplies the single 20--4 joule exciter with two input voltages from the aircraft: 28 Vdc to power the intermittent duty circuit, and 115 Vac to power the continuous duty circuit. The intermittent circuit fires both igniter plugs with 20 joules. The continuous circuit fires one igniter plug with 4 joules. A spark gap assembly ensures that voltage is supplied to only one igniter during continuous circuit operation. -- The engine start switch is a guarded, momentary switch used to control the starter shutoff valve and the electrical supply to the ignition exciter. -- All ground starts and inflight air starts should be made using the 20--joule DC exciter (firing both igniter plugs). -- For optimum life of the ignition system components, the operating duty cycle is 2 minutes ON, 3 minutes OFF, 2 minutes ON and 23 minutes OFF. In addition, the 4--joule system should be off during normal flight -- The 4--joule system should be used for protection against flameout during takeoff and prior to activating the engine inlet anti--icing system. -- “GRD START AND CONT” provides 20 joules to both igniters when start switch and fuel shutoff lever are ON. When start switch is released, system goes to 4 joules to one igniter. -- “IGN OVRD” provides 20 joules to both igniters bying start switch and fuel shutoff levers. -- SB 5880 permits an operator to convert to an improved 20 joule exciter that permits continuous duty operation.

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* 20 Joule Ignition System (Optional) -- The dual 20--joule exciters receive 115 Vac power from the aircraft. Each exciter has an input and an output connector. Both exciters are bolted together but function independent of one another. -- The 20--joule ignition system is controlled manually by a switch on the overhead in the flight deck. The exciters can be switched on individually or simultaneously. -- The start switch is a guarded, momentary switch used to control the starter air shutoff valve and the AC supply to the ignition exciters * Switch ”A” or ”L” provides 115 V, 400 Hertz electrical power to exciter A when the fuel shutoff lever is ON. -- Switch ”B” or ”R” provides 115 V, 400 Hertz electrical power to exciter B when the fuel shutoff lever is ON. -- BOTH provides 115 V, 400 Hertz electrical power to exciters A and B when the fuel shutoff lever is ON. -- OVRD provides 115 V, 400 Hertz electrical power to exciters A and B, bying the fuel lever and start switch. -- For optimum life of the ignition system components, the operating duty cycle is 10 minutes ON, 20 minutes OFF. -- An improved 20 joule exciter that permits continuous duty operation is available. Refer to SB 5803. The seven combustion chambers without igniter plugs have their fuel/air mixtures ignited by the flame that goes (propagates) through the cross--over tubes between adjacent cans. Pratt & Whitney recommends that both igniters be used for ground starts and in--flight restarts to ensure rapid, dependable starting and The ignition system be turned OFF during normal flight. Maintenance SB 5961 provides an alternate supplier of the exciter box. All of the ignition system components are LRU‘s. The gap assembly is LRU. The igniters can be cleaned.

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Figure 2 SCL

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Ignition Electrical Flow Page: 5

BOEING 737 - 200 JT8D -- 7 to -- 17 74 -- 00

NOTES:




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BOEING 737 - 200 JT8D -- 7 to -- 17 80 -- 00

STARTING SYSTEM



POWER PLANT STARTING SYSTEM

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BOEING 737 - 200 JT8D -- 7 to -- 17 80 -- 00

STARTING SYSTEM General The pneumatic starting system provides means for rotating the engine to the RPM range, where starting can be accomplished when ignition and fuel are supplied. Components The pneumatic start system consists of a pneumatic starter located on the left aft face of the accessories drive case and solenoid controlled, pneumatically operated start valve located above the starter. The start valve control switch is located on the P5, forward overhead . The two engine start systems utilize the pneumatic manifold and valves for air supply to starters.

Using air supply from ground air supply cart. Air from ground cart through engiuo No. 2 bleed valve is supplied to No. 2 start valve or through isolation valve and engine No. 1 bleed valve to No. 1 start valve. Using bleed air from opposite operating engine. Adequate pressure is provided from the 8th stage bleed of the engine operating at approximately 80% N2 RPM. Air from the operating engine through its bleed valve, isolation valve and bleed valve of engine to be started, is supplied to the start valve.


