✰ FM 1-506

FIELD MANUAL
NO 1-506

HEADQUARTERS
DEPARTMENT OF THE ARMY
Washington, DC, 30 November 1990
FUNDAMENTALS OF AIRCRAFT POWER PLANTSi

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PREFACEThis manual provides information on the operation, components, and systems of aircraft power plants. The
turbine engine is relatively new to the aviation field and its technology is growing rapidly. This manual will bringusers up-to-date on new developments in the field.
Propellers have been deleted from this manual since the Army no longer works on them.
This manual is for use by Army aviation mechanics worldwide.
For specific instructions on the aircraft power plants in particular types and models of aircraft, refer to applicablemaintenance manuals. Should the information in this field manual and that in a specific aircraft maintenancemanual conflict, the latter takes precedence.
The proponent of this publication is HQ TRADOC. Submit changes for improving this publication on DA
Form 2028 (Recommended Changes to Publications and Blank Forms) and forward it to Assistant Commandant, US Army Aviation Logistics School, ATTN ATSQ-LTD-L, Fort Eustis, Virginia 23604-5421.
Unless otherwise stated, whenever the masculine gender is used, both men and women are included.
This manual contains copyrighted material as follows:
Pages 3-7 thru 3-8,3-11 thru 3-12,3-14 thru 3-15,3-17 thru 3-20; 4-17; 8-1,8-5 thru
8-8; 9-1,9-3, thru 9-4,9-6, 9-7, thru 9-12; Fig 3-19, Fig 3-20: Copyright ©, 1951,
1974, United Technologies Corporation.

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PART ONE. FUNDAMENTALS OF POWER PLANTS
CHAPTER 1

BASIC REQUIREMENTS

Aircraft power plants must meet exacting requirements for dependability and endurance. Many difficult
engineering problems have been overcome in an effort tosatisfy these requirements with further advances being
made each day. Requirements imposed on aircraft
power plants in an effort to obtain engines suitable foraircraft include –Reliabtity.
Durability.
Compactness.
Low weight per horsepower.
High specific power output.
Reasonable cost.
High thermal efficiency.
Freedom from vibration.
Ease of maintenance.
Operating flexibility.
DEFINITION OF TERMS
Reliability
Reliability is the most important fundamental power
plant requirement. In the air each working part, nomatter how small is important. Only by careful attention
to the smallest detail can aircraft power plant manufacturers and mechanics assure power plant reliability.
Durability
Durability is the measure of reliable engine life. The
durability realized by an engine depends largely on thetype or condition of operation. Intelligent application ofoperation and maintenance procedures results in greatly
improved power plant durability.
Compactness
Compactness is essential to power plant design inorder to lower parasitic drag and to attain higher speeds.

The overall form an engine takes is determined to a highdegree by the compactness required. The degree ofcompactness that may be achieved is limited by thephysical requirements of the engine. For example, compactness is limited on radial air-cooled engines duet to thefrontal area required for sufficient cooling of engine
cylinders.
Low Weight per Horsepower
Minimum weight per horsepower (HP) is a primary
requirement in aeronautics. The weight of a power plantmust be kept as low as possible. This allows the aircraft
to carry a large useful load with a satisfactory margin ofsafety in proportion to gross weight. The larger modern
aircraft reciprocating engines have attained a horsepowerto-weight ratio of 1 horsepower to 1 pound weight. Gas
turbine engines currently used by the Army have a greater
horsepower-to-weight ratio. A good example is the
T-55-L-712, which develops 4500 shaft horsepower(SHP) and weighs 750 pounds(dry) (6.0 HP per pound).
High Specific Power Output
Power output is based on engine size, RPM, andweight for the fuel-air mixture. Size and RPM are limitedin the reciprocating engine. Therefore, an increase in theeffective working pressure in the cylinders is one of themost valuable ways to increase the specific power output.
Greater pressure increases are possible by supercharging
the engine (comprising the mixture before it enters thecylinders). The limiting factors in increasing cylinderpressure are resistance of the fuel to detonation and the
maximum allowable cylinder pressure.
Reasonable Cost
Despite perfection of design and quality ofworkmanship, no power plant will be desirable if it is too
costly in a competitive market. A primary factor determining the usefulness of an engine is its cost. Because of

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the raw materials and the great number of man-hours
involved, complex designs drive up the cost. The mostsatisfactory design is generally the simplest that will meetrequirements.
High Thermal Efficiency
Thermal efficiency is a measure of the losses sufferedin converting heat energy in the fuel into mechanicalwork; it is the ratio of the heat developed into useful work
to the heating value of the fuel. High thermal effciency,
therefore, means high fuel economy-something ofgreat importance in aircraft engines. The less fuel required for a military mission, the greater the military loadthat can be carried and the lower the fuel cost.
Freedom From Vibration
A power plant that is free from vibration is importantin the light, somewhat flexible aircraft structure since
severe engine vibration will in some cases reduce the lifeof certain structural parts. The need for freedom fromvibration is met usually by using a large number ofcylinders to offset the vibration torque delivered by theindividual cylinders. Counterweights are installed on
crankshafts to balance rotating masses. These are usually
hinged to provide dynamic damping of vibration which
results from power impulses and to counteract undesirable torsional or twisting vibration. Also, flexible
engine mount isolators are used to permit certain movements of the power plant that are harmful to aircraft
structures.
Ease of Maintenance
The requirement of ease of maintenance is especially important to military operations in the field.
Simplicity of design and use of standard parts, whenpossible, assist in keeping maintenance at a low level.
Operating Flexibility
Flexibility is the ability of a power plant to runsmoothly and provide the desired performance at all
speeds from idling to maximum power output. The wide
range of operating requirements demanded of aircraft
engines presents difficulties rarely encountered in otherpower plant fields. In addition to the requirement of
unfailing reliability, the engine must operate in widely
varying positions, altitudes, and atmospheric conditions.
FUELS
Requirements
An engine fuel must be tailored to an engine and viceversa since there must be enough quantities of fuel available
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to the engine. Some significant properties of aviation fuels
are discussed below.
Heat Energy Content or Net Heating Value. Theenergy content or heating value of a fuel is expressed inheat units (British thermal units [BTUs]). A fuel satisfactory for aircraft engines must have a high heat energy
content per unit weight. A high heat energy content
causes the weight of fuel carried to be lower than a lowheat energy content. Then more of the load-carrying
capacity is available for the payload Aviation gasoline
and JP fuels are very desirable from this standpoint. The
heat energy content for aviation gasoline is about 18,700
BTUs/pound, and for JP fuels about 18,200
BTUs/pound. The various alcohols, which have maximum energy content of about 12,000 BTUs/pound, dopossess some other desirable characteristics as an internal combustion engine fuel.
Volatility. A volatile liquid is one capable of readilychanging from a liquid to a vapor when heated or whencontacting a gas into which it can evaporate. Since liquid
fuels must be in a vaporous state to burn volatility is an
important property to consider when choosing a suitable
fuel for an aircraft engine. Volatility determines the starting accelerating vapor-locking and distribution characteristics of the fuel. Gasoline and JP fuels are verysatisfactory because they can be blended during the refining process to give the desired characteristics. Because of
the nature of constant pressure combustion in gas turbine
engines a highly volatile fuel is not necessary. JP fuels areof rather low volatility while aviation gasoline is highly
volatile. Comparing a highly volatile fuel like aviation
gasoline to a less volatile one like JP fuel the following
effects become apparent. The highly volatile fuel –Starts easier in cold temperatures.
Has a slightly better combustion efficiency.
Leaves less deposit in the combustion chamberand on the turbine blades.
Is a greater fire hazard.
Creates a greater danger of vapor lock of thefuel system.
Has high evaporation losses through thebreather of the fuel tank at high altitudes.
NOTE: The last two difficulties are practically nonexistent with fuels having low
volatility.
Stability. The fuels used in aircraft engines must be
stable. Because aviation fuels are sometimes stored forlong periods, they must not deposit sediment. The gums

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that are normally formed are insoluble in gasoline and JP
fuels and may cause restrictions in fuel strainers andliners. Aviation fuel must also retain its original properties during storage.
Purity. Aviation fuel must be free from water, dirt,
and sulfur. Small amounts of water will not usually cause
any difficulty because water can be removed from the fuelsystem by draining. Large amounts of this impurity, however, can cause complete engine failure. It is very important that corrosive sulfur be eliminated from fuel. The
sulfur content of fuel may form corrosive acids whenbrought in contact with the water vapor formed in the
combustion process.
Flush Point. The flash point is the lowest temperature at which fuel will vaporize enough to forma combustible mixture of fuel vapor and air above the fuel. It isfound by heating a quantity of fuel in a special container
while passing aflame above the liquid to ignite the vapor.
A distinct hash of flame occurs when the flash pointtemperature has been reached.
Fire Point. The fire point is the temperature whichmust be reached before enough vapors will rise to
produce a continuous flame above the liquid fuel. It is
obtained in much the same manner as the flash point.
Reid Vapor Pnessure. Reid vapor pressure is theapproximate vapor pressure exerted by a fuel whenheated to 100% This is important because it is used todetermine when a fuel will create a vapor lock.
Specific Gravity. Specific gravity is the ratio of thedensity (weight) of a substance (fuel)compared to thatoof an equal amount of water at 60 F. Specific gravity isexpressed in terms of degrees API. Pure water has a
specific gravity of 10. Liquids heavier than water have anumber less than 10. Liquids lighter than water have anumber greater than 10. An example is JP-4, whose
specific gravity in degrees API is 57. The American
Petroleum Institute (API) has chosen pure water bywhich to measure the specific gravity of fuels.
NOTE: Both flash and fire points give a
relative measure of the safety properties offuel a high flash point denotes that a hightemperature must be reached before
dangerous handling conditions are encountered. The minimum flash point permitted in a fuel is usually written into thespecifications.

Grades
Turbine fuels are high-quality fuels covering thegeneral heavy gasoline and kerosene boding range. They
do not contain dyes or tetraethyl lead.
One of the major differences between the wideboiling and kerosene types is the fuel volatility. JP-4
fuels have a wider boding range with an initial boilingpoint considerably below that of kerosene. As a groupthese fuels have lower specific gravities than kerosene
types. Wide-boil-range fuels have Reid vapor pressures of 2 to 3 pounds and flash points below room
temperature. Kerosene-type fuels have Reid vapor pressuresof less
than 0.5 pound and flash points higher thanool00 F (38 C). Wide-boiig-range fuels generally havelower freezing points than kerosene fuels.
The fuel authorized for Army aircraft gas turbine
engines is JP-4. The letters “JP" stand for jet propulsion;
the number 4 indicates fuel grade.
Military specification MIL-T-5624 covers JP-4, JP-5,
and JP-8 fuels. Jet A, Jet Al, and Jet B are commercial
fuels which conform to the American Society for Tinting
Materials specification ASTM-D-1655.
Jet B is a JP-4 typefuel;o its freezing point is -56°Foo(-49 C) instead of -72 F (-58 C) for JP-4.
JP-5, Jet 4 and Jet A-1 are kerosene-type fuels.
ASTM Jet A and A-1 differ primarily in their fuelfreezing points. Jet A o is considered
suitable down
to ofuelootemperatures of -36 F (-38 C); Jet A-1, to -54 F (48 C).
JP-4 is a fuel consisting of approximately 65 percent
gasoline and 35 percent light petroleum distillate, withrigidly specified properties. JP-4 is currently the Armystandard fuel for turbine engines.
JP-5 is a specially refinedkerosene having a minioomum flash point of 140 F and a freezing point of -51 F(-46°C).
JP-8 is a speciallyrefined kerosene with a minimumoflash point of 110 F and a freezing point of -54°F (-48°C).
This fuel is being classified as a total replacement fuel forall of NATO. It will replace all fuels currently used inmilitary equipment from generators to tanks to aircraft
and even to trucks. This classification will ease logistics incombat. Having only one fuel for all equipment preventsaccidentally mixing or using the wrong fuels. To datetesting of JP-8 is proceeding well with the total single-fuelconcept on its way to full fielding.
JP fuels vary from water white to light yellow; color
coding, however, does not apply to these fuels.

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Additives in JP fuels include oxidation and corrosion
inhibitors, metal deactivators, and icing inhibitors. Icing
inhibitors also function as biocides to kill microbes inaircraft fuel systems.
Should mixing of JP fuels become necessary, there
is no need to drain the aircraft fuel system before adding
the new fuel. Due to the different specific gravities ofthese fuels, mixing them will affect the turbine engine’s
performance. Be sure to consult appropriate technicalmanuals for additional information and procedures.
When changing to a fuel with a different specificgravity, externally adjusted fuel controls and fuel flow
dividers on some engines may require retrimming or
readjustment for optimum performance.
COMBUSTION
Mixing fuel with air and burning it would seem to bea very simple process, but this apparent simplicity is
deceptive. Problems encountered are with distribution
knot.kj ignition timing, and so forth. In an internal combustion engine, the combustion process is the rather rapid
reaction between fuel and oxygen. This process liberates
the potential energy contained in the fuel supplied to theengine. In a gas turbine engine atmospheric air is taken
in and compressed; fuel is then burned in the compressed
air, which then expands through a turbine that drives acompressor.
The combustion problem would not be so great if
weight and space were not so important in aircraft gas
turbine engines. Without such limitation the air supply
for the compressors could be divided. A portion suitable
to the desired power output could be burned at approximately the chemically correct fuel-air ratio in a
low-velocity combustion chamber. Design of the combustion chamber is such that less than a third of the total
volume of air entering the chamber is permitted to mix
with the fuel. The excess air bypasses the fuel nozzles andis used to cool the hot surfaces and to mix with and cool
the burned gases before they enter the turbines.
Temperature Limitations
Gas turbine engines produce work in proportion to
the amount of heat released internally. Most of this heat
is obtained by burning fuel although some heat originates
when air is compressed in the compressor. Low fuel-air
ratio is required to keep the temperature of the gases
delivered to the turbine down to a value which the turbine
wheel can tolerate. With present materials of construction, the highly stressed blades or buckets of the turbine
wheel cannot stand a temperature of more than about
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1500 F (815 C); therefore, the gases at entrance
to theooturbine vanes cannot exceed about 1600 F (898 C) for
more than brief periods.
Fuel-Air Ratio
The fuel-air ratio is so low in gas turbine engines that
if fuel and air were uniformly mixed, the mixture would
not ignite. Complete combustion of both fuel and air is
obtained with isooctane at a fuel-air ratio of 0.066. Thisis the theoretical value which would produce complete
combustion if given sufficient time to reach equilibrium.
In practice, complete combustion of either fuel or air
requires an excess of one or the other. Thus, if all of the
air is to be burned excess fuel must be present and afuel-air ratio of about 0.080 is required. Control of power
output is largely determined by means of fuel-air ratio.
Increasing fuel-air ratio increases the quantity of air andthe temperature at which it is discharged at the exhaust
jet pipe.
Rich-Mixture Blowout
Combustion efficiency in a well-designed combustion chamber using the most favorable fuel may be as highas 98 percent at sea level. On the other hand, it may fall
to as low as 40 percent at extremely high altitudes with abadly designed combustion chamber and unsuitable fuelAs combustion efficiency is reduced, a point is reachedwhen the turbine does not develop enough power to drive
the compressor. Increasing the fuel supply in order tomaintain or increase engine speed may not result inincreased engine RPM, and the unburned fuel may extinguish the flame. This is known as rich-mixture blowout.
Lean-Mixture Die-Out
In contrast to rich-mixture blowout, lean-mixture
die-out occurs when the mixture is too lean to bum underconditions of efficient combustion. When combustion
efficiency is very low, the mixture may be ignited from anexternal source (such as an igniter plug). It will theneither be extinguished when the spark ceases or may bum
so slowly that the flame is carried out through the turbine.
Lean-mixture die-out can also occur when the fuel supply
is reduced in order to decrease engine speed.
Rich mixture blowout and lean mixture die-out havebeen virtually eliminated through the refinement of fueldelivery systems.
LUBRICATION
Lubrication is a very important part of power plant
operation. An engine allowed to operate without