Air Supply The air for the pneumatic starter can be supplied by: A. Auxiliary Power Unit B. Ground air supply cart C. Bleed air from opposite engine (if operating). Before an engine start is attempted, the air conditioning system pack valves and wing thermal anti--icing valves must be closed, isolation valve and engine bleed valves must be open. Switches for these valves are located on the P5, forward overhead . Air pressure for engine start is Indicated on a dual pointer pressure gage on P5 . Operation Using air supply from Auxiliary Power Unit. With APU operating, the APU bleed valve is opened by a switch on the P5, forward overhead . Air from APU through the isolation valve and engine No. 2 bleed valve is supplied to the No. 2 start valve, or through engine No. 1 bleed valve to the No. 1 start valve. During the time when the APU bleed valve, isolation valve, engine No. 1 bleed valve and engine No. 2 bleed valve are open, the DUAL BLEED amber light on P5 will be illuminated. Tho ”dual bleed” condition is allowed to exist only during engine starting. Upon completion of engine starting, the APU bleed valve must be closed and the light will be extinguished.

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Figure 1 SCL

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Starting System Schematic Page: 3



BOEING 737 - 200 JT8D -- 7 to -- 17 80 -- 00

START CONTROL CIRCUIT General The engine start circuit is used for engine starting on the ground, continuous ignition (if necessary) during engine operation and engine re--start in flight. Components Start system components consist or a solenoid controlled, pneumatically operated start valve located on loft side of engine, pneumatically driven starter located on left aft side of accessories drive case, ignition exciter unit on left side of diff case and two igniters located in No. 4 and No. 7 combustion chambers. A four position, (GRD, OFF, LOW IGN, FLT), rotary type engine start switches with holding coils are located on P5 forward overhead . Operated by the engine start lever are two ignition switches located in the control pedestal. The engine start circuit receives 28 Vdc and 115 Vac electrical power from P6--2 circuit breaker . Operation Pushing and rotating the engine start switch to GRD position closes two s. 28 Vdc through the first energizes start swttch holding coil, ground for coil is through cutout switch in engine starter. The start switch is now held in GRD position. Also start valve solenoid is energized and start valve opens, supplying air to starter, evidenced by pressure drop on duct pressure gage. The si arter drives tim high presuure compressor and turbine through the accessories drive. At approximately 16 -- 17 % RPM of high pressure compressor and providing that low pressure compressor is rotating, the start lever is advanced from CUT--OFF to IDLE position. 28 Vdc through second of start switch and ignition switch is supplied to the ignition exciter unit. Output of exciter unit is 20 Joules of electrical energy to each igniter at approximately 5 sparks per 10 seconds. At cutout speed, the centrifugally operated switch inside the starter opens, start switch holding coil releases and start switch returns to OFF. 28 Vdc is interrupted to ignition system and start va)ve solenoid, start valve closes, evidenced by pressure recovery on duct pressure gage. The engine continuos to accelerate to idle speed.

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On No. 1 engine only, the starter cutout switch provides power to the stall warning heating system. If the start valve is opened” manually, it must be closed when high pressure compressor speed reaches 35--4O % RPM. If the engine start switch does not hold in GRD position, it indicates that either the holding solenoid is not energized or the starter cutout switch has failed open. The engine can still be started by holding manually the start switch in GRD position. It the engine start switch does not return to OFF at starter cutout speed, it indicates that cutout switch failed closed. Place start switch to OFF manually at cutout speed. Continuous Ignition Continuous or low energy ignition is used during takeoff, turbulance and landing. Pushing and rotating the engine start switch to LOW IGN position closes fourth . 115 Vac through fourth of start switch and ignition switch is supplied to the ignition exciter unit. Output of exciter unit is 4 Joules of electrical energy to igniter in No. 7 combustion chamber only at approximately one spark every two seconds. If the low ignition system has failed, the low ignition spark gap on the exciter unit has to be changed. Flight Start In order to re--start the engine in flight, the starter is not operated, since the engine with be windmilling. Pushing and turning the engine start switch to FLT position closes third . 28 Vdc through third of start switch and ignition switch is supplied to the ignition exciter unit. Output of exciter unit is the same as during the ground start. When engine reaches idle speed, place switch to LOW IGN. THE ENGINE START SWITCH IS HELD MECHANICALLY IN LOW IGN AND FLT POSITIONS. The engine is stopped by placing the start lever at cutoff. NOTE:

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Figure 2 SCL

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Start Control Circuit Page: 5



BOEING 737 - 200 JT8D -- 7 to -- 17 80 -- 00

ENGINE STARTER General The engine is started using low pressure air(supplied to the starter through the start valve. The pneumatic starter is a turbine type air motor which converts energy of compressed air into torque sufficient to accelerate the engine to starting speed. The starter drives the high pressure compressor through the accessories drive case. Location The starter is located on the left aft face of the accessories drive case. It is attached by five keyhole type bolts, provisions are for six, but the bolt at 1 o’clock position is not used due to limited accessibility. Components The starter consists of a scroll assembly, turbine wheel, reduction gear assembly, engaging mechanism, output shaft and a centrifugal cutout switch. The air exits the starter through a safety screen and a containment deflector.


Operation When the valve is open, air is itted to the starter scroll assembly and is directed radially inwards through the turbine wheel imparting high speed rotation. The reduction gear train reduces the high speed, low torque of the turbine wheel into low speed, high torque through a pawl and ratchet engagement mechanism to the output shaft. When the speed of the output shaft exceeds the speed of the internal gear hub, the clutch mechanism overruns thus providing automatic disengagement. Overspeed control is maintained by the centrifugally operated cutout switch, which automatically closes the starter valve when the output shaft reaches a speed corresponding to 35--40 % N2 RPM.

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Figure 3 SCL

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Engine Starter Page: 7



BOEING 737 - 200 JT8D -- 7 to -- 17 80 -- 00

PNEUMATIC START VALVE General The start valve allows airflow from the airplane’s pneumatic system to the starter during engine starting operation. The valve is solenoid controlled, pneumatically operated non--.regulating. Location The pneumatic start valve is located on the heft side of the engine just above the engine starter. Components The start valve consists of a valve body assembly with a lever operated butterfly valve, diaphragm type pneumatic actuator assembly, filter and solenoid operated switcher with a manual override.


Operation Placing the engine start switch, on the P5 overhead , to GRD posjtion, allows start valve solenoid to be energized and moving the switcher. Air from the pressurized pneumatic manifold es through the filter and switcher to the lower chamber of actuator assembly, the upper chamber is vented to ambient. Pressure acting on the bottom of the diaphragm overcomes the spring force, and by linkage moves the butterfly to open position. Deenergizing the solenoid ( normally by the cutout switch in the starter) , repositions the switcher. The lower chamber of actuator assembly is now routed to ambient through the switcher. the spring in upper chamber and torsion spring on butterfly close the valve.

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Figure 4 SCL

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Start Valve Schematic Page: 9

BOEING 737 - 200 JT8D -- 7 to -- 17 80 -- 00

START VALVE Manual Override If the solenoid fails to operate normally, the manual solenoid override button can be used. It is accessible through a small hole in the left cowling, very near to the CSD oil service access door, with a suitable tool (screwdriver) using the cowling as a pivot point. The override must be released when engine speed reaches 35--40% N2 RPM. A second override method requires the left copling to be open. The valve butterfly position indicator can be rotated with an Allen key to open the valve. The valve must be closed whet engine speed reaches 35--40% N2 RPM.