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lubrications certain to fail. Lubrication not only combats
friction but also acts as a cooling agent.
WARNING
Never mix reciprocating engine oils and gas turbine
engine oils; they are not compatible. Mixing themcauses engine failure!.
The primary purpose of a lubricant is to reduce
friction between moving parts. Because liquid lubricants(oils) can be circulated readily, they are used universallyin aircraft engines. In theory, fluid lubrication is based on
actual separation of the surfaces so that no metal-tometal contact occurs. As long as the oil film remains
unbroken, metallic fiction is replaced by the internal
fluid friction of the lubricant. Under ideal conditions
friction and wear are held to a minimum. In addition to
reducing friction, the oil film acts as a cushion between
metal parts. This cushioning effect is particularly important for such parts as reciprocating engine crankshaft andconnecting rods, which are subject to shock loading. As
oil circulates through the engine, it absorbs heat from theparts. Pistons and cylinder walls in reciprocating engines
especially depend on oil for cooling. Oil also aids informing a seal between the piston and cylinder wall toprevent gas leaks from the combustion chamber. Oils also
reduce abrasive wear by picking up foreign particles andcarrying them to a filter to be removed
Friction Reduction
Lubricating oil decreases friction by preventing
metal-to-metal contact at bearing points throughout theengine. Separating mating surfaces of moving parts by athin film of oil changes dry or solid friction to fluid friction.
The result is less heat generated in the moving parts anddecreased wear on the parts.
Cushioning EffectLubricating oil cushions bearing surfaces by absorbing the shock between them.
Cooling
It has been noted that reducing friction results in lessheat being generated. Also, as oil is circulated throughbearings and splashed on various engine parts, it absorbs
a great amount of heat. Lubrication is particularly important in reciprocating engines to cool the piston andcylinder. An efficient lubrication system will absorb asmuch as 10 percent of the total heat content of fuelconsumed by the engine. By carrying away this heat, theoil flow reduces operating temperatures of internal parts
not directly cooled by the engine cooling system.
Sealing Action
Oil helps seal mating surfaces in the engine, and thefilm of oil on various surfaces is an effective pressure seal.
In reciprocating engines the oil film between the cylinderwall and piston and piston rings is important in retaining
the high gas pressure in the cylinder.
Cleaning EffectOil cleans the engine by picking up carbon and otherforeign particles as it passes through and around engine
parts. It carries these particles through the system to astrainer where they are filtered from the oil.
LUBRICATING OIL REQUIREMENTS
The conditions which the engine operates under
determine the requirements for lubricating oil. Conditions like temperature, contact pressure, and type andrate of motion vary so much that one lubricant cannot
provide ideal lubrication for all components. Using alubricating oil with all the desirable properties in*degrees will provide satisfactory results. TB 55-9150-200-24
specifies engine oils for use in Army aircraft. Some desirable
lubrcating oil qualities are –Viscosity.
Antifriction ability.
Cooling ability.
Chemical stability.
Viscosity
The degree of resistance of an oil flow at a specified
temperature indicates its viscosity. An oil that flows slowly
is described as a viscous oil or an oil of high viscosity. An
oil that flows readily is said to possess a low viscosity. The
viscosity of all oils is affected by temperature. AS thetemperature increases, oil becomes thinner. The rate at
which an oil resists viscosity changes through a given
temperature range is called its viscosity index The viscosity of aircraft engine oil is fairly high because of highoperating temperatures high bearing pressures, andrelatively large clearances inside an aircraft engine. Since
aircraft engines are also subjected to a wide range oftemperatures, an oil with a high viscosity index is required.
Antifriction
The theory of fluid lubrication is based on the actual
separation of metallic surfaces by an oil film. Lubricantsshould have high antifriction characteristics to reduce1-5

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frictional resistance of the moving parts and high antiwear
properties to resist the wearing action that occurs during
engine operation.
Cooling Ability
Lubricating oil should have maximum cooling ability
in order to absorb as much heat as possible from all
lubricated surfaces.
Chemical Stability
The extreme operating conditions and high dollar
value of aircraft engines make it necessary to use lubricating oil of the very best quality. The following chemicaltransformations can occur in a lubricating oil and makeit unfit for service.
Acidity. Acidity in oils is dangerous chiefly whenhigh temperatures and moisture are present as is the casein aircraft engines. The results of a high acid content arecorrosion of metal and the formation of sludges, emulsions, and deposits in the oil system.
Oxidation. All lubricating oils tend to oxidize whenin contact with air. The compounds formed by oxidation
are undesirable and harmful since they are generally of agummy or acid character.
Sulfur. Sulfur may occur in lubricating oil as freesulfur or as sulfonates. Free sulfur may be present throughcareless or improper refining or by actual addition. Freesulfur is corrosive in nature and impairs the stability ofthe oil. The presence of sulfonates indicates overtreatment with sulfuric acid or inadequate washing of the oil
to free it of chemicals during the refining process. Sulfonates are strong emulsifying agents that tend topromote sludge formation in the oil system. Lubricatingoils containing high percentages of sulfur oxidize moreeasily than those with low sulfur content.
Carbon Residue. Petroleum lubricating oils arecomplex mixtures of hydrocarbons that vary widely intheir physical and chemical properties. Owing to thesedifference some oils may vaporize under atmosphericconditions without leaving any appreciable residues.
Other oils leave a nonvolatile carbon residue upon
vaporization. This carbon residue is the result of a partial
breakdown of the oil by heat, which is caused by destructive distillation of the oil without air entering into thereaction. Many parts in the engine operate at a temperature high enough to cause this reaction and to form
carbon deposits. These deposits are undesirable as they
may restrict passages. In reciprocating engines they may
cause sticking piston rings and sticking valves.

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Moisture. Corrosion of bearing metals is almost
always due to moisture in the oil. Although it is possible
for corrosion to occur from acidity, it is unlikely unlessmoisture is also present. Therefore, to prevent corrosion,
it is important to eliminate moisture from the oil as muchas possible. This is especially true at high temperatures
because they increase the rate of corrosion.
Because of the accumulation of these harmful substances common practice is to drain the entire lubrication system at regular intervals and refill it with new oil.
The time between oil changes varies with each make andmodel aircraft and engine combination.
WARNING
When handling oil used in gas turbine engines, do notallow oil to remain on skin any longer than necessary. It
contains a toxic additive that is readily absorbed
through skin.
GAS TURBINE ENGINE LUBRICANTS
There are many requirements for turbine engine
lubricating oils; but because of the small number ofmoving parts and the complete absence of reciprocating
motion, lubrication problems are less complex in theturbine engine. This, together with the use of ball androller bearings, requires a less viscous lubricant. The
turboprop engine, while using essentially the same typeof oil as the turbojet, must use a higher-viscosity oil
because of the higher bearing pressures introduced by
the highly loaded propeller reduction gearing.
Gas turbine engine oil must have high viscosity forgood load-carrying ability but must also have viscosity low
enough to provide good flow ability. It must also be of low
volatility to prevent loss by evaporation at the high altitudes at which the engine operates. In addition, the oilshould not foam and should be essentially nondestructive
to natural or synthetic rubber seals in the lubricating
system. Also, with high-speed antifriction bearings, theformation of carbons or varnishes must be held to aminimum.
The many requirements for lubricating oils are met
in the synthetic oils developed specifically for turbine
engines. Synthetic oil has two principal advantages overpetroleum oil: it tends to deposit less lacquer and cokeand evaporates less at high temperatures. Its principal
disadvantage is that it tends to blister or remove paintwherever it is spilled Painted surfaces should be wipedclean with petroleum solvent after spills.

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Oil-change intervals for turbine engines vary widelyfrom model to model. They depend on the severity of oil
temperature conditions imposed by the specific airframe
installation and engine configuration. Follow the applicable manufacturer’s instructions.
Synthetic oil for turbine engines is usually suppliedin sealed l-quart or l-gallon metal cans. Although thistype of container was chosen to minimize contamination,
it is often necessary to filter the oil to remove metal slivers,
can sealants, and so forth, which may occur when opening
the can.
Some oil grades used in turbojet engines may containoxidation preventives, load-carrying additives and substances that lower the pour point, in addition to synthetic
chemical-base materials.
BEARINGSIn Army aircraft bearings are found throughout thepower train system from the engine to the rotor orpropeller. The failure of anyone of these bearings would
place the entire aircraft in jeopardy. It is crucial that they
be properly serviced and maintained. In fact, bearings are
considered so important that a major air raid was conducted at considerable sacrifice during World War II in
an attempt to eliminate one of Germany's principal bearing manufacturing centers. If the Allies could havedestroyed Germany’s capacity to produce antifriction
bearings, its entire aviation effort would have come to an
immediate standstill. Today, our economy could not function without bearings.
Bearings have the following functions. They–
Support the load on a shaft. The load may be awheel, pulley, or gear, or it may be the turning
shaft itself.
Reduce friction created by turning. This is accomplished both by design and by lubricationand is one of the most important functions ofbearings.
Reduce friction created by thrust. A speciallydesigned bearing is required for this purpose.
Hold a shaft in rigid alignment. A high-speed
rotating shaft has a tendency to “whip” unlessadequately supported by bearings.
Provide a place for adjustment.
Bearings are classified into two broad categories
Friction– also commonly known as a plain or
babbitt type; it makes a sliding contact with theshaft.
Antifriction– contains rollers or balls; it makes
a rolling contact with the shaft.

The term “antifriction” has long been used to differentiate rolling and sliding bearings. The word is
gradually being dropped in deference to the word “rolling" to describe ball and roller bearings. The term’’rolling
bearings’’ will be used in this manual to describe all
bearings consisting of ball or roller elements that rollbetween concentric inner and outer rings. The term
“plain or journal bearings” will be used to describe twopiece bearings where the two rotating surfaces are sliding
with respect to each other.
Rolling Bearings
A brief review of history reveals the long gradualdevelopment of rolling bearing. They were developed toreduce friction, thereby increasing work output while
reducing energy input. Rolling-type bearings changed
sliding friction into rolling motion with a greatly reduced
friction level.
Assyrians and Babylonians as far back as 1100 BC
used round logs as rollers to move huge monuments andstones. With the rollers placed under the load, it could
more easily be pulled. AS the rollers came out of the rearof the load, they were carried around to the front andplaced in the oncoming path of the load
The wheel was an extremely important development; however, it was based on sliding friction. The useof lubricants did reduce friction and increase load andlife expectancy of the wheel. But it was not until the axle
rested on balls or rollers that sliding friction was finallychanged to rolling motion and the fiction level significantly reduced.
Modern rolling bearing had their origin with thegreat inventor, Leonardo da Vinci, about 1500 AD. Manyof his drawings show the use of balls and rollers andforeshadow such modem technology as surface finish,
raceway grooves, and conical pivots.
Rolling bearings are also classified into two types
ball and roller.
Ball Bearings. The ball bearing is one of the mostcommon used in aircraft. With relatively minor variations
it can be adapted to many different uses. It creates theleast amount of friction of any common bearing because
the ball itself is the best antifriction rolling device known.
The ball maintains point contact with the surface it rolls
on and reduces friction to a minimum. It is, therefore,
best suited to high-speed applications.
Roller Bearing. The roller bearing makes use of acylindrical-shaped roller between the fiction surfaces.
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Since it is a cylinder, it will make line contact rather thanpoint contact. It is therefore more suited to heavy loads
because the weight is distributed over a larger cotact
area.
Construction Features. Refer to Figuree 1-1. The
rolling elements in rolling bearings are provided with

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both an inner track and an outer track on which to roll.
These tracks are known as “races.” The races form aprecision, hardened, and true surface for the balls orrollers to ride in. The balls or rollers are held together andspaced evenly around the bearing by means of a cage orseparator. Seals are used on some bearings to keep out
dirt and to keep grease in.

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Dimensional interchangeability does not necessarily
indicate functional interchangeability. Therefore, somebearings may be suited for thrust or axial loads whileothers are not. Bearing design (size and number of balls,
depth and type of groove, width and thickness of races,
construction of the separators) will determine the loadand speed for which the bearing can be used. Bearings
made to take both radial and thrust loads will be increased on one side of the outer race and will usually bestamped “thrust.”
Plain Bearings
Plain bearings are classified as split type or solid type.
Split Type. The lower end of an engine connecting
rod is a good example of a split bearing. It may beinstalled and used on a shaft in locations that preclude
use of other bearing types. Another example of split
bearings is the main bearing mounts supporting the engine crankshaft.
Solid Type. An engine connecting rod also furnishes a good example of a solid friction bearing thepiston pin bearing or bushing. Bearings of this type canbe used only where it is possible to slip them over the endof the shaft on which they run.
Construction Features. Refer to Figure 1-2. A basic
requirement for the plain bearing is that it and the shaft
be made of dissimilar metals. A steel bearing surface
could never be used with a steel shaft. The bearing metalis always softer than the steel shaft, yet it must be hardenough to provide adequate heat transfer and possessgood wear qualities. The metal used in bearing is known
as babbitt.

placed in a split-type bearing housing produce a unit
which is both efficient and easily maintained. Bearing
inserts can be mass-produced to very close tolerances andrequire little or no fitting during assembly. The bearing
is renewed by merely taking the housing apart, replacing
the inserts, and bolting it up again.
Inserts on aviation engines are often of the tri-metal
variety. The steel shell will have a layer of silver on its
inner surface and a layer of babbitt on top of the silver.
This bearing is durable and provides excellent heat transfer. It is typically used for crankshaft and rod end bearings (Figure1-3).

Bearing inserts must always be secured in the housing they must never turn with the shaft. Split-type inserts
will have a tang on each half which will fit into a notch in
the housing. Solid-type inserts (bushings) are pressed
into the housing. Either method will preclude the chanceof turning in the housing. Inserts that must absorb thrust
or axial loads are designed with flanges along the sides totake this side load. Bearings without this side flange mustnever be subjected to thrust or axial loads.
Teflon-Lined Bearings
A new type of bearing finding increasing use is the
Teflon-lined bearing. It offers high reliability and easy
maintenance. It is self-lubricating, chemically inert, andshock-resistant; and it has a low coefficient of friction. It
is designed as a spherical or as a journal bearing and also
as a rod end bearing. Refer to Figures 1-4 and 1-5.

A typical friction bearing contains babbitt metal
inserts in the bearing housing. A bearing insert consistsof a steel shell on which a layer of babbitt has been
applied The shell is actually two half shells that when1-9

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Bearing Loads
The two types of bearing loads are –Radial. When a wheel turns, it producesocentrifugal force. When this force is 90 fromthe shaft, it is known as radial loading.
Axial/Thrust. When the force or load on a bearing is in a direction parallel to the shaft, it is
known as axial/thrust loading
NOTE: Thrust loads are the same as axial
loads. Whenever the term “thrust load” isused throughout this manual, it will also
refer to axial loads.
Many times a bearing is subject to a combination ofboth radial and thrust leak; for example, the engine
crankshaft main bearings. The throws on the crankshaft
produce radial load while the pull of the propeller
produces thrust load. On high-horsepower engines thethrust load produced by the propeller is so great that theprop shaft could not turn unless this load (and the resulting friction) were absorbed by a special thrust bearing in
the engine front section.

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CHAPTER 2BACKGROUND, DEVELOPMENT, AND THEORYIn their early attempts at flight humankind soon
learned that the human body was totally inadequate as a
propulsion device. They also learned that heavier-thanair aircraft would need power to attain and sustain flight.
Early efforts to develop a satisfactory engine included
rocket, steam, jet, and reciprocating engines, but it was
the reciprocating engine that first pushed the Wright
brothers aloft. Today’s Army aircraft are propelled by gas
turbine and reciprocating engines. While the technology
which led to these engines is relatively new, the idea is not.
EARLY DEVELOPMENT
The Aeolipile
Hero, an Egyptian scientist from Alexandria,
developed the first "jet engine" about the year 100 BC.
Known as the "aeolipile" (Figure 2-1), it consisted of aboiler, two hollow bent tubes mounted to a sphere, andthe sphere. Steam coming from the boiler enteredthrough the two hollow tubes supporting the sphere. The
steam then exited through the bent tubes on the sphere,causing it to spin. Hero is said to have used this invention
to pull-open temple doors.

The Chimney Jack
Leonardo Da Vinci designed a device called the"chimney jack" around the year 1500 AD. The chimney
jack was used to turn a roasting skewer. This reactiontype turbine worked on the principle of heat raising gases
from the roasting tire. As the hot air rose, it passed
through fanlike blades that turned the roast through aseries of gears.
Rockets
With the discovery of gunpowder around 1000 AD
the Chinese, probably accidentally through a defective
firecracker developed the rocket. Rockets were used as
early as 1232 by the Mongols to create fear among enemy
troops. A Chinese scholar, Wan Hu attempted to propel
his sled with rockets. He became the first martyr inhumankind’s attempts to fly. Rockets were used in the
Napoleonic Wars and the War of 1812. The Germans
used V-2 rockets in World War II The most noted use ofrockets is to launch space vehicles.
Branca’s Stamping MillIn 1629 an Italian engineer, Giovanni Branca, wasprobably the first to invent an actual impulse turbine.
This device, a stamping mill (Figure 2-2), was generated
by a steam-powered turbine. A jet nozzle directed steam
onto a horizontality mounted turbine wheel, which thenturned an arrangement of gears that operated his mill.

The Steam Wagon
In 1687 Isaac Newton attempted to put his newly
formulated laws of motion to the test with his “steam

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wagon" (Figure 2-3). He tried to propel the wagon bydirecting steam through a nozzle pointed rearward
Steam was produced by a boiler mounted on the wagon.
Due to lack of power from the steam, this vehicle didn’t
operate.

MODERN AVIATION
EnglishIn January 1930 an Englishman, Frank Whittle, submitted a patent application for a gas turbine for jet
propulsion. It wasn’t until the summer of 1939 that the
Air Ministry awarded Power Jets Ltd a contract to design
a flight engine. In May 1941 the Whittle W1 engine(Figure 2-5) made its first flight mounted in the Gloster
Model E28/39 aircraft. This airplane would later achievea speed of 370 MPH in level flight with 1000 pounds ofthrust.

The First Gas Turbine
In 1791 John Barber, an Englishman. was the first to
patent a design that used the thermodynamic cycle of themodem gas turbine. His design (Figure 2-4) contained
the basics of the modem gas turbine it had a compressor,
a combustion chamber, and a turbine. The main difference in his design was that the turbine was equipped
With a chain-driven reciprocating type of compressor.
He intended its use for jet propulsion.

GermanThe Germans, Hans von Ohain and Max Hahn,
students at Gottingen, seemingly unaware of Whittle’s
work patented a jet propulsion engine in 1936. Ernst
Heinkel Aircraft Company adapted their ideas and flew
the second aircraft engine of this development in an
HE-178 aircraft (Figure 2-6) on August 27, 1939. This
was the first true jet-propelled aircraft. The engine,
known as the Heinkel HES-36, developed 1100 poundsof thrust and hurled the HE-178 to speeds of over 400
MPH. This engine used a centrifugal flow compressor.
Later development produced a more advanced turbine
engine that used an axial flow compressor. This turbine
was used to power the ME262 jet fighter to 500 MPH.
These planes were introduced in the closing stages of2-2

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World War II. The more modern engine features of blade
cooling, ice prevention, and the variable-area exhaust
nozzle were incorporated into the ME262 aircraft engine.
0n September 30, 1929, using a modified glider and Opelrockets, the Germans were the first to achieve flight using
a reaction engine.

Italian
An Italian, Secundo Campiri of the Caproni Company, developed a turbine engine that used a 900-HP
reciprocating engine to drive its three-stage compressor.
This turbine was installed in the Caproni-Campiri CC-2,
but it only achieved a disappointing 205 MPH. In August
1940 the CC-2 made its first flight, with the whole projectending just eight years later.
AmericanWith the help of the British, America entered the jet
propulsion field in September 1941. The W.lX engine, acomplete set of plans for the W.2B engine, and a groupof Power Jets engineers were flown to the United Statesfrom Britian. A contract was awarded General Electric
Corporation to develop an American version of The
W.IX. One year later (October 1942) the Bell XP-59A(Figure 2-7), fitted with two General Electric l-A engines(rated at 1300 pounds of thrust each) made this country’s
first jet propulsion flight.

Although General Electric introduced America
to jet propulsion with its modified Whittle design,
Westinghouse Corporation gave America its own gas

turbine engine. The Westinghouse engine included anaxial compressor and an annular combustion chamber.
These two Westinghouse designs (axial compressor andannular combustion chamber) or variations thereof arestill being used in turbine engines.
NEWTON’S LAWS OF MOTIONFirst Law
Newton’s first law states a body in a state of rest
remains at rest; a body in motion tends to remain inmotion at a constant speed and in a straight line unlessacted upon by some external force.
The first part of this law is evident from everyday
experience for instance, a book placed on a table stays
on the table. The second part of the law is more difficult
to visualize. It states that if a body is set in motion and left
to itself, it virtually keeps on moving without the action of
any further force. The statement is correct; the body
would continue to move without any reduction in velocity
if no force actcd upon it. However, experience shows thata retarding force (friction) is always present. A block ofwood thrown to slide along a cement-paved road comesto rest after sliding only a short distance because thefriction is great; along a waxed floor it would slide farther
because the friction is less, along a sheet of ice it would
slide much farther because the fiction is much less. From
examples like these it is theorized that if friction were
eliminated entirely, a body once set in motion on a levelsurface would continue to move indefinitely with undiminished velocity.
Second Law
Newton’s second law states: an unbalance of forceson a body produces or tends to produce an accelerationin the direction of the greater force; the acceleration is
directly proportional to the force and inversely proportional to the mass of the body.
The meaning of the second law is illustrated by theexample of two identical boxes being moved across afloor. If more force is applied to one box than to the other,
the one subjected to the greater force will have greater
acceleration. Or if a full box and an empty one are drawn
across the floor with exactly equal forces, the empty box
has the greater acceleration. In general, the greater theunbalanced force and the smaller the mass, the greater
the acceleration.
Third Law
Newton’s third law states: for every action there isan equal and opposite reaction; the two actions are

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directed along the same straight line. In this statement
the term "action" means the force that one body exerts
on a second, and "reaction" means the force that thesecond body exerts on the first. That is, if body A exerts
a force on body B, then B must exert an equal andopposite force on A. Note that action and reaction,
though equal in magnitude and opposite in direction,never neutralize or cancel each other for they always act
on different objects.
The recoil of a rifle demonstrates this law of actionreaction. The gunpowder in a charge is ignited by theignition cap, combustion takes place, and the bullet is
rapidly accelerated from the rifle. As a result of thisaction, the rifle is accelerated rearward against theshoulder of the person firing it. The recoil felt by the
person is the reaction to the action which ejected thebullet.
The principle of jet propulsion can be illustrated bya toy balloon (Figure 2-8). When inflated with the stem
sealed, pressure is exerted equally on all internal surfaces.
Since the force of this internal pressure is balanced, there
is no tendency for the balloon to move. If the stem isreleased, the balloon will move in a direction away fromthe escaping jet of air. Although the flight of the balloon
may appear erratic, it is moving in a direction away fromthe open stem at all times.
The balloon moves because of an unbalanced condition existing in it. The jet of air does not have to pushagainst the outside atmosphere it could function betterin a vacuum. When the stem area of the balloon isreleased, a convergent nozzle is created. As the airflows
through this area, velocity is increased accompanied by adecrease in air pressure. In addition, an area of theballoon’s skin against which the internal forces had beenpushing removed. On the opposite internal surface ofthe balloon, an equal area of skin still remains. The higher
internal pressure acting on this area moves the balloon ina direction away from the open stem. The flight of theballoon will be of short duration, though, due to theamount of air contained within it. If a source of pressurized air were provided, it would be possible to sustain
the balloon’s flight.
THE BRAYTON CYCLE
The "Brayton cycle" (Figure 2-9) is the name given
to the thermodynamic cycle of a gas turbine engine to
produce thrust. This is a varying-volume, constantpressure cycle of events and is commonly called the"constant-pressure cycle" or "open cycle." A more
recent term is "continuous-combustion cycle."

2-4

The four continuous and constant events are: intake,compression, expansion (includes power), and exhaust.
These cycles will be discussed as they apply to a gas
turbine engine.
In the intake cycle air enters at ambient pressure andin a constant volume and leaves at an increased pressureand decreased volume. In the compressor section air isFM 1-506

received from the intake at increased pressure slightly
above ambient and slightly decreased volume. Air enters
the compressor where it is compressed, leaving with alarge increase in pressure and decrease in volume. Thisis caused by the mechanical action of the compressor.
The next step (expansion) takes place in the combustion
chamber by burning fuel, which expands the air. Pressure
remains relatively constant, but a marked increase involume takes place. The expanding gases move rearward
through the turbine assembly and are converted fromvelocity energy to mechanical energy by the turbine.
BERNOULLI’S THEOREM
Bernoulli’s theorem states: when a gas or fluid isflowing through a convergent duet (as in nozzle stator
vanes or venturi), its speed will increase and its temperature and pressure will decrease (Figure 2-10). If this area
is a divergent duct (as in a diffuser or rotor blade), its
speed will slow, and its temperature and pressure will
increase. The total energy in a flowing gas is made up of
static and dynamic temperatures and pressures. A nozzle or a diffuser does not change the total energy level but
rather changes one form of energy to another. By varying
the area of a pipe, velocity can be changed into pressure
and pressure into velocity. A turbine engine is just sucha pipe, with areas where air pressure and velocity are
constantly being changed to achieve desired results.
To state this principle simply: the convergent duct
increases velocity and decreases pressure. The divergent
duet can be associated with the compressor where the air
is slowing and pressurizing. In the combustion area theopposite is true. There, the velocity is increasing and thepressure is decreasing.

BOYLE’S LAWThis law states that if the temperature of a confinedgas is not changed, the pressure will increase in direct
relationship to a decrease in volume. The opposite is also

true – the pressure will decrease as the volume is increased. A simple demonstration of how this works may
be made with a toy balloon. If you squeeze the balloon,
its volume is reduced, and the pressure of air inside theballoon is increased. If you squeeze hard enough, thepressure will burst the balloon.
CHARLES’ LAWThis law states that if a gas under constant pressure
is so confined that it may expand, an increase in thetemperature will cause an increase in volume. If you holdthe inflated balloon over a stove, the increase in temperature will cause the air to expand and, if the heat is
sufficiently great, the balloon will burst. Thus, the heat of
combustion expands the air available within the combustion chamber of a gas turbine engine.
PRESSURE AND VELOCITY
Air is normally thought of in relation to its temperature, pressure, and volume. Within a gas turbine engine
the air is put into motion and another factor must be
considered, velocity. Consider a constant airflow
through a duct. As long as the duct cross-sectional area
remain unchanged, air will continue to flow at the samerate (disregard frictional loss). If the cross-sectional areaof the duet should become smaller (convergent area), theairflow must increase velocity if it is to continue to flow
the same number of pounds per second of airflow(Bernoulli’s Principle). In order to obtain the necessary
velocity energy to accomplish this, the air must give up
some pressure and temperature energy (law of conservation of energy). The net result of flow through thisrestriction would be a decrease in pressure and temperature and an increase in velocity. The opposite would betrue if air were to flow from a smaller into a larger duct
(divergent area); velocity would then decrease, and pressure and temperature would increase. The
throat of an automobile carburetor is a good example ofthe effect of airflow through a restriction (venturi); evenon the hottest day the center portion of the carburetor
feels cool. Convergent and divergent areas are usedthroughout a gas turbine engine to control pressure andvelocity of the air-gas stream as it flows through theengine.
GAS TURBINE ENGINE VS RECIPROCATINGENGINEA cycle is a process that begins with certain conditions and ends with those same conditions. Reciprocating and turbine engines have similiarties. Both power
plants are air-breathing engines. Both engines have thesame series of events (intake, compression, power, and2-5

FM 1-506

exhaust). The difference is that in a turbine engine all ofthese events happen simultaneously, whereas in thereciprocating engine each event must follow the preceding event. Another difference is that in a turbine engine
each operation in the cycle is performed by a separate
component designed for its particular function; in thereciprocating engine all of the functions are performed
in one component.
Note that the reciprocating engine obtains its work
output by employing very high pressures (as much as 1000
pounds per square inch [psi]) in the cylinder during
combustion. With these high pressures a greater amountof work can be obtained from a given quantity of fuel,thereby raising the thermal efficiency of this type engine.
On the other hand, a turbine engine’s thermal efficiency

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is limited by the ability of its compressor to build up highpressures without excesive temperature rises. Ideally, aturbine engine should burn as much fuel as possible inorder to raise the gas temperature and increase the useful
output.
Energy added in the form of fuel is more thanenough to drive the compressor. The remaining energy
produces the thrust or power for useful work.
NOTE: Thermal efficiency is defined as
the relationship between the potential heat
energy in the fuel and the actual energy output of the engine.

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PART TWO. GAS TURBINE ENGINESCHAPTER 3

PRINCIPLES, OPERATION, AND CONSTRUCTION

A gas turbine engine is an air-dependent, thermal
jet propulsion device that uses exhaust-gas-driven turbine
wheels to drive a compressor, making continuous operation of the engine possible.
PRINCIPLES
Although the gas turbine engine differs radically inconstruction from the conventional four-stroke, fiveevent cycle reciprocating engine, both operate on thesame principle. In the piston (reciprocating) engine, thefunctions of intake, compression, combustion, and exhaust all take place in the same combustion chamber
therefore, each must completely occupy the chamber
during its respective part of the combustion cycle. In thegas turbine engine, a separate section is devoted to eachfunction; all functions are performed at the same time
without interruption. These sections are the air-inlet
section the compressor section, the combustion section,
and a turbine and exhaust section. The compressor
brings in compresses, and forces air into the combustion
section. Fuel is then injected into the combustion area,
where it mixes with the compressed air. The fuel and air
mixture is ignited by an electrical ignition plug; however,once started, ignition is shut off and combustion is continuous and self-supporting. In gas turbine engines
designed for Army use, nearly two-thirds of the energy
available from the products of combustion is necessary
to drive the compressor. The power turbine is designed
to extract as much of the remaining energy as possible
from the gases and transform it into shaft horsepower
which is used to drive an output shaft or propeller froma power shaft.
A basic gas turbine engine is illustrated in Figure 3-1.
A container of compressed air has been provided with an
opening (nozzle) through which the air passes off into theatmosphere until the pressure is exhausted (A). This
brief high-speed airflow (velocity) through the node
produces a mild momentary reaction force (solid arrow).

The density of the air can be increased by heating (B).
The result of this increased volume is a greater velocity
as indicated by the long arrow (heated air) as comparedto the short broken arrow. In (C) the heated air (solid
arrow) is directed (impinged) onto the airfoil-shaped
blades attached to the rim of the disc (turbine) causing
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the turbine to rotate on its shaft. To provide for thecontinuous airflow, an opening is placed in the forward
end of the container(D). A mechanical pump (compressor) forces air (solid arrow) into the container (E).With
high-velocity air driving the turbin, the rotating motionof the turbine can be used to drive the compressor by
joining the two rotating parts with a shaft (F). Up to thispoint the air container has only provided for the heatingof the air by an outside means. The container becomesan engine combustion chamber when fuel is introduced
mixed with the air, and ignited internally. Not all of the
air entering the combustion chamber is needed for combustion. However, all of it is heated and expands rapidly.
It exhausts as hot, high-velocity gas through the nozzle
and is directed (impinged) onto the turbine. As the gases
pass across the turbine, approximately two-thirds of theavailable energy is used The remaining energy still possesses a high work potential. Therefore, a second turbine
is placed in the flow path of the hot, high-velocity gases
and is driven in the same manner as the comprcssordriven turbine. The second turbine (power turbine), bymeans of a shaft system mechanically independent of thecompressor turbine system drives a system of reduction
gears and the output shaft. The output shaft is externallysplined for turboprop application or internallys splined forhelicopter application (G).
ENGINE OPERATION
The front or inlet duct is almost entirely open topermit outside air to enter the front of the engine. The
compressor works on the incoming air and delivers it to
the combustion section with as much as twelve times ormore the pressure the air had when it entered the engine.
In the combustion section fuel is sprayed and mixed withthe compressed air. The air-fuel mixture is then ignitedby igniter or spark plugs. Once the mixture is lighted theigniter is turned off as the burning process will continue
without further assistance, providing the engine is supplied with the right air-fuel mixture. This air-fuel mixturebums at a relatively constant pressure with only about 25
percent of the air taking part in the actual combustion
process The rest of the air is used for cooling some of
the components of the combustion section, This heatedexpanding air is then directed to the turbine. The turbines are connected to the compressor and accessories
They will extract a major portion (nearly two-thirds) of
the energy to drive them. The remaining energy will be
used to either propel the aircraft through a jet nozzle or
to turn other times (power turbines). Power turbines
will take this energy and transform it into shaft horsepower, which in turn drives an output shaft or propeller
from a power shaft.
3-2

TYPES OF GAS TURBINE ENGINESThe gas turbine engines most commonly used today
are divided into four types:
Turbojet engine – a gas turbine engine in which
the net energy available is used by the air or hot
gas solely in the form of the jet issuing through apropelling nozzle or nozzles or a jet engine
whose air is supplied by aturbine-driven compressor, the turbine being driven by exhaustgases.
Turboshaft engine– an engine in which the net
energy available is transmitted from the exhaustgas-driven turbine wheel through a planetary
gear, which may be integral to the engine orexternally provided by the aircraft manufacturer, to turn a helicopter transmissiondriven powertrain system.
Turboprop engine – a turboshaft engine inwhich power is transferred from the power shaft
through reduction gearing to a propeller. The
major difference between the two engines is thegear reduction.
Turbofan engine – in principle, the same as aturboprop, except that the geared propeller isreplaced by a duct-enclosed axial-flow fan
driven at engine speed (Figure 3-2).

FM 1-506

The term "turbo" means "turbine." Therefore, atruboshaft engine is an engine that delivers power
through a shaft, which, in turn is powered by a turbine
wheel. Army aircraft gas turbine engines are generally
the free-type power turbine turboprop or turboshaft
engines. Having a free power turbine enables the power
output shaft to turn at a constant speed while the powerproducing capability of the engine can be varied to accommodate the increased loads applied to the power
output shaft. Turbine engines may also be classified into
two general groups, centrifugal-flow and axial-flow,
depending on the type of compressor used. However,most gas turbine engines in Army aircraft employ a
combination of both types.
AIR INLET SECTION
The amount of intake required by a gas turbine engine
is approximately 10 times that required by a reciprocating
engine. The air entrance is designed to conduct incoming
air to the compressor with minimum energy loss resulting
from drag or ram pressure loss, that is, the flow of air into
the compressor should be free of turbulence to achieve
maximum operating efficieney. Proper design contributes materially to aircraft performance by increasing

the ratio of compressor discharge pressure to duct inlet
pressure.
The amount of air passing through the engine
depends on the–Speed of the compressor RPM.
Forward speed of the aircraft.
Density of the ambient air.
Inlets (Figures 3-3, 3-4) are classified as–Nose inlets located in the nose of the fuselage,
power lant Pod, or nacelle.
Winq inlets– located along the leading edge of
the wing usually at the root for single-engine
installations.
Annular inlets – encircling in whole or in partthe fuselage or power plant pod or nacelle.
Scoop inlets -projecting beyond the immediate
surface of the fuselage or nacelle.
Flush inlets – recessed in the side of thefuselage, power plant or or nacelle.
Bellmouth inlets– a bell-shaped funnel withcarefully rounded Shoulders, mounted to frontof engine.