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Figure 5 SCL

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Manual Override Page: 11

BOEING 737 - 200 JT8D -- 7 to -- 17 80 -- 00

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

EXHAUST



POWER PLANT EXHAUST

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

THRUST REVERSER General The thrust reverser assembly is located on the aft end of each engine and is for use on the ground only. When actuated, the reversers provide reverse thrust as a means of decelerating the aircraft during landing roll or after a rejected takeoff. Reverse thrust is provided by two deflector doors opening aft, which deflect the exhaust gases forward. The reversers are hydraulically powered, electrically controlled and monitored.



POWER PLANT EXHAUST

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00


POWER PLANT EXHAUST

Figure 1 SCL

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T/R View Page: 3


POWER PLANT EXHAUST

BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

THRUST REVERSER INSTALLATION General The thrust reverser assembly is attached to the tailpipe extension, which in turn is attached to the engine. The reverser assemblies are interchangeable between the engines.


Components A double flanged tailpipe extension is installed between the aft end of the engine and the forward end of the reverser, these flanges are fireshields. The thrust reverser assembly consists of an integral tailpipe and shroud assembly, two doors with locks, two lock and door actuators, actuating linkage and interlock push--pull cable with rigging provisions. At the forward end of the reverser assembly are located the hydraulic components (check valve restrictors) and lines for pressure and return flows, for the two lock and door actuators.

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POWER PLANT EXHAUST

Figure 2 SCL

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T/R Component Locations Page: 5


POWER PLANT EXHAUST

BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

THRUST REVERSER DOOR LOCK General The thrust reverser door lock assembly maintains the thrust reverser in the forward thrust position except when reverse thrust operation is selected. Location The thrust reverser door lock assemblies, two for each reverser, are located circumferentially around the junction of the jet pipe and reverser under the nacelle skin. Components The door lock assembly consists of a latch, a hydraulically operated lock actuator, a striker pin, a torsion spring and a tension spring. Extension of the latch is an actuating plate for the proximity switch sensor unit. Attached to the door is a latch fitting with a roller which is engaged by a latch.


Operation During the selection of reverse thrust, the lock actuator is pressurized moving the latch, against torsion spring, and thus unlocking the door. With the door moving the retracting arm is pulled by the tension spring and engages the striker pin, holding the latch In unlocked position. During the selection of forward thrust, the rock actuator is again pressurized to a reset position, without moving the latch since the actuator rides in a slot in the latch. When the door is stowed, the retracting arm is moved by the latch fitting, releasing the latch. The torsion spring now rotates the latch which engages the roller on the latch fitting.

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00


POWER PLANT EXHAUST

Figure 3 SCL

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T/R Door Lock Page: 7


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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

THRUST REVERSER DOOR LINKAGE General The thrust reverser door linkage is a four--bar linkage system used to stow and deploy the thrust reverser doors. The linkage assembly is powered by the hydraulically operated actuators. Location The thrust reverser door linkage assemblies, two for each reverser, are housed behind aerodynamic fairings at the rear of the nacelle. Components The linkage assembly consists of a driver link, idler link, and an overcenter links. The drive links are attached to the forward part of the doors and to the overcenter links, the idler links are attached to the aft part of the doors and assembly. Attached to the overcenter links is the reverser actuator. On the left actuator only is attached the interlock rod.


Operation With the reverser in the forward thrust position, the linkage is maintained in the overcenter position, thus keeping the reverser locked. During the selection of reverse thrust, the reverser actuators are pressurized moving the overcenter links, which in turn rotate the drive links. Rotation of the drive links moves the doors to the open position. During the selection of forward thrust, the rfeverser doors are closed and locked by the actuator.

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00


POWER PLANT EXHAUST

Figure 4 SCL

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T/R Door Linkage Page: 9


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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

THRUST REVERSER HYDRAULIC / ELECTRICAL SCHEMATIC General The thrust reverser operation is controlled by a reverser system which is hydraulically operated and electrically controlled. The actuating medium is pressurized hydraulic fluid supplied by the aircraft landing gear down line portion of the A hydraulic system or alternatively from the standby hydraulic system. Indication of reverser operation is provided in the control cabin.