There are two basic types of air entrances in use:
single entrance and divided entrance. Generally, it is best
to use a single entrance with an axial-flowengine to obtainmaximum ram pressure through straight flow. Single
entrance is used almost exclusively on wing or external
installations where the unobstructed entrance lends itself
readily to a single short, straight duct (Figure 3-5).

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A divided entrance offers greater opportunity toditluse the incoming air and enter the plenum chamberwith the low velocity required to utilize efficiently adouble-entry compressor (refer back to Figure 3-3).
(The plenum chamber is a storage place for ram air,usually associated with fuselage installations.) It is also
advantageous when the equipment installation or pilot
location makes the use of a single or straight duct impractical. Inmost cases the divided entrance permits the useof very short ducts with a resultant small pressure drop
through skin fiction.
The air inlet section of turboprop and turboshaft
engines (Figure 3-6) also incorporates some type of particle separator or inlet screens to protect compressors
from foreign object damage (FOD). Systems will vary
among manufacturers. Consult the aircraft or engine
technical manuals for a descriptionof the inlet duct and.
its particular air inlet protective device.

depends on the speed of the engine. On most engines thevanes are hollow to allow hot air or oil to flow through toprevent ice buildup.
COMPRESSOR SECTION
The compressor section of the turbine engine has
many functions. Its primary function is to supply enoughair to satisfy the requirements of the combustion burners.
The compressor must increase the pressure of the massof air received from the air inlet duct and then discharge
it to the burners in the required quantity and pressure.
A secondary function of the compressor is to supply
bleed air for various purposes in the engine and aircraft.
The bleed air is taken nom any of the various pressure
stages of the compressor. The exact location of the bleed
port depends on the pressure or temperature required
for a particular job. The ports are small openings in thecompressor case adjacent to the particular stage fromwhich the air is to be bled. Varying degrees of pressureand heat are available simply by tapping into the appropriate stage. Air often bled from the final or highest
pressure stage because at this point pressure and airtemperature are at a maximum. At times it may benecessary to cool this high-pressure air. If it is used forcabin pressurization or other purposes where excess heatwould be uncomfortable or detrimental the air is sent
through a refrigeration unit.
Bleed air has various uses including driving the
remote-driven accessories. Some current applications of
bleed air are–
In cabin pressurization heating and cooling.
In deicing and anti-icing equipment.
For pneumatic starting of engines.
In auxiliary drive units (ADUs).
In control booster servo systems.
As power for running instruments.

Air inlet ducts have an anti-icing system. Turbine
engine air inlets use hot engine oil, hot bleed air, or a
combination of both These systems prevent icing in aturbine engine air inlet. They are not designed to melt icethat has already formed on or in the inlets.
Guide vanes are included in some turbine engines
These vanes direct air coming through the inlet into thecompressor at the most efficient angle. The angle

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Compressor section location depends on the type ofcompressor. In the centrifugal-flow engine the compressor is between the accessory section and the combustion
section; in the axial-flow engine the compressor is between the air inlet duct and the combustion section.
Centrifugal-Flow Compressor
The centrifugal-flow compressor basically consistsof an impeller (rotor), a diffuser (stator), and a compressor manifold. The impeller and the diffuser are the twomain functional elements. Although the diffuser is aseparate component positioned inside and secured to theFM 1-506

manifold, the entire assembly (diffuser and manifold) isoften referred to as the diffuser.
The impeller’s function is to pick up and accelerate
air outward to the diffuser. Impellers may be either of
two types – single entry or double entry. Both are similar
in construction to the reciprocating engine supercharger
impeller. The double-entry type is similar to two back-to
back impellers. However, because of much greater combustion air requirements in turbine engines, these
impellers are larger than supercharger impellers.
The principal differences between the two types ofimpellers are size and ducting arrangement. The doubleentry type has a smaller diameter but is usually operated
at a higher rotational speed to ensure enough airflow.
The single-entry impeller permits convenient ducting
directly to the impeller eye (inducer vanes) as opposedto the more complicated ducting necessary to reach therear side of the double-entry type. Although slightly
more efficient in receiving air, the single-entry impeller
must be large in diameter to deliver the same quantity of
air as the double-entry type. This of course, increases theoverall diameter of the engine.
Included in the ducting for double-entry compressor
engines is the plenum chamber. This chamber is necessary for a double-entry compressor because air mustenter the engine at almost right angles to the engine axis.
To give a positive flow, air must surround the engine
compressor at a positive pressure before entering thecompressor.
Multistage centrifugal compressors consist of two or
more single compressors mounted in tandem on the same
shaft. The air compressed in the first stage passes to thesecond stage at its point of entry near the hub. This stage
will further compress the air and pass it to the next stage
if there is one. The problem with this type of compression
is in turning the air as it is passed from one stage to thenext.
The diffuser is an annular chamber provide with anumber of vanes forming a series of divergent passages
into the manifold. The diffuser vanes direct the flow of
air from the impeller to the manifold at an angle designed
to retain the maximum amount of energy imparted by the
impeller. They also deliver the air to the manifold at avelocity and pressure satisfactory for combustion chambers.
The compressor manifold diverts the flow of air fromthe which, which is an integral part of the manifold, into
the combustion chambers. The manifold will have one
outlet port for each chamber so that the air is evenly

divided. A compressor outlet elbow is bolted to each of
the outlet ports. These air outlets are constructed in theform of ducts and are known by a variety of names
including "air outlet ducts", "outlet elbows," and "combustion chamber inlet ducts.” These outlel ducts perform avery important part of the diffusion process. Theychange the airflow direction from radial to axial. The
diffusion process is completed after the turn. To help theelbows perform this function efficiently, turning vanes(cascade vanes) are sometimes fitted inside the elbows.
The vanes reduce air pressure losses by presenting asmooth, turning surface.
The centrifugal compressor is used best on smaller
engines where simplicity, flexibility, and ruggedness areprimary requirements. These have a small frontal area
and can handle high airflows and pressures with low lossof efficiency.
Centrifugal-flow compressors have the following advantages:
High pressure rise per stage.
Efficiency over wide rotational speed range.
Simplicity of manufacture with resulting low
cost.
Low weight.
Low starting power requirements.
They have the following disadvantages:
Large frontal area for given airflow.
Impracticality if more than two stages becauseof losses in turns between stages.
Axial-Flow Compressor
Axial-flow compressors have have main elements arotor (drum or disc type) and a stator. These compressors are constructed of several different materials
depending on the load and operating temperature. The
drum-type rotor consists of rings that are flanged to fit
one against the other so that the entire assembly can beheld together by through bolts. This type of construction
is satisfactory for low-speed compressors where
centrifugal stresses are low (Figure-3-7). The rotor (disctype) assembly consists of—
Stub shafts.
Discs.
Blades.
Ducts.
Air vortex spoilers.
Spacers.
Tie bolts.
Torque cones.
3-5

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Rotor blades are generally machined from stainlesssteel forgings, although some maybe made of titanium inthe forward (colder) section of the compressor (Figure3-8). The blades are attached in the disc rim by different

3-6methods using either the fir-tree-type, dovetail-type, orbulb-type root designs. The blades are then locked into
place with screws, peening, locking wires, pins, keys, orplates (Figure 3-9). The blades do not have to fit too

FM 1-506

tightly in the disc because centrifugal force during engine
operation causes them to seat. Allowing the blades somemovement reduces the vibrational stresses produced byhigh-velocity airstreams between the blades. The newest
advance in technology is a one-piece design machined
blade disc (combined disc and blade);both disc and rotor
blade are forged and then machined into one (refer toFigure 3-8 again).
Clearances between rotor blades and the outer caseare very important to maintain high efficiency. Becauseof this, some manufacturers use a “wear fit” design between the blade and outer case. Some companies design
blades with knife-edge tips that wear away to form theirown clearances as they expand from the heat generated
by air compression. Other companies coat the innersurface of the compressor case with a soft material(Teflon) that can be worn away without damaging theblade. Rotor discs that are joined together by tie bolts
use serration splines or curve coupling teeth to preventthe discs from turning in relation to each other. Anothermethod of joining rotor discs is at their rims.
Axial-flow compressor casings not only support
stator vanes and provide the outer wall of the axial paths

the air follows but also provide the means for extracting
compressor air for various purposes. The stator andcompressor cases show great differences in design andconstruction. Some compressor cases have variable
stator vanes as an additional feature. Others (compressor cases) have fixed stators. Stator vanes may be either
solid or hollow and mayor may not be connected at theirtips by a shroud. The shroud serves two purposes. First,
it provides support for the longer stator vanes located in
the forward stages of the compressor second, it provides
the absolutely necessary air seal between rotating andstationary parts. Some manufacturers use split compressor cases while others favor a weldment, which forms a
continuous case. The advantage of the split case is thatthe compressor and stator blades are readily available forinspection or maintenance. On the other hand the continuous case offers simplicity and strength since it requires no vertical or horizontal parting surface.
Both the case and the rotor are very highly stressedparts. Since the compressor turns at very high speeds thediscs must be able to withstand very high certrifugal
forces. In addition the blades must resist bending loads
and high temperatures. When the compressor is constructed each stage is balanced as a unit. The compressor case in most instances is one of the principal
structural, load-bearing members of the engine. It may
be constructed of aluminum steel, or magnesium.
Axial-flow compressors have the following advantages:
High peak efficiency.
Small frontal area forgiven airflow.
Straight-through flow, allowing high ram efficiency.
Increased pressure rise due to increased number of stages with negligible losses.
They have the following disadvantages:
Good efficiency over narrow rotational speed
range.
Difficulty of manufacture and high cost.
Relatively high weight.
High starting power requirements (this has
been partially overcome by split compressors).
The air in an axial compressor flows in an axial
direction through a series of rotating (rotor) blades andstationary (stator) vanes that are concentric with the axisof rotation. Unlike a turbine, which also employs rotor
blades and stator vanes the flow path of an axial compressor decreases in cross-sectional area in the directionof flow. This reduces the volume of air as compression
progresses from stage to stage.
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FM 1-506

After being delivered to the face of the compressor
by the air inlet duct incoming air passes through the inlet
guide vanes. Upon entering the first set of rotating
blades, the air, which is flowing in a general axial direction is deflected in the direction of rotation. The air isarrested and turned as it is passed on to a set of stator
vanes. Following that it is picked up by another set ofrotating blades and soon through the compressor. Air
pressure increases each time it proses through a set ofrotors and stators.
The rotor blades increase the air velocity. When air
velocity increases, the ram presure of air passing througha rotor stage also increases. This increase in velocity andpressure is somewhat but not entirely mullified by diffusion. When air is forced past the thick sections of therotor blades static pressure also increases. The larger
area at the rear of the blades (due to its airfoil shape) acts
as a diffuser.
In the stators velocity decreases while staticpressure increases. As air velocity decreases, the pressure due to velocity or ram that has just been gained in
Passing through preceding rotor stage decreases
somewhat; however, the total pressure is the sum of staticpressure and pressure due to ram. Successive increases
and decreases in velocity as air leaves the compressor
are usually only slightly greater than the velocity of the
air at the entrance to the compressor. As the pressureis built up by successive sets of rotors and stators, lessand less volume is required. Thus, the volume within
the compressor is gradually decreased. At the exit of

3-8

the compressor, a diffuser section adds the final stage tothe compression process by again decreasing velocity andincreasing static pressure just before the air enters theengine burner section.
Normally, the temperature change caused by diffusion is not significant by itself. The temperature rise
which causes air to get hotter and hotter as it continues
through the compressor, is the result of the work being
done on the air by the compressor rotors. Heating of the
air occurs because of the compression process and because some of the mechanical energy of the rotor isconverted to heat energy.
Because airflow in an axial compressor is generally
diffusing it is very unstable. High efficiency is maintained
only at very small rates of diffusion. Compared to aturbine, quite a number of compressor stages are needed
to keep the diffusion rate small through each individual
stage. Also, the permissible turning angles of the blades
are considerably smaller than those which can be used inturbines. These are the reasons why an axial compressor
must have many more stages than the turbine which
drives it. In addition, more blades and consequently
more stages are needed because the compressor, incontrast to a turbine, is endeavoring to push air in a
direction that it does not want to go in.
Dual Compressor
The dual compressor is a combination either of two
axial compressors or of an axial and a centrifugal compressor (Figure 3-10). The dual-axial compressor consists of

FM 1-506

a low-pressure compressor in front and a high-pressurecompressor in the rear. Both compressors(lowand high)are driven by two different shafts that connect to different
turbines. The starter is usually connected to the high-pressure compressor because it reduces the torque required to start the engine. With the rear (high-pressure)compressor turning at governed Speed, the front (lowpressure) compressor (not governed) is automaticallyrotated by its turbine. Rotation speed is whatever speed
will ensure an optimum flow of air through the compressor. With the front and rear compressor rotors working
in harmony instead of interfering with each other, compression rates can be increased without decreasing efficiency. Due to the added length of the engine this typeof compressor is found on turbojet aircraft.
Most gas turbines in Army aircraft have a combination of an axial compressor (front) and a centrifugal
compressor (rear). The usual combination is a five- or
seven-stage axial- flow compressor and a centrifugal-flow
compressor. The axial compressor and centrifugal compressor combination is mounted on the same shaft; thecompressors turn in the same direction and at the same
speed. By combining the the manufacturer makes themost of the advantages of both compressors small frontaI area, increased compression ratios, and shortened
overall engine length. Using the centrifugal-flow compressor boosts compression and increases efficiency of
the turbine engine. The centrifugal compressor also
shortens the length of the engine. If the centrifugal compressor were not added, the manufacturer would have to
add more stages of axial compression to equal that of thecentrifugal compressor.
COMBUSTION SECTION
The combustion section contains the combustion
chambers, igniter plugs, and fuel nozzle or fuel injectors.
It is designed to burn a fuel-air mixture and to deliver
combusted gases to the turbine at a temperature notexceeding the allowable limit at the turbine inlet.
Theoretically, the compressor delivers 100 percent of its
air by volume to the combustion chamber. However, thefuel-air mixture has a ratio of l5 parts air to l part fuel byweight. Approximately 25 percent of this air is used toattain the desired fuel-air ratio. The remaining 75 percent is used to form an air blanket around the burning
gases and to dilute the temperature, which may reach as
high as 3500ºF, by approximately one-half. This ensuresthat the turbine section will not be destroyed by excessiveheat.

The air used for burning is known as primary air; thatused for cording is secondary air. Secondary air is controlled and directed by holes and louvers in the combustion chamber liner. Igniter plugs function during starting
only; they are shut off manually or automatically. Combustion is continuous and self-supporting. After engine
shutdown or failure to start, a pressure-actuated valve
automatically drains any remaining unburned fuel fromthe combustion chamber. The most common type usedin Army gas turbine engines is the external annularreverse-flow type.
The primary function of the combustion section is,of course, to bum the fuel-air mixture, thereby adding
heat energy to the air. To do this efficiently, the combustion chamber must–
Provide the means for mixing the fuel and air toensure good combustion.
Bum this mixture efficiently.
Cool the hot combustion products to a temperature which the turbine blades can withstandunder operating conditions.
Deliver the hot gases to the turbine section.
The location of the combustion section is directly
between the compressor and turbine sections. The combustion chambers are always arranged coaxially with thecompressor and turbine, regardless of type, since thechambers must be in a through-flow position to function
efficiently.
All combustion chambers contain the same basic
elements:
A casing
A perforated inner liner.
A fuel injection system.
Some means for initial ignition.
A fuel drainage system to drain off unburned
fuel after engine shutdown.
There are currently three basic types of combustion
chambers, varying in detail only:
The multiple-chamber or can type.
The annular or basket type.
The can-annular type.
Can-Type Combustion Chamber
The can-type combustion chamber is typical of thetype used on both centrifugal and axial-flow engines(Figure 3-11). It is particularly well suited for thecentrifugal compressor engine since the air leaving thecompressor is already divided into equal portions as it
3-9

FM 1-506

leaves the diffuser vanes. It is then a simple matter toduct the air from the diffuser into the respective combustion chambers arranged radially around the axis of
the engine. The number of chambers will vary; in thepast as few as 2 and as many as 16 chamber’s havebeen used. The present trend is about 8 or 10 combustion chambers. Figure 3-11 illustrates the arrangementof can-type combustion chambers. On American-built

3-10

engines these chambers are numbered in a clockwise
direction facing the rear of the engine with the No. 1
chamber at the top.
Each can-type combustion chamber consists of anouter case or housing with a perforated stainless steel(highly heat-resistant) combustion chamber liner or inner
liner (Figure 3-12). The outer case is divided for ease of

FM 1-506

liner replacement. The larger section or chamber bodyencases the liner at the exit end; the Smaller chamber coverencases the front or inlet end of the liner.
The interconnector (flame propagation) tubes area
necessary part of can-type combustion chambers. Since
each can is a separate burner operating independently ofthe others, there must be some way to spread combustion
during the initial starting operation. This is done byinterconnecting all the chambers. The flame is started by
the spark igniter plugs in two of the lower chambers; itpasses through the tubes and ignites the combustible
mixture in the adjacent chamber. This continues until allchambers are burning. The flame tubes will vary inconstruction details from one engine to another although
the basic components are almost identical.
The interconnector tubes are shown in Figure 3-12.
Bear in mind that not only must the chambers be interconnected by an outer tube (in this case, a ferrule), butthere must also be a slightly longer tube inside the outer
one to interconnect the chamber liners where the flame
is located The outer tubes or jackets around the interconnecting flame tubes not only afford airflow between
the chambers but also fulfill an insulating function aroundthe hot flame tubes.
The spark igniters are normally two in number.
They are located in two of the can-type combustion
chambers.
Another very important requirement in the construction of combustion chambers is providing the meansfor draining unburned fuel. This drainage prevents gum
deposits in the fuel manifold, nozzles, and combustion
chambers. These deposits are caused by the residue leftwhen fuel evaporates. If fuel is allowed to accumulate
after shutdown there is the danger of afterfire.
If the fuel is not drained, a great possibility existsthat at the next starting attempt excess fuel in the combustion chamber will ignite and tailpipe temperature will
go beyond safe operating limits.
The liners of can-type combustors have perforationsof various sizes and shapes, each hole having a specificpurpose and effect on flame propagation in the liner. Air
entering the combustion chamber is divided by holes,
louvers, and slots into two main streams – primary andsecondary air. Primary (combustion) air is directed inside the liner at the front end where it mixes with the fueland bums. Secondary (cooling) air passes between theouter casing and the liner and joins the combustion gases
through larger holes toward the rear of the liner, coolingthe combustion gases from about 3500ºF to near 1500ºF.