Components The system consists of the following components: A. Hydraulic fluid supply through a shuttle valve in the main wheel well. B. Isolation valve, located at forward end of air conditioning bay, receiving power from circuits in Engine ory Unit Module M523. C. Pressure switch, located in leading edge o wing/body fairings electrically connected to circuits in M5281 D. Selector valve, located in wing leading edge above engine, operated by the reverse thrust lever and thrust lincage. E. Lock actuator, door actuation and check vflve/restrictors located on the thrust reverser. F. Engine Accessory Unit, M528 located on E3-- rack receiving inputs from Air/Grd switch, engine low oil pressure switch, hydraulic pressure switch and proximity switches located on reverser doors. The unit contains circuits for operation o isolation valve, ISOLATION VALVE light (P5) and REVERSER UNLOCKED big t (P2). G. Engine fire switch on P8 and reverse override switch on P5 . H. The power supplies are 28 Vdc from P6--2 circuit breaker .

Whenever the reverser doors are unlocked, the proximity switches energize circuits in Engine Accessory Unit M528. Through these circuits REVERSER UNLOCKED amber light on P2 is illuminated. Stowing of the reverser is accomplished by repositioning of the selector valve by the reverse thrust lever. Hydraulic fluid is supplied to reposition the lock actuators and then to the door actuators, return flow is through the selector valve. When the isolation valve is deenergized, the lock and door actuators are slowly depressurized through the restrictor and isolation valve. The pressure switch, located downstream of the isolation valve, is connected to circuits, in Engine Accessory Unjit, M528 for the ISOLATION VALVE light on P5 Aft Overhead . With isolation valve energized, illuminated light indicates low pressure. With isolation valve deenergized, illuminated light indicates high pressure. During maintenance practices, with the reverser in the open or closed position, it is necessary to depressurize thef actuators to prevent the inadvertant use of reverser. To depressurize the reverser system, the manual lockout plunger on the isolation valve is depressed and held depressed by a pin and lockout mechanism. This action connects the actuators to the return line.

Operation In order to operate the thrust reverser, the isolat ion valve must be energized. The valve is normally energized when the fire switch is in the NORMAL position, airplane is on the ground and engine is operating. An override switch is provided for energizing the isolation valve when airplane is not on the ground and/or engine is not operating. With hydraulic pressure available, movement of reverse thrust lever allows selector valve to port fluid to lock actuators first by virture of check valve/restrictors and then to the door actuators. The return flow from the actuators is through the selector valve.

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00


POWER PLANT EXHAUST

Figure 5 SCL

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T/R Control Schematic Page: 11


POWER PLANT EXHAUST

BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

THRUST REVERSER ISOLATION VALVE General The thrust reverser isolation valve (one per engine), is a solenoid operated valve which allows hydraulic supply to and return from the reverser actuators. Location The isolation valve is located on the forward bulkhead of each air conditioning bay.


Components The valve consists of a shuttle valve operated by a 28 Vdc powered solenoid for opening and a spring for closing. It has three ports for hydraulic pressure, supply to actuators and hydraulic return. On the outside of the valve is a manual lockout plunger with provisions for a locking pin. Operation When the solenoid is deenergized, the valve is closed by the spring and the supply line to actuators is connected to the return line. This depressurizes the actuators on the reverser. In order to energize the solenoid, the airplane must be on the ground and engine operating or the override switch placed to the OVERRIDE position. When the solenoid is energized, the shuttle valve is repositioned to connect the supply line to the actuators to open or close the reverser. During the performance of the maintenance practices, the reverser may be in open or closed position with the actuators pressurized. This condition is very undersirable, since inadvertant movement of reverse thrust lever will move the reverser. Hence prior to working in the vicinity of the reverser, the isolation valve manual lockout plunger should be depressed and the ground lock assembly with the locking pin should be installed. The valve is now locked in closed position depressurizing the reverser actuators and movement of thrust reverser lever will not operate the reverser.