Holes around the fuel nozzle in the dome or inlet end of
the can-type combuster liner aid in atomization of thefuel. Louvers are also provided along the axial length ofthe liners to direct a cooling layer of air along the insidewall of the liner. This layer of air also tends to control the
flame pattern by keeping it centered in the liner, preventing burning of the liner walls.
Annular- or Basket-Type Combustion Chamber
Figure 3-13 illustrates the flow of air through thelouvers in the double-annular combustion chamber.
Some provision is always made in the combustion
chamber case or in the compressor air outlet elbow for
installation of a fuel nozzle. The fuel nozzle delivers thefuel into the liner in a freely atomized spray. The freer thespray, the more rapid and efficient the burning process.
Two types of fuel nozzles currently being used in thevarious types of combustion chambers are the simplex
nozzle and the duplex nozzle.
The annular combustion chamber consists basicallyof a housing and a liner, as does the can type. The liner
consists of an undivided circular shroud extending all theway around the outside of the turbine shaft housing. The
chamber may be constructed of one or more baskets. If
two or more chambers are used, one is placed outside theother in the same radial plane; hence, the term "doubleanmular chamber.”
The spark igniter plugs of the annular combustion
chamber are the same basic type used in the can combustion chambers, although construction details may vary.
There are usually two plugs mounted on the boss
provided on each of the chamber housings. The plugs
must be long enough to protrude from the housing into
the outer annulus of the double-annular combustion
chamber.
The annular-type combustion chamber is used inmany engines designed to use the axial-flow compressor.
It is also-used by engines incorporating dual-type compressors (combinations of axial flow and centrifugalflow). Its use permits building an engine of small
diameter. Instead of individual combustion chambers, thecompressed air is introduced into an annular space formed
by a combustion chamber liner around the turbine shaft.
Usually, enough space is left between the outer liner walland the combustion chamber housing to permit the flow ofcooling air from the compressor. Fuel is introduced
through nozzles or injectors connected to a fuel manifold.
The nozzle opening may face upstream or downstream
to airflow depending on engine design. Various meansare provided to introduce primary (compressed) air to
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FM 1-506

the vicinity of the nozzle or injectors to support combustion and additional air dowmstream to increase the mass
flow. Secondary cooling air reduces the temperature ofgases entering the turbine to the proper level,
Some axial compressor engines have a single annular
combustion chamber similar to that shown in Figure 3-14.
The liner of this type of burner consists of continuous,
circular, inner and outer shrouds around the outside of
the compressor drive shaft housing. Holes in the shrouds
allow secondary cooling air to enter the center of the
combustion chamber. Fuel is introduced through a seriesof nozzles at the upstream end of the liner. Because of
their proximity to the flames all types of burner liners areshort-lived in comparison to other engine components;
they require more frequent inspection and replacement.
This type of burner uses the limited space available
most effectively, permitting better mixing of the fuel andair within a relatively simple structure. An optimum ratioof burner inner surface area to volume is provided; thisensures maximum cooling of the gases as combustion
occurs. The design also tends to prevent heat warping
However, the burner liner on some engines cannot bedisassembled without removing the engine from theaircraft – a distinct disadvantage.
The latest annular combustion system for military
use is a low-pressure fuel injection system with vortex air
swirlers to mix fuel and compressor discharge air before
combustion. The fuel injector is positioned into the center of an air swirler in the dome of the liner. Fuel leaving
the injectors (which has been swirled) is surrounded by a

3-12

concentric air vortex pattern. This breaks fuel particles
down to an extremely small size before they reach the
combustion zone, This creates excellent fuel-air mixingthat ensures a low smoke level in the exhaust. The lowpressure fuel system does not have tine nozzle orifices
and can handle contaminated fuel without clogging.

FM 1-506

Can-Annular-Type Combustion Chamber
The can-annular-type combustion chamber wasdeveloped by Pratt and Whitney for use in their JT3
axial-flow turbojet engine. Since this engine features the
split-spool compressor, it needed a combustion chambercapable of meeting the stringent requirements of maximum strength and limited length plus high overall efficiency. These were necessary because of the high airpressures and velocities in a split-spoolcompresmr alongwith the shaft length limitations explained below.
The split compressor requires two concentric shafts
to join the turbine stages to their respective compressors.
The front compressor joined to the rear turbine stages
requires the longer shaft. Because this shaft is inside theother, a limitation is imposed on diameter. The distance
between the front compressor and the rear turbine mustbe limited if critical shaft lengths are to be avoided.
Since the compressor and turbine are not susceptible to appreciable shortening the necessary shaft length
limitation had to be absorbed by developing a new typeof burner. A design was needed that would give thedesired performance in much less relative distance than
had previously been assigned.
Can-annular combustion chambers are arranged
radially around the axis of the engine in this instance therotor shaft housing. The combustion chambers are
enclosed in a removable steel shroud that covers theentire burner section. This feature makes the burners
readily available for any required maintenance.
The burners are interconnected by projecting flame
tubes. These tubes make the engine-starting processeasier. They function identically with those previously
discussed but differ in construction details.
Each combustion chamber contains a central bulletshaped perforated liner. The size and shape of the holes
are designed to admit the correct quantity of air at thecorrect velocity and angle. Cutouts are provided in twoof the bottom chambers for installation of the spark
igniters. The combustion chambers are supported at theaft end by outlet duct clamps. These clamps secure themto the turbine nozzle assembly.
The forward face of each chamber presents six apertures which align with the six fuel nozzles of the corresponding fuel nozzle duster. These nozzles are thedual-orifice (duplex) type. They require a flow divider(pressurizing valve) as was mentioned above in the cantype combustion chamber discussion. Around each nozzle are preswirl vanes for imparting a swirling motion to

the fuel spray. This results in better atomization burningand efficiency.
Swirl vanes perform two important functions. They
cause—High flame speed-provides better nixing of
air and fuel and ensures spontaneous burning.
Low air velocity axially-swirling prevents theflame from moving axially too rapidly.
Swirl vanes greatly aid flame propagation because ahigh degree of turbulence in the early combustion andcooling stagea is desirable. Virous mechanical mixingof fuel vapor with primary air is necessary; mixing bydiffusion alone is too slow. Mechanical mixing is alsodone by other measure; for example,placingcoarse screensin the diffuser outlet as is done in most axial-flow engines.
Can-annular combustion chambers must also havefuel drain valves in two or more of the bottom chambers.
This ensures drainage of residual fuel to prevent its being
burned at the next start.
The flow of air through the holes and louvers of the
can-annular chambers is almost identical with the flow
through other types of burners. Special baffling is used
to swirl the combustion airflow and to give it turbulence.
Performance Requirements
Performance requirements include–
High combustion efficiency. This is necessary
for long range.
Stable operation. Combustion must be freefrom blowout at airflows ranging from idle tomaximum power and at pressures representing
the aircraft’s entire altitude range.
Low pressure loss. It is desirable to have as
much pressure as possible available in the exhaust nozzle to accelerate the gases rearward
High pressure losses will reduce thrust and increase specific fuel consumption.
Uniform temperature distribution. The
average temperature of gases entering the turbine should be as close as possible to thetemperature limit of the burner material to obtain maximum engine performance. High
local temperatures or hot spots in the gas stream
will reduce the allowable average turbine inlet
temperature to protect the turbine. This will
result in a decrease in total gas energy and acorresponding decreasem engine performance.
Easy starting. Lowpressures and high velocities
in the burner make starting difficult. A poorly
designed burner will start only within a small
3-13

FM 1-506

range of light speeds and altitudes, whereas awell designed burner will permit easier air
restarts.
Small size. A large burner requires a large engine housing with a corresponding increase in
the airplane’s frontal area and aerodynamic
drag. This will result in a decrease in maximum
flight speed Excessive burner size also results inhigh engine weight, lower fuel capacity andpayload, and shorter range. Modem burners
release 500 to 1000 times the heat of a domestic
oil burner or heavy industrial furnace of equal
unit volume. Without this high heat release theaircraft gas turbine could not have been madepractical.
Low-smoke burner. Smoke not only annoys
people on the ground, it may also allow easy
tracking of high-flying military aircraft.
Low carbon formation. Carbon deposits canblock critical air passages and disrupt airflow
along the liner walls, causing high metal
temperatures and low burner life.
All of the burner requirements must be satisfied overa wide range of operating conditions. For example,
airflows may vary as much as 50:1, fuel flows as much as
30:1, and fuel-air ratios as much as 5:1. Burner pressures
may cover a ratio of 100:1, while burner inlet temperatures may vary by more than 700ºF.
The effect of operating variables on burner performance is—
Pressure.
Inlet air temperature.
Fuel-air ratio.
Flow velocity.
TURBINE SECTION
All turbines in modern jet engines, regardless of thetype of compress used, are of of axial-flow design. They
consist of one or more stages located immediately to therear of the engine burner section. Turbines extract kinetic
energy from the expanding gases as the gases come fromthe burners. They convert this energy in to shaft horsepower
to drive the compressor and engine accessories. In a turboshaft or turboprop engine one or more turbines will also
furnish the power required to turn the engine drive or
propeller shaft. Nearly three-fourths of all of the energy
available from combustion is needed to drive the compressor or compressors in the case of a dual-compressor engine.
This includes the fan of a turbofan engine. If the engine
3-14

is a turboshaft or turboprop, the turbines are designed to
extract as much energy as possible from the gases passing
through the engine. So efficient are the turbines in suchengines that the propeller in a turboprop aircraft provides
approximately 90 percent of the propulsive force with
only 10 percent supplied by jet thrust.
The axial-flow turbine has two main elements: turbine rotors (or wheels, as they are sometimes called) andstationary vanes. The stationary part of the assembly
consists of a plane of contoured vanes, concentric withthe axis of the turbine and set at an angle to form a seriesof small nozzles. These nozzles discharge the gases onto
the blades in the turbine rotors. The stationary vaneassembly of each stage in the turbine is usually referred
to as the turbine nozzle guide vanes. The turbine nozzle
area is the most critical part of the turbine design. If thenozzle area is too large, the turbine will not operate at its
best efficiency. If the area is too small the nozzle will havea tendency to choke and lose efficiency under maximum
thrust conditions. The turbine nozzle area is defined as
the total cross-sectional area of the exhaust gas passages
at their narrowest point through the turbine nozzle. It is
calculated by measuring and adding the areas between
individual nozzle guide vanes.
There are three types of turbines: impluse reaction
and a combination of these two known as reaction-impulse.
In the impulse type there is no net change in pressure
between the rotor inlet and the rotor exit. The blade’s
relative discharge velocity will be the same as its relative
inlet velocity. The nozzle guide vanes are shaped to form
passages which increase the velocity and reduce thepressure of the escaping gases. In the reaction type, thenozzle guide vanes do little more in relation to the rotor
than alter flow direction. The decrease in pressure andincrease in velocity of gases are caused by the convergent
shape of the passage between the rotor blades. In a jetengine the turbine is usually a balanced combination ofboth types known as a reaction-impulse turbine. Its
design is intended to achieve both a small diameter anda proper match with the compressor.
Turbines may be either single or multiple stage.
When the turbine has more than one stage, stationary
vanes are inserted between each rotor wheel and therotor wheel downstream. They are also placed at theentrance and exit of the turbine unit. Each set of stationary vanes forms a nozzle-vane assembly for the turbine
wheel that follows. The exit set of vanes serves tostraighten the gas flow before passage through the jetFM 1-506

nozzle. The wheels may or may not operate independently of one another, depending on engine type andturbine power requirements.
Shaft RPM, gas flow rate, turbine inlet and outlet
temperature and pressure, turbine exhaust velocity, andrequired power output must all be considered by thedesigner of the turbine. If the engine is equipped with adual compressor, the turbine must also be dual or "split."
In this event, the forward part of the turbine (which drives
the high-pressure compressor) can be single-stage because it receives high-energy gases directly from theburner and turns at a higher RPM than the turbine forthe low-presssure compressor. By the time the gases
reach the rear part of the turbine (which drives the
low-pressure compressor), they have expanded. Considerably more blade area is needed if work or energybalance is to be maintained. To do a multistage
turbine is used for the second part of the turbine (Figure3-15).

high expansion ratio. This results in a large temperaturedrop in gases passing through the turbine and a coolturbine exhaust. If the engine is equipped with an afterburner, a cool exhaust enables more fuel to be burned in
the afterburner without exceeding the temperature limitof the construction materials used in the afterburner.
The turbine wheel is a dynamically balanced unit
consisting of super alloy blades attached to a rotating disc.
The base of the blade is usually a "fir tree" design toenable it to be firmly attached to the disc and still allow
room for expansion. In some turbines the rotating blades
are open at their outer perimeter. More commonly, theblade is shrouded at the tip. The shrouded blades forma band around the perimeter of the turbine wheel, which
serves to reduce blade vibrations. The weight of the
shrouded tips is offset because the shrouds permit thinner, more efficient blade sections than are otherwise
possible because of vibration limitations. Also, by acting
in the same manner as aircraft wing tip fences, theshrouds improve the airflow characteristics and increase
the efficiency of the turbine. The shrouds also serve to
cut down gas leakage around the tips of the turbine
blades.

Turbines are subjected to high speeds and hightemperatures. High speeds result in high centrifugal
forces. Turbines must operate close to temperature limits
that, if exceeded, lower the strength of the materials they
are constructed of. Turbine blades undergo distortion or
lengthening known as "creep." Creep means that theblade stretches or elongates. This condition is cumulative. The rate of creep is determined by the load imposedon the turbine and the strength of the blade. The strengthof the blade is determined by the temperature within theturbine. Since changes in pitch and creep are more
pronounced if engine operating limits are not respected
the pilot or flight engineer must closely observe thetemperature and RPM limits stipulated by the manufacturer.
Construction
The turbine wheel is one of the most highly stressed
engine parts. Not only must it operate at temperatures
up to approximately 1700ºF, but it must do sounder severe
centrifugal loads imposed by high rotational speeds ofover 40,000 RPM for small engines and 8000 RPM forlarger ones. Consequently, engine speed and turbine
inlet temperature must be accurately controlled to keep
the turbine within safe operating limits.
The turbine assembly is made of two main parts: thedisc and blades. This disc or wheel is a statically anddynamically balanced unit of specially alloyed steel. It
usually contains large percentages of chromium, nickel,
and cobalt. After forging, the disc is machined all overand carefully inspected using X rays, magnetism, andother inspection methods for structural integrity. The
blades or buckets are attached to the disc by means of a
fir tree design to allow for different rates of expansion
between the disc and the blade while still holding theblade firmly against centrifugal loads. The blade is kept
from moving axially either by rivets, special locking tabs
or devices, or another turbine stage.
Some turbine blades are open at the outer perimeter(Figure 3-16); in others a shroud is used. The shroud acts
to prevent blade-tip losses (gas leakage around the tipsof the turbine blade) and excessive vibration. By acting
in the same manner as aircraft wing tip fence, theshrouds improve airflow characteristics and increase turbine efficiency.
Shrouds reduce resistance to distortion under highloads, which tend to twist the blade toward low pitch. The
shrouded blade has an aerodynamic advantage; thinner
blade sections can be used and tip lines can be reduced
by using a knife-edge or labyrinth seal at this point.