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00


POWER PLANT EXHAUST

Figure 6 SCL

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Jan -- 2004

T/R Isolation Valve Page: 13


POWER PLANT EXHAUST

BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

THRUST REVERSER ISOLATION VALVE AND LIGHT CIRCUIT General In order to operate the thrust reverser, it is necessary to energize the thrust reverser isolation valve. Normally, the isolation valve is energized when the fire switch is in the NORMAL position, airplane is on the ground and engine is operating. An override switch is provided for energizing the isolation valve when airplane is not on the ground and/or engine is not operating. A light is provided for the isolation valve.


Components The circuit components consist of the following: A. Fire switch located on P8 in control cabin. B. Air/Grd relay in Landing Gear Accessory Unit M338 on E3--2 rack. C. Isolation valve and light control circuits in Engine Accessory Unit M528 located on E3--2 rack. D. Thrust reverser isolation valve, located at forward end of air conditioning bay. E. Thrust reverser OVERRIDE switch (NORMAL and OVERRIDE) guarded in NORMAL position and ISOLATION VALVE amber light located on P5 aft overhead . F. The power supplies are 28 Vdc from P6--2 circuit breaker . A. Normal With airplane on the ground, relay K5 in M338 is deenergized and its s are in GRD position. With engine operating, engine oil pressure is normal. Relay K2 in M528 is energized, power for relay is 28 Vdc from Battery Bus through NORMAL position of fire switch and ground through the oil pressure switch. 28 Vdc from Battery Bus through NORMAL position of fire switch, of relay K5 in GRD position in M338, of relay K2 in M528 to energize the isolation valve. B. Override Placing the thrust reverser override switch to OVERRIDE position, 28 Vdc from Batt Bus through NORMAL position of tire switch is supplied to energize the isolation valve, bying the airplane on the ground and/or engine operating conditions.

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Isolation Valve Light Operation The ISOLATION VALVE light provi4es visual indication of isolation valve condition (energized or deenergized) and the corresponding hydraulic pressure (high or low) downstream of the valve. The light operation is as follows: AIRPLANE ISOLATION HYDRAULIC ISOLATION POSITION VALVE PRESSURE VALVE LIGHT A. Ground Energized High Off B. Ground Energized Low On C. Air Deenergized Low Off D. Air Deenergized High On Case A With isolation valve energized normally or by the OVERRIDE switch in OVERRIDE position, transistor Q3 is biased. Bias power for transist Q4 is removed by transistor Q3 and hydraulic pressure switch in high pressure position. Ground for ISOLATION VALVE light is not available, since transistor Q4 is not conducting. Case B. With isolation valve energized normally or by the OVERRIDE switch in OVERRIDE position, transistor Q3 is biased. Bias power for transistor Q4 is now available since hydraulic pressure switch is in low pressure position. Ground for ISOLATION VALVE light is available through transistors Q4 and Q3. Case C. With isolation valve deenergized, transistor Q3 is not biased. Bias power for transistor Q4 is available since hydraulic prcssuru switch is in low pressure position and transistor Q4 not conduct tng. Ground for ISOLATION VALVE light is not available through transistor Q3 or hydraul to pressure switch. Case D. With isolation valve deenergized, transistor Q3 is not biased. Bias power for transistor Q4 is available since transistor Q4 Is not conducting. Ground for ISOLATION VALVE light is available through transistor Q4 and hydraulic pressure switch.

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BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00


POWER PLANT EXHAUST

Figure 7 SCL

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Jan -- 2004

T/R Isolation Valve Circuit Page: 15


BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00


POWER PLANT EXHAUST

Figure 8 SCL

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Jan -- 2004

Hush Kit and Exhaust Mixer Page: 16

BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00 BOTH MAIN GEAR UP



POWER PLANT EXHAUST

Figure 9 SCL

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Jan -- 2004

T/R Summary Page: 17

BOEING 737 - 200 JT8D -- 7 to -- 17 78 -- 00

NOTES:



POWER PLANT EXHAUST

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