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FM 1-506

Shrouding however, requires that the turbine run cooler
or at reduced RPM because of the extra mass at the tip.
On blades that are not shrouded, the tips are cut orrecessed to a knife-edge to permit a rapid “wearing-in”
of the blade tip to the turbine casing with am-responding
increase in turbine efficiency.
Blades are forged from highly alloyed steel. They
are carefully machined and inspected before being certified for use. Many engine manufacturers will stamp amoment weight number on the blade to retain rotor
balance when replacement is necessary.
Another method for increasing efficiency is the useof honeycomb shrouding (Figure 3-17). This shroud
works as a labyrinth sealing the unshrouded turbine

3-16. a stator support,
tips. These shrouds are all. . housed. bywhich, in turn, is supported by the engine outer casing.
This design is currently in use in the new General Electricturboshaft engines.
Nozzle vanes may be either case or forged. Somevanes are hollow (Figure 3-18) to allow a degree ofcooling by compressor bleed air. In all cases the nozzle
assembly is made of very high-temperature, highstrength steel to withstand the direct impact of the hothigh-pressure, high-velocity gas flowing from the combustion chamber.

FM 1-506

Some manufacturers are experimenting withtranspiration-cooled nozzle and turbine blading in which theairflows through thousands of small holes in a porous airfoil
made from a sintered wire mesh material (Figure 3-19).
The performance of the gas turbine engine depends
largely on the temperature at the inlet. Increasing this0temperaturefrom the present limit of about 1750 F to the02500 F possible with transpiration-cooled blades will
result in about a 100 percent increase in specific horsepower. Transpiration cooling may be a promisingdevelopment in gas turbine design.
Cooling
Design engineers use every device at their command
to increase the allowable inlet temperature. On practically all large engines, one such device is to cool thefret-stage turbine inlet guide vanes and the first-stage
rotor blades. This is done by conducting compressor
bleed air through passages inside the engine to the turbine area. There, the air (the coolant) is led to thelongitudinal holes, tubes or cavities in the first-stage
vanes and blades.
After entering the vane and blade passages, the air(coolant) is distributed through holes at the leading andtrailing edges of the vanes and blades. The air impinges
along the vane and blade surfaces and then passes out of

the engine with the exhaust. Although bleed air comingfrom the compressor may be hot, it is cool in relation to thetemperature at the turbine inlet. This air, therefore servesto cool the vanes and blades. This permits gases comingfrom the burner section to enter the turbine at higher
temperatures than would otherwise be permissible.

Cooling is necessary only in the turbine inlet area because
enough energy is extracted from the exhaust gases by the
first or first and second stages of the turbine to reduce thetemperature to a tolerable level.
EXHAUST SECTION
The term “exhaust duet” applies to the engine exhaust pipe or tail pipe including the jet nozzle of a
non-after-burning engine (Figure 3-20). Altough anafterburner might also be considered a type of exhaust
duct, after burning is a subject in itself.
If the engine exhaust gases could be discharged
directly to the outside air in an exact axial direction at the
turbine exit, an exhaust duct might not be necessary. This,
however, is not practical. A larger total thrust can beobtained from the engine if the gases are discharged fromthe aircraft at a higher velocity than that permissible atthe turbine outlet. An exhaust duct is added to collectand straighten the gas flow as it comes from the turbine.
It also increases the velocity of the gases before they are

3-17

FM 1-506

discharged from the exhaust nozzle at the rear of the duct.
Increasing gas velocity increases its momentum and thethrust produced.

3-18

An engine exhaust duct is often referred to as theengine tail pipe. The duct is essentially a simple, stainlesssteel, conical or cylindrical pipe. The engine tail cone andstruts (Figure 3-21) are usually included at the rear of theturbine. The struts support the rear bearing and impartan axial direction to the gas flow, the tail cone helps
smooth the flow. Immediately aft of the turbine outlet
and usually just forward of the flange to which the exhaust
duct is attached, the engine has a sensor for turbine
discharge pressure. In large engines, it is not practical tomeasure internal temperature at the turbine inlet. Therefore, the engine is usually also instrumented for exhaustgas temperature at the turbine outlet. One or more
thermocouples preinserted in the exhaust case to provide
adequate sampling of exhaust gases. Pressure probes arealso inserted in this case to measure pressure of gases
coming from the turbine. The gradually diminishingcross-seclional area of a conventional convergent type of
exhaust duct is capable of keeping the flow through theduct constant at velocities not exceeding Mach 1.0 at theexhaust nozzle.

FM 1-506

Exhaust Ducts
Turboshaft engines in helicopters do not develop
thrust using the exhaust duct. If thrust were developedby the engine exhaust gas, it would be impossible tomaintain a stationary hover; therefore, helicopters use
divergent ducts. These ducts reduce gas velocity anddissipate any thrust remaining in the exhaust gases. On
find-wing aircraft, the exhaust duct may be the convergent type, which accelerates the remaining gases to
produce thrust. This adds additional SHP to the engine
rating. Equivalent shaft horsepower (ESHP) is the combination of thrust and SHP.
Conventional Convergent Exhaust Nozzle
The rear opening of the exhaust duct is the jet nozzle,
or exhaust nozzle as it is often called The nozzle acts asan orifice, the size of which determines velocity of gases as
they emerge from the engine. In most non-after-burning
engines, this area is critical; for this reason, it is fixed at
the time of manufacture. The exhaust (jet) nozzle area
should not be altered in the field because any change in
the area will change both the engine performance and theexhaust gas temperature. Some early engines, however,were trimmed to their correct RPM or exhaust gas
temperature by altering the exhaust-nozzle area. When
this is done, small tabs that may be bent as required areprovided on the exhaust duct at the nozzle opening. Or
small, adjustable pieces called "mice" are fastened asneeded around the perimeter of the nozzle to change thearea. Occasionally, engines are equipped with variablearea nozzles which are opened or closed, usually automatically, with an increase or decrease in fuel flow. The
velocity of the gases within a convergent exhaust duct isusually held to a subsonic speed. The velocity at thenozzle approaches Mach 1.0 (the velocity at which thenozzle will choke) on turbojets and low-bypass-ratio turbofans during most operating conditions.
Convergent-Divergent Exhaust Nozzle
Whenever the pressure ratio across an exhaust nozzle is high enough to produce gas velocities which might
exceed Mach 1.0 at the engine exhaust nozzle, morethrust can be gained by using a convergent-divergent typeof nozzle (refer back to Figure 3-21). This can be done
provided the weight penalty is not so great that the benefitof the additional thrust is nullified. The advantage of aconvergent-divergent nozzle (C-D nozzle) is greatest at
high Mach numbers becausce of the resulting higher pressure ratio across the engine nozzle. If the pressure ratio
through a subsonic exhaust duct is great enough (this willbe the case when the pressure at the entrance to theexhaust duct becomes approximately twice that at theexhaust nozzle), the change in velocity through the duct
will be enough to cause sonic velocity (Mach 1.0) at thenozzle. At very high flight Mach numbers, the pressure
ratio becomes much more than 20. If a C-D nozzle isused, the velocity at the exhaust nozzle becomes correspondingly greater than Mach 1.0. This is a distinct
advantage, provided the nozzle can effectively handle
these high velocities.
When a divergent duct is employed in combination
with a conventional exhaust duct, it is called a convergentdivergent exhaust duct (Figure 3-21). In the C-D nozzle,
the convergent section is designed to handle the gases
while they remain subsonic and to deliver them to thethroat of the nozzle just as they attain sonic velocity. The
divergent section handles the gases after they emerge
from the throat and become supersonic further increasing their velocity.
Pressure generated within an engine cannot be converted to velocity, particularly when a convergent nozzle
is used. The additional pressure results in additional
thrust which, as has been shown, must be added when the
total thrust developed by the engine is computed. The
additional thrust is developed ineffliciently. It would bemuch better to convert all of the pressure within theengine to velocity and develop all of the engine thrust bymeans of changes in momentum. In theory, a C-D nozzle
does this. Becuase it develops this additional part of thetotal thrust more efficiently, it enables an engine to
produce more total net thrust than the same basic engine
would generate if it were equipped with a conventional
convergent duct and nozzle. The C-D nozzle would benearly ideal if it could always be operated under the exact
conditions for which it was designed. However, if the rateof change in the duct area is either too gradual or too
rapid for the calculated increase in weight of the gases,
unsteady flow downstream of the throat will occur withan accompanying loss of energy. This ultimately means
loss of thrust. If the rate of increase in area of the duct is
too little, the maximum gas velocity that can be reachcd
will be limited. If the rate of increase is too great, the gas
flow will break away from the surface of the nozzle, andthe desired increase in velocity will not be obtained. As
exhaust gases accelerate or decelerate with changing
engine and flight conditions, their pressure fluctuates
above or below the pressure ratio for which the nozzle
was designed When this occurs, the nozzle no longer
converts all of the pressure to velocity, and the nozzle
begins to lose efficiency.
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FM 1-506

The solution to this dilemma is a C-D nozzle with avariable cross-sectional configuration which can adjust
itself to changing pressure conditions. Several types of
C-D nozzles have been tried, and a few have been usedsuccessfully on production aircraft. As the actual designand operation of such nozzles is usually either classified
military information or proprietary information of the
manufacturer, the nozzles cannot be described here.
THRUST REVERSERS
The difficult problem of stopping an aircraft after
landing increases many times with the greater grossweights common to large, modern aircraft with theirhigher wind loadings and increased landing speeds.
Wheel brakes alone are no longer the beat way to slow
the aircraft immediately after touchdown. The reversiblepitch propeller solved the problem for piston engine andturboprop-powered airplanes. Turbo rbojet and turbofan
aircraft, however, must rely on some device such as aparabrake or runway arrester gear or some means ofreversing the thrust produced by their engines.
Although sometimes used on military aircraft, theparabrake or drag parachute has distinct disadvantage
The parabrake is always subject to either a premature
opening or a failure to open at all. The parabrake must
be recovered and repacked after each use and, if
damaged or lost must be repaired or replaced. Once theparabrake has opened, the pilot has no control over theamount of drag on the aircraft except to release theparachute completely.
Arrester gears are primarily for aircraft carrier deck
operation although they are sometimes used by militarybases as overshoot barriers for land runways. They would
hardly be suitable for commercial airline operation at abusy municipal airport.
OIL SYSTEM SEALS
The significance of oil system seals in aircraft engines
is great. A leaking seal in a turbine engine could causetire, bearing failure, or cockpit fumes, to name a few
dangers. There are three main types of oil system seals:
synthetic, labyrinth, and carbon.
Synthetic
Synthetic seals (neoprene, silicone, Teflon, and synthetic rubber) are used throughout the engine’s oil systern. They are used where metal-to-metal contact would
not provide satisfactory sealing to withstand pressures insuch items as filters, turbine, and fittings. Seals come inmany sizes and shapes and are not normally reused. New
replacement seals are received from supply channels
3-20

usually in a package that prevents damage. In most cases
the packages will have a “cure date” stamped on theoutside. (Cure date is the date of manufacture of theseals.) This date is particularly important when installing
seals retie of rubber, which has a tendency to deteriorate
more rapidly than synthetic material. Just as importantis to use the proper seal with the correct part number fora specific installation. Never use a seal from anothersystem just because it looks like the right seal. The
composition or military specifications may be entirely
different which could cause the seal to fail at a crucial
moment.
Some synthetic seals coming into contact with synthetic oils such as MIL-7808 or ML-23609, have a tendency to swell; others might deteriorate completely.
Occasionally, seals are referred to as "packings" or "gaskets." However, there is a difference between the two.
Packing is used to provide a running seal; a gasket is used
between two stationary parts to create a static seal. Somemanufacturers refer to a gasket as a packing and viseversa. These terms should not be taken literally, Always
go strictly by the part number when using seals.
LabyrinthLabyrinth or air seals are designed to allow a small
amount of air to flow across the sealing surface.o This
helps prevent oil (or lower-pressured air) seepage acrossthe same surface (Figure 3-22). Air seals have two
separate parts. One part forms a plain or honeycomb
surface; the corresponding part is a circular seal withannular grooves. These grooves may use a soft metal as
the basic composition or be machined into a surface.
Matching the two together (one rotating portion or racewith one stationary) forms an air pressure seal. A seriesof soft metal knifelike edges rides very close to the seal
surface or cuts a path into a stationary honeycomb orsilver alloy air seal (Figure 3-23).
NOTE: When honeycomb or silver alloy isused, it is bonded to the stationary portionof the air seal.
Air for this seal is normally bled from the compressor and then forced between the sealing surface
and the seal. The effect of pressurization prevents oil (orlower-pressured air) from seeping from one section toanother during engine operation. Air seals work only
when the engine is operating when the engine is shut
down seal leakage will occur. Be extremely careful whenworking in or around the seal area because seals arecomposed of very soft metal. Avery small nick or groovein a seal may cause a serious oil leak which may requirea premature engine change.

FM 1-506—

3-21

FM 1-506

CarbonCarbon oil seals are used to contain the oil within the
bearing areas on most jet turbine engines. All carbon
seals form a sealing surface by having a smooth carbon
surface rub against a smooth steel surface. The steel
surface is called a “seal race” or "faceplate," dependingon the engine manufacturer. All carbon seals arepreloaded. That is, the carbon must in some way bepressed against the steel surface. Three common
preload methods are spring tension, centrifugal force,and air pressure. During operation, the seal may be aided
by allowing a small amount of oil to flow into the rubbing
surface. The oil also cools the seal as a certain amount ofheat is built up by the carbon rubbing on the steel surface.
The carbon oil seal shown in Figure 3-24 consists of
two rows of carbon segments (seal ring and back ring)
mounted in a housing and held together around theircircumference by extension springs. These springs not
only hold the segments together by circling the outside
but also serve as the preload necessary to press the seals
inward. The seal segments nearest the bearing have a lipthat forms the seal; the positioning pads contact the steel
race and maintain the proper sealing positions of thesegments. These positioning pads are sometimes
referred to as “wear blocks” because the seal lip is verythin and without the pads would have a short wear life.
The grooves between the pads are staggered to reduce
airflow toward the sump. On this particular seal, the seal
race contact surface is cooled by a spray of oil. The key

and spiral pin shown in the figure keep the carbon segments from turning within the seal housing the compression springs press the seal segments into the housing. The
entire assembly is held together by the spring retainer andsnap ring. This type of seal is stationary and rubs against
an inner, rotating seal race.
Other configurations of carbon seals may haveseveral seals on each side of the bearing. They may also
rub on the side or outer surface of the seal rather than the
inner surface as the one illustrated does. Seals such as
the one in Figure 3-25 can be rebuilt by replacing thesegments in groups; other seals however, must be
returned to an overhaul facility when they are damaged
and must be replaced with a complete seal assembly.

MATERIALS AND METHODS OF CONSTRUCTION
High-temperature, high-strength materials andunique methods of manufacture have made the gas
turbine engine a practical reality in a few decades. The
performance of turbojet and turboprop engines depends
largely on the temperature at the inlet to the turbine.
Increasing the turbine inlet temperature from the present
limit (formost highly
produced engines) of approximate00ly 1700 F to 2500 F will result in a specific thrust increaseof approximately 130 percent along with a corresponding
decrease in specific fuel consumption. For this reason
high cycle temperatures are desirable. Not all materials
can withstand the hostile operating conditions found inparts of the gas turbine engine.
Properties and Characteristics
Metallurgists have been working for almost 50
years improving metals for use in aircraft construction. Each type of metal or alloy has certain properties
3-22

FM 1-506

and characteristics which make it desirable for a particular use, but it may have other qualities which are
undesirable. The metallurgist’s job is to build up thedesirable qualities and tone down the undesirable ones.
This is done by the alloying (combining) of metals and by
Various heat-treating processes. It is not necessary for theairframe repairer to be a metallurgist, but it is advantageous to have a general knowledge of the properties
used in their development. The repairer should befamiliar with a few metallurgical terms. The following
terms are used in describing the physical properties andcharacteristics of metals.
Hardness. Hardness refers to the ability of ametal to resist abrasion, penetration, cutting action, or permanent distortion. Hardness may be
increased by working the metal and, in the caseof steel and certain aluminumalloys, by heattreatment and cold-working. Structuralparts are often formed from metals in their soft
state and are then heat-treated to harden themso the finished shape is retained. Hardness andstrength are closely associated properties ofmetals.
Brittlemem. Brittleness is the property of ametal which allows little bending or deformation without shattering. A brittle metal is apt tobreak or crack without change of shape. Because structural metals are often subjected toshock loads, brittleness is not a desirable
property. Cast iron or cast aluminum and very
hard steel are brittle metals.
Malleability. A metal which can be hammered, rolled, or pressed into various shapes
without cracking breaking or other detrimental
effects is said to be malleable. This property isnecessary in sheet metal which is worked into
curved shapes such as cowling, andwingtips. Copper is an example of a malleable
metal.
Ductility. Ductility is the property of a metal
which permits it to be permanently drawn, bent,
or twisted into various shapes without breaking.
This property is essential for metals used in
making wire and tubing. Ductile metals are
preferred for aircraft use because of their easeof forming and resistance to failure undershock loads. For this reason, aluminum alloys
are used for cowl rings, fuselage and wing skin,and formed or extruded parts such as ribs, spars,and bulkheads. Chrome-molybdenum steel isalso easily formed into desired shapes. Ductility
is similar to malleability.
Elasticity. Elasticity is that property whichenables a metal to return to its original shapewhen the force which causes the change ofshape is removed. This property is extremely
valuable, as it would not be desirable to have apart permanently distorted after an applied load
was removed. Each metal has a point known as
the elastic limit beyond which it cannot beloaded without causing permanent distortion.
In aircraft construction, members and parts are
so designed that the maximum loads to which
they are subjected will never stress them beyond
their elastic limits. This desirable property ispresent in spring steel.
Toughness. A material which possesses
toughness will withstand tearing or shearing andmay be stretched or otherwise deformed
without breaking. Toughness is a desirable
property in aircraft materials.
Conductivity. Conductivity is the property
which nables a metal to carry heat or electricity.
The heat conductivity of a metal is especially
important in welding as it governs the amountof heat that will be required for proper fusion.
Conductivity of the metal, to a certain extent,
determines the type of jig to be used to control
expansion and contraction. In aircraft,electrical conductivity must also be considered
in conjunction with bonding to eliminate radio
interference. Metals vary in their capacity to
conduct heat. Copper, for instance, has a relatively high rate of heat conductivity and is a goodelectrical conductor.
Fusibility. Fusibility is the ability of a metal to
become liquid by the application of heat. Metals
are fusedin welding.
Steels fuse at approximateooaluminumalloys fuse atly 2500 F (1371 C), andooapproximately 1100 F (593 C).
Density. Density is the weight of a unit volume
of a material. In aircraft work, the actual weightof a material per cubic inch is preferred, as thisfigure can be used in determining the weight ofapart before actual manufacture. Density is an
important consideration when choosing amaterial to be used in the design of a part inorder to maintain the proper weight andbalance of the aircraft.
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FM 1-506

Contraction and expansion. contraction andexpansion are reactions produced in metals as
the result of heating or cooling. A high degreeof heat applied to a metal will cause it to expand
or become larger. Cooling hot metal will shrink
or contract it. Contraction and expansion affectthe design of welding jigs, castings, andtolerances necessary for hot rolled material.
Strength. Strength refers to the ability of ametal to hold loads (or forces) without breaking
Strength is a property that sums up many of thedesirable qualities of metals. Strength with
toughness is the most important combination ofproperties a metal can possess. Metals
having this combination of properties are usedin vital structural members that may become
overloaded in service.
In a discussion of metal properties, stress and strain
should be mentioned. Stress is a force placed upon abody and is measured in terms of force per unit area, theforce being expressed in pounds and the unit of area insquare inchs; in other words, pounds per square inch(psi). Stress may be in the form of compression, tension,
torsion, bending, shearing loads, or a combination of two
or more of these. All parts of an aircraft are subjected tostresses. When a part fails to return to its original form
after being stressed, it is said to be strained. The various
stresses acting on parts of an aircraft, while in flight, havean important&wing on the metals used:Tension. Tension is the resistance to pullingapart or stretching produced by two forces pulling in opposite directions along the samestraight line. The cables of a control system areplaced under tension when the controls areoperated.
Compression. Compression is the resistance topushing together or crushing produced by twoforces pushing toward each other in the samestraight line. The landing struts of an aircraft areunder compression when landing and, to asmaller degree, when supporting the weight of
the aircraft as it rests on the ground.
Torsion. Torsion is the resistance to twisting. A
torsional force is produced when an engine
turns a crankshaft. Torque is the force that
produces torsion.
Bending. Bending is a combination of tension
and compression. The inside curve of the bent
object is under compression, and the outside
curve is under tension. The main spars of themain rotor blades are subjected to bending.
3-24

The blades droop while the rotor head is at restand bend upward when rotating.
Shear. Shear is the stress exerted when twopieces of metal fastened together are separated
by sliding one over the other in opposite directions. The stress cuts off a bolt or a rivet like apair of shears. In general, rivets are subjected toshear only; bolts, to both shear and tension.
There is internal shear in most bending elements and in the skin of sheet metal structures.
Metallurgical and Metalworking Terms
Some of the more commonly used terms in the fieldof mettalurgy are listed below:
Strength.
— Creep strength-the only ability of a metal
to resist slow deformation due to stress, butless than the stress level needed to reach theyield point. Creep strength is usually statedin terms of time, temperature, and load.
Yieldstrength – the point reached when—metal exhibits a permanent set under load.
— Rupture strength– the point where metal
will break under a continual load applied for
periods of l00 and 1000 hours. Metals areusually tested at several temperatures.
— Ultimate tensile strength– the load underwhich metal will break in a short time.
Ductility– the ability of a metal to deform
without breaking.
Coefficient of expansion – a measure of how
much a metal will expand or grow with theapplication of heat.
Thermal conductivity-a measure of the abilityof a metal to transmit heat.
Corrosion and oxidation resistance– indicates
how well a metal can resist the corrosive effectsof the hot exhaust stream.
Melting point-the temperature at which metal
becomes a liquid.
Critical temperature – point where as it iscooled a metal’s internal structure and physicalproperties are altered. The rate of cooling will
greatly influence the ultimate properties of themetal.
Heat treatability– a measure of how the metal’s
basic structure will vary under an operation orseries of operations involving heating and cooling of the metal while it is in a solid state. Ferritic, austenitic, and martensitic steels all vary inheat treatability.

FM 1-506

Thermal shock resistance—the ability of a metal
to withstand extreme changes in temperature in
short periods of time.
Hardness—the ability of a metal to resist
abrasion penetration cutting action or permanent distortion. Hardness may be increased
by working the metal or by heat treatment.
Alloying—the combining of metals in their molten state to create desirable qualities or reduceor eliminate undesirable qualities.
Common metal working terms include the following:
Casting—a process whereby molten metal
solidifies in a mold.
Forging—a process of plastic deformation
under pressure that may be slowly or quickly
appliedElectrochemical machining (ECM)—controlled high-speed deplating using a shaped tool
(cathode), an electricity-conducting solution
and the workpiece (anode).
Machining—any process where metal is formedby cuttin,- hot or cold rolling pinching, punching, grinding, or by laser beams.
Extrusion—pushing through a die to form
various cross-sectional shapes.
Welding—a process of fusing two pieces ofmetal together by locally melting part of thematerial by arc welders, plasmas, lasers, orelectron beams.
Pressing—a process of blending, pressing,sintering (fusing the power particles together
through heat), and coining metals out of prealloyed powders.
Protective finish and surface treatment—includes plating by electrical and chemical
processes ceramic coatings, or painting. Surface treatments for increased wear may take theform of nitriding, cyaniding, carburizing, diffusion coating and flame plating.
Shot peening—a plastic flow or stretching of ametal’s surface by a rain of round metallic shot
thrown at high velocity. Shot peening also
serves to work-harden metals, especiallyaluminum alloy.
Heat treatment—a series of operations involving the controlled heating and cooling of metals
in the solid state. Its purposes is to change amechanical property or combination of properties so that the metal will be more useful,serviceable, and safe for a specific purpose. It

includes normalizing annealing stress relievingtempering and hardening.
Inspection (strictly speaking not a part of themetalworking process but integrally associatedwith it) – includes magnetic particle and dye
penetrant inspection, X-ray inspection dimensional and visual inspection, and inspection by
devices using sound, light, and air.
Heat Ranges of Metals
The operating conditions inside a gas turbine engine
vary considerably, and metals differ in their ability tosatisfactorily meet these conditions.
Aluminum Alloys. Aluminum and its alloys areused in temperature ranges up to 500ºF. With low density
and good strength-to-weight ratios, aluminum forgingsand castings are used extensively for centrifugal compressor wheels and housings, air inlet sections, accessory
sections, and for the accessories themselves.
Magnesium Alloys. Magnesium is the world’s
lightest structural metal. Aluminum is 15 times heavier,titanium 25 times heavier, steel 4 times heavier, andcopper and nickel alloys are 5 times heavier. Magnesium
is combined with small amounts of certain other metals,
including aluminum, manganese, zinc, zirconium,
thorium, and others, to obtain the strong lightweightalloys needed for structural purposes.
Titanium Alloys. Titanium and its alloys are used
for axial-flow compressor wheels, blades, and otherforged components in many large, high-performanceengines. Titanium combines high strength with low density and is suitable for applications up to 100ºF.
Steel Alloys. This group includes high-chromium,molybdenum, high-nickel, and iron-base alloys in addition to low-alloy steels. Because of the relatively low
material cost, ease of fabrication, and good mechanicalproperties, low-alloy steels are commonly used for bothrotating and static engine components such as compressor rotor blades, wheels, spacers, stator vanes and structural members. Low-alloy steels can be heat-treated andcan withstand temperatures up to 100°F. High nickelchromium iron-base alloys can be used up to 1250ºF.
Nickel-Base Alloys. Nickel-base alloys are some of
the best metals for use between 1200ºF and 1800ºF. Most
contain little or no iron. They develop high-temperaturestrength by age hardening and are chacterized by longtime creep-rupture strength as well as high ultimate andyield strength combined with good ductility. Many ofthese materials, originally developed for turbine bucket
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FM 1-506

applications, are also being used in turbione wheels, shafts,
spacers, and other parts. Their use is somewhat
restricted because of cost and because of their requirement for critical materials.
Cobalt-Base Alloys. Colbalt-base alloys form
another important group of high-temperature, highstrength, and high-corrosion- resistant metals. They contain little or no iron. These alloys are used in afterburnerand other parts of the engine subjected to very hightemperatures.
Chemical Elements of Alloys
The number of materials used in alloys is large.
Some of the most commonly used elements are listed
below.

horsepower needed to drive the compressor. There isalso a severe temperature gradient (difference) of several
hundred degrees between the central portion of the disk
and its periphery.
Many metals which would be quite satisfactory atroom temperatures will lose much of their strength at theelevated temperatures encountered in the engine’s hot
section. The ultimate tensile strength of a metal at one
temperature does not necessarily indicate its ultimate
tensile strength at a higher temperature. For example, at
l000ºF Inconel X has an ultimate tensile strength ofapproximately 160,000 psi; and S 816 at the same
temperature has an ultimate tensile strength of 135,000
psi. At 1500ºF their positions are reversed. Inconel X has
an untimate tensile strength of 55,000 psi; S 816 has an

The percentages of elements used partially determines the physical and chemical charteristics of the alloy
and its suitability to a particular application, Temperingand other processes determine the rest. Three characteristics that must be considered are—High-temperature strength.
Resistance to oxidation and corrosion.
Resistance to thermal shock.
High-Temperature Strength. The most highly stressedparts of the gas turbine engine are the turbine blades anddiscs. Centrifugal forces tending to break the disc varywith the square of the speed. For example, the centrifugal
force on a disc rotating at 20,000 RPM will be four times
that at 10,000 RPM. Blades weighing only 2 ounces mayexert loads of over 4000 pounds at maximum RPM.
Blades must also resist the high bending loads applied by
the moving gas stream to produce the thousands ofultimate tensile strength of 75,000 psi. The creep
strength, which is closely associated with ultimate tensile
strength, is probably one of the most important considerations in the selection of a suitable metal for turbine
blades. Engine vibration and fatigue resistance will also
have some influence on the selection and useful life of
both discs and blades.
Many materials will withstand the high temperaturesencountered in a gas turbine engine (carbon columbium,
molybdenum, rhenium, tantalum, and tungsten all havemelting points above 4000ºF). However, the ability towithstand high temperatures while maintaining
reasonable tensile strength is not the only consideration.
All of the following qualities must be taken into accountwhen selecting a particular metal:
Critical temperature.
Rupture strength.

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FM 1-506

Thermal conductivity.
Coefficient of expansion.
Yield strength.
Ultimate tensile strength.
Corrosion resistance.
Workability.
Cost.
Resistance to Oxidation and Corrosion. Corrosion
and oxidation are results of electrical and chemical reactions with other materials. The hot exhaust gas stream
encountered in the engine speeds up this reaction. While
all metals will corrode or oxidize, the degree of oxidation
is determined by the base alloy and the properties of theoxide coating formed. If the oxide coating is porous or
has a coefficient of expansion different from that of thebase metal, the base metal will be continually exposed tothe oxidizing atmosphere. One solution to oxidation atelevated temperatures is ceramic coatings. Ceramiccoated afterburner liners and combustion chambers arein use today. The ceramic coating has two basic functions:
Sealing the base metal surface against corrosion.
Insulating the base metal against high temperatures.
These coatings are not without disadvantages:
They are more susceptible to thermal shock.
They must have the same coefficient of expansion as the base metal.
They are brittle.
They have low tensile strength which restricts
their use in the engine.
Some promising work is being done with various
metal-ceramic combinations called Cermets or
Ceramels. Materials being used with ceramics include
aluminum, beryllium, thorium, and zirconium oxides, toname a few.
Resistance to Thermal Shock. Many materials which
would otherwise be quite suitable must be rejectcd because of their poor thermal shock characteristics. Several
engine failures have been attributed to thermal shock on
the turbine disc. Ceramic coating in particular are vulnerable to this form of stress. Improved fuel controls,
starting techniques, and engine design have lessened thisproblem.
Transpiration Cooling
The effort to achieve higher turbine inlet temperatures (and therefore higher thermal efficieney) has been

approached from two directions: (1) high-temperaturematerials and (2) cooling methods. A common methodof cooling the nozzle guide vanes on gas turbine engines
is to pass compressor bleed air through the hollow blades
to cool them by convective heat transfer. Some engines
also use air bled from the compressor to cool the front
and rear face of the turbine discs and the hollow turbine
blades.
Transpiration cooling is a novel and efficient methodof allowing the turbine blades and other parts within the
hot section to operate at much higher turbine inlet
temperatures. The Wright Corporation has constructed
and run turbine blades at an inlet temperature of 2500ºF.
In this type of cooled blade the air passes throughthousands of holes in a porous airfoil made from asintered wire mesh material. Since the sintered wire
mesh is not strong enough by itself, an internal strut isprovided as the main structural support carrying all air foil
and centrifugal loads. Fabrication techniques involve
rolling layers of woven wire mesh and then sintering theselayers to forma porous metal sheet. The sheet is thenrolled into an airfoil shape.
Porous materials have been tested for use in combustion chambers and for afterburner liners. A similarmaterial called Rigimesh has also been used in rocket
engines to help keep the fuel nozzles cool. Many
manufacturers are experimenting with other types ofporous materials for use in blades in an attempt to obtainhigher turbine inlet temperatures.
Other Materials
Relatively new materials called composites are coming into use in both airframes and engines. In theseproducts graphite, glass or boron filaments are embedded in an epoxy-resin matrix or base substance.
Other types of filaments and materials are being tried tomeet the demands of higher temperatures and stress
The chief advantage of composite material is its very
favorable strength-to-weight ratio, which can lead tolighter weight in many structural parts. For example, alighter fan blade allows a lighter fan disc, which in turn
permits a lightening of other parts all the way down theline. Composite materials may be used in conjuction withother load-bearing materials to provide a support function. Typical of this type of structure are fan blades witha steel spar and base and with an airfoil composite shell.
Manufacturing Techniques
Basic parts of the engine are produced by several
casting and forging processes, literally dozens of machine3-27

FM 1-506

operations, and fabrication procedures using a variety ofmetal-joining methods.
Casting. Several engine parts are cast in aluminum,magnesium, or steel alloys. These parts include intake
and compressor housings, accessory- and blading.
to name a few. Casting methods differ. They include –Sand casting.
Spin Casting.
Lost-wax or investment casting.
Resin-shell mold casting.
Slip casting.
Mercasting.
Sand castings. Sand casting uses a wood or metal
pattern around which a clay-he sand has been packed
to form the mold. The mold is then split, the patternremoved, the mold reassembled, and any cores that arenecessary added. Molten metal at a precise temperature
is poured into the mold and allowed to cool. The moldis removed, and various heat treatments may be performed to obtain the desired physical characteristics.
The casting may be spun while being poured.
Spin casting. Spin casting results in a denser,
sounder casting. Spinning is normally performed on
small ring sections. Cooling of the metal radially inward
results in fewer stresses.
Lost-wax or investment casting. Basically, the investment casting process involves the use of heatdisposable wax or plastic patterns which are surroundedwith a refractory material to form a monolithic mold
Patterns are removed from the mold in ovens, and molten
metal is poured into the hot mold. Sometimes this pouring is done in a vacuum furnace. After cooling the moldmaterial is quite fragile and easily removed from thecasings. Because the finished product duplicates thepattern exactly, the fabrication of patterns is critical.
They are made by injecting molten wax or plastic into
metal dies. The finished castings have an exceptionallysmooth surface finishing and require very little further
machining. Incidentally, this process is not new. It wasused by the ancient Greeks and Egyptians to cast
lightweight statues, intricate bowls, and pitchers.
Resin-shell mold casting. Resin-shell mold castingis a high-production method similar to investment casting
except that the tolerances are not held as closely. In many
ways it rivals sand casting in economy.
Slip casting. Slip casting (which was borrowed
from the ceramics industry) is used for super heatresistant materials. Often it is the only way certain
3-28

materials can be shaped. Metal ceramics, silicon nitride,
and refractory metals cast this way can withstandtemperatures over 2200ºF.
Mercasting. The Mercast process is a precision
casting technique. It is essentially the same kind of
method as the lost-waxer investment process except thatfrozen mercury is used as a pattern instead of wax. Liquidmercury is poured into a master mold where it is frozen
at temperatures below 40ºF. Then it is removed andcreated with a cold refractory slurry to a thickness of 1/8
inch or more. The refractory shell is dried at low
temperature then the shell and mercury are brought toroom temperature and the mercury is melted out. The
refractory shell is fired to give it strength and then is usedas the mold for a usual casting process. Complicatedparts can be made with the Mercast process. Very close
tolerances and excellent surface finish can be obtained
The cost, however, is higher than that of some othermethods.
Forging. Disks, drive Shafts, gears, vanes,
blades, and numerous other parts of the gas turbine
engine are manufactured by forging. This process allows
the development of a grain structure and results in a finely
grained, more ductiIe, strong dense product. Forging is
by rapid hammering or slow pressing. The choice oftechnique depends on the resistance of metal to rapid
deformation. The workpiece is generally heated to improve plasticity and reduce forging forces. It will oftenpass through several different dies before the final shapeis obtained. All ductile materials can be forged but theirforgeability varies considerably. Forgeability generally
depends upon –Melting point.
Ductility.
Yield strength.
Crystallographic structure.
Recovery from forging stresses
Surface reactivity.
Die friction.
Cost.
Machining. Common tools used to manufacturegas turbine parts include lathes, mills, broaches, grinders,shapers and planers, polishers and buffers, drills, saws,
shears, filers, threaders, contour machines of all kinds,
and a host of other devices to cut and form metal. Manyof these devices use a numerical tape control or otherautomatic control devices to reduce human error andproduce a more uniform, less expensive product. Robots
equipped with computers also assist in machining parts.

FM 1-506

Some nontraditional machining techniques forremoving metal from super hard and super tough alloys
and other materials whose complex shapes preclude
conventional metal-cutting tools include—
Chemical milling.
Electrochemical machining (ECM).
Electric-discharge machining (EDM).
Electron-beam machining.
Laser-beam machining.
Other nonconventional machining includes everything from abrasive jet cutting to ultrasonic machining.
Chemical milling. Chemical milling involves theremoval of metal by dissolving it in a suitable chemical.
Those areas that are not to be dissolved are masked withnonreactive materials. The process can be used on mostmetals, including aluminum, magnesium, titanium, steels,and superalloy for surface sculpturing. Both sides of theworkpiece can be chemically milled simultaneously. In
addition, the process can be used to machine very thinsheets.
Electrochemical machining. Electrochemicalmachining is basically a chemical deplating process inwhich metal, removed from a positively charged
workpiece using high-amperage, low-voltage DC, isflushed away by a highly pressurized electrolyte before it
can plate out on the cathode tool. The cathode tool ismade to produce the desired shape in the workpiece;
both must be electrically conductive. The work proceeds
while the cathode and workpiece are both submerged inan electrolyte such as sodium chloride. A variation andextension of electrochemical machining is electro-stream
drilling. In this process a negatively charged electrolyte,
usually an acid, drills holes in a workpiece that has been
positively charged Holes as small as 0.005 inch indiameter and 0.5 inch deep in superalloys can be drilledin this manner.
Electric-discharge machining. In electric-dischargemachining high voltages are used to produce a highelectrical potential between two conductive surfaces (theworkpiece and electrode tool) both of which are immersed in a dielectric fluid. Material is removed fromboth the electrode and the workpiece by a series of very
short electric discharges or sparks between the two andis swept away by the dielectric fluid. More material isremoved from the workpiece than from the tool byproper selection of the two materials. This process canbe used to shape complex parts to very close tolerancesfrom refractory metals and alloys that were formerly
impossible to machine. The use of electric-dischargemachining is limited in that it is slower than electrochemical machining tool replacement can become expensive,
and the surface of the workpiece is damaged as a resultof the sparks. On the other hand, the EDM process isless expensive than the ECM process.
Electron- and laser-beam machining. Electronbeam machining and laser-beam machining are being
used experimentally. They may find future use in the
production of gas turbines and other aerospace components.
Fabrication. Welding is used extensively to fabricate and repair many engine parts. Fabricated sheet steel
is used for combustion chambcrs, exhaust ducts, compressor casings, thrust reversers, silencers, and so forth.
Common methods include resistance and inert-gas(usually argon) welding. Less common methods use
plasmas and lasers. Electric resistance welding is used tomake spot, stitch (overlapping spots), and continuousseam welds. Inert-gas welding employs a nonconsumable electrode (tungsten-thorium alloy) surroundedby some inert gas such as argon or helium. The gas
prevents an adverse reaction with the oxygen present in
the normal atmosphere. The inert gas can be applied in
the immediate area of the arc. In the case of production
runs the workpiece or the entire welding machine can be
enclosed in a thin plastic balloon, sometimes as large asa room. The entire plastic bubble is filled with andsupported by the inert gas. The operator stands on theoutside and works through specially designed armholes.
After welding, many parts must be stress-relieved.
Where temperature or working loads are not large, brazing or silver soldering may be used to join such parts as
fittings and tube assemblies.
Electron-beam welding is showing great promise asa method of fabricating parts from heretofore difficult to
weld or unweldable materials. Electron-beam welding
uses a stream of focused electrons traveling at speeds
approaching 60 percent of the speed of Iight. Even
though the mass of electrons which form the beam issmall they are traveling at such speeds that they containa great amount of kinetic energy. When the beam strikes
the workpiece, the kinetic energy is transformed into heat
energy. The welding usually takes place in a vacuum,
although nonvacuum techniques can be used. The following characteristics make this welding process a valuable one in the gas turbine manufacturing area:
Deep narrow welds with a very narrow heataffected zone in the base metal.

3-29

FM 1-506

The ability to weld materials as thin as 0.00025
inch and as thick as 4 inches of stainless steel.
The ability to weld many different types ofmaterials.
Finishing. The basic material, the properties
desired in the finished product, and the kind of protection
desired determine the type of surface and internal treatment received. The variety is considerable and includes
the following:
Chemical treatment. Chrome pickling is themost commonly used of all chemical treatmentsof magnesium. The part is dipped in a solutionof sodium bichromate, nitric acid, and water.
Electrochemical treatment. Anodizing is acommon surface treatment for aluminum alloys
in which the surface aluminum is oxidized to anadherent film of aluminum oxide.
Painting. A thin preservative resin varnish coating is used to protect internal steel, aluminum,and magnesium parts. The characteristic colorof this shiny, transparent coating is usually green
or blue-green. A graphite powder may bemixed with the varnish to act as an antigellant.
Grey, black, or aluminum enamel or epoxy
paint is also used extensively as a protective
finish.
Shot peering. This procedure can increase thelife of apart many times. It is essentially a plastic
flow or stretching of a metal’s surface by a rainof round metallic shot thrown at high velocity byeither mechanical or pneumatic means. The
0.005- to 0.035-inch stretched layer is placed ina state of compression with the stress concentration uniformly distributed over the entire surface. This aids to increase surface hardness.
Glass beads are sometimes used as the shot forcleaning purposes.
Plating. A great number of plating materials
and procedures are used. Plating materials involving the use of chemical or electrochemicalsolutions include cadmium, chromium, silver,
nickel, tin and others. The exact procedure isdetermined by the plating and base metal.
Aluminizing. This is another plating method
whereby pure molten aluminum is sprayed onto
the aluminum alloy base material to form aprotective coating against oxidation and corrosion.
The Coating Service of Union carbide Corporation
has developed and is producing machines for applying
3-30

extremely wear-resistant and other specialized coating
to gas turbine parts, tools, and other machines. The
different coatings are applied by either of two methods–
the detonation gun (D-gun) or the plasma gun. Fourtimes a second, a spark ignites the mixture and creates adetonation which hurls the coating particles, heated to aplastic state by the 6000ºF temperature in the gun, out of
the barrel at a speed of 2500 feet per second. The part tobe plated is kept below 300ºF by auxiliary cooling streams.
The high-level sound of 150 decibels necessitates housing
the gun in a double-walled, sound-insulated construction.
Operation is controlled from outside this enclosure.
The plasma gun or torch produces and controls ahigh-velocity, inert-gas stream that can be maintained attemperatures above 20,000ºF. Unlike the D-gun process
no combustion takes place. The high-temperature plasma is formed by ionizing argon gas in the extreme heat ofan electric arc. Gas molecules absorb heat energy fromthe arc split into atoms, and then further decompose into
electrically charged particles called ions. The hot gas
stream can melt any known material, without decomposition. When the molten particles, which are introduced in
powdered form, strike the part being coated, a permanent welded bond is formed. While the D-gun is apatented Union Carbide machine, other manufacturers
make and distribute a variety of plasma-plating and cutting devices.
Heat Treatments. All of the following procedures
alter the mechanical properties of steel to suit the end
Normalizing– steel is heated to a temperature
above the critical range and allowed to coolslowly. Normalizing promotes uniformity ofstructure and alters mechanical properties.
Annealing-consists of heating to a point at ornear the critical range, then cooling at apredetermined rate. It is used to develop softness, improve machine abliity, reduce stress, improve or restore ductility, and modify otherproperties.
Stress relieving–metal is heated throughout toa point below the critical range and slowly
cooled. The object of this treatment is to restore
elastic properties or reduce stresses that may
have been induced by machining cold working,or welding.
Solution heat treatment – done with aluminumalloys. Consists of heating the metals to thehighest temperature without melting andFM 1-506

holding the metals at that temperature longenough to produce a solid solution of the hardening components in the alloy. This is followed by rapid quenching to retain the
produced condition.
Hardening– involves heating the metal to atemperature above the critical range and thenquenching The cooling rate will determine
hardness.
Tempering– steel is usually too brittle for use
after quenching. Tempering restores some ofthe ductility and toughness of steel at thesacrifice of hardness or strength. The process isaccomplished by heating the hardened steel to a
specific point below the critical temperature,holding it at that temperature for a specifictime, and then cooling material in a controlled
area.

Nonmetallic Materials
Teflon, nylon, carbon, rubber, Bakelite, and a host of
plastic materials are used in the gas turbine engine mainly
as sealing and insulation materials. For example, nylon
and Teflon are used to insulate and protect the shielded
electrical wiring located on the outside of the engine.
Teflon is also used on the J-79 F4 Phantom engine for thesealing the variable- stator-vane actuators. Carbon isused largely inside the engine in the form of carbon
rubbing seals. Some of these “face” carbon rubbing seals
must be flat to within two helium light bands, or approximately 23 millionths of an inch Rubber and rubberized fabric materials makeup the sealing edge of thefire seal which divides the hot and cold sections of theengine when mounted in the nacelle. Synthetic rubber isused extensively throughout the engine in the form ofO-rings or other shaped seals.

3-31

FM 1-506

CHAPTER 4

FUEL SYSTEM

The fuel system for gas turbine and reciprocating
engines is the same. It delivers to the engine fuel meteringsystem a uniform flow of clean fuel at the proper pressure
and in the necessary quantity to operate the engine.
Despite widely varying atmospheric conditions, the fuelsupply must be adequate and continuous to meet thedemands of the engine during flight.
DESCRIPTION
The fuel system is one of the more complex aspectsof a gas turbine engine. The variety of methods used tomeet turbine engine fuel requirements makes reciprocating engine carburetion seem simple by comparison.
Thrust
It must be possible to increase or decrease the power
at will to obtain the thrust required for any operating
condition. In turbine-powered aircraft, this control isprovided by varying the flow of fuel to the combustion
chambers. However, turboprop and turboshaft aircraft
also use variable-pitch propellers or helicopter rotors;
thus, the selection of thrust is shared by two controllable
variables, fuel flow and propeller blade or rotor blade
pitch.
Fuel Supply
The fuel supply must be adjusted automatically tocorrect for changes in ambient temperature or pressure.
An excessive fuel supply in relation to mass airflow
through the engine can cause the limiting temperature of
the turbine blades to be exceeded. It can produce compressor stall and surge.
Fuel Delivery
The fuel system must deliver fuel to the combustion
chambers in the right quantity and in the right condition
for satisfactory combustion. Fuel nozzles form part of thesystem and atomize or vaporize the fuel so that it tiff
ignite and burn efficiently. The fuel system must also
supply fuel so that the engine can easily be started on the
ground and in the air. This means fuel must be injected
into the combustion chambers in a combustible condition

when the engine is being turned over slowly by the starting
system. Combustion must be sustained while the engine
is accelerating to its normal running speed.
Acceleration
Another critical condition to which the fuel system
must respond occurs during a quick acceleration. Whenthe engine is accelerated, energy must be furnished to theturbine in excess of that necessary to maintain a constant
RPM. However, if the fuel flow increases too rapidly, anoverrich mixture maybe produced causing a surge.
Fuel Control
Turbojet, turbofan, turboprop, and turboshaft engines are equipped with a fuel control unit which automatically satisfies the requirements of the engine.
Although the basic requirements apply generally to allgas turbine engines, the way in which individual fuelcontrols meet these needs cannot be conveniently
generalized. Each fuel control manufacturer has its own
way of meeting the engine demands.
TURBINE ENGINE FUEL SYSTEMCOMPONENTS
Main Fuel Pumps, Engine-Driven
Main fuel pumps deliver a continuous supply of fuelat the proper pressure during operation of the aircraft
engine. Engine-driven fuel pumps must be able to deliver
the maximum flow needed at high pressure to obtainsatisfactory nozzle spray and accurate fuel regulation.
Fuel pumps for turbojet engines are generally positive displacement gear or piston types. The term “positive displacement” means that the gear or piston will
supply a fixed quantity of fuel to the engine for every
revolution of the pump gears or for each stroke of thepiston.
These fuel pumps may be divided into two distinct
system categories constant displacement and variabledisplacement. Their use depends on the system used toregulate the flow of fuel. This maybe a pressure relief
valve (barometric unit) for constant displacement (gear)

4-1

FM 1-506

pump or a method for regulating pump output in thevariable displacement (piston) pumps.
Constant-Emplacement Pump
Gear-type pumps have approximately straight-line
flow characteristics, whereas fuel requirements fluctuatewith flight or ambient air conditions. A pump of adequate capacity at all engine operating conditions will haveexcess capacity over most of the range of operation. Thisis the characteristic which requires the use of a pressure
relief valve for disposing of excess fuel. The impeller,
which is driven at a greater speed than the high-pressure
elements, increases the fuel pressure from 15 to 45 psi,
depending on engine speed.
This fuel is discharged from the boost element (impeller) to the two high-pressure gear elements. Each ofthese elements discharges fuel through a check valve to acommon discharge port. High-pressure elements deliver
approximately 51 gallons per minute at a discharge pressure of 850 psig.
Shear sections are incorporated in the drive systemsof each element. If one element fails, the other element
continues to operate. Check valves prevent circulation
through the inoperative element. One element can supply enough fuel to maintain moderate aircraft speeds.
A relief valve is incorporated in the discharge portof the pump. This valve opens at approximately 900 psi
and is capable of bypassing the total flow at 960 psi.
Excess fuel is recirculated. The bypass fuel is routed tothe inlet side of the two high-pressure elements.
Variable-Displacement Pump
The variable-displacement pump system differsfrom the constant-displacement pump system. Pump
displacement is changed to meet varying fuel flow requirements; that is, the amount of fuel discharged fromthe pump can be made to vary at anyone speed. With apump; of variable flow, the applicable fuel control unit
can automatically and accurately regulate the pump pressure and delivery to the engine.
Where variable-displacement pumps are installed
two similar pumps are provided and connected in parallel. Each pump can carry the bad if the other fails during
normal parallel operations. At times one pump maybe
insufficient to meet power requirements. Pump duplication increases safety in operation, especially during
takeoff and landing
The positive-displacement, variable-stroke-typepump incorporates a rotor, piston, maximum speed
governor, and relief valve mechanism.
4-2

Fuel Filters
All gas turbine engines have several fuel filters at
various points along the system. It is common practice toinstall at least one filter before the fuel pump and one on
the high-pressure side of the pump. In most cases thefilter will incorporate a relief valve set to open at a
specified pressure differential to provide a bypass for fuelwhen filter contamination becomes excessive.
Paper Cartridge. This filter is usually used on the
low-pressure side of the pump (Figure 4-1). It incorporates a replaceable paper filter element which iscapable of filtering out particles larger than 50 to 100
microns (the size of a human hair). (One micron equals
0.000039 inch, or 25,400 microns equal one inch.) The
cartridge protects the fuel pump from damage due to fuelcontamination.
Screen Disc. This disc (Figure 4-2) is located on theoutlet side of the pump. This filter is composed of a stackof removable fine wire-mesh screen discs which must be
dissassembled and cleaned periodically in an approved
solvent.
Screen. This screen is generally used as a lowpressure fuel filter. Some of these filter screens areconstructed of sinter-bonded stainless steel wire clothand are capable of filtering out particles larger than 40
microns (Figure 4-3).
In addition to the mainline filters, most fuel systems
will incorporate several other filtering elements. They
may be located in the fuel tank, fuel control, fuel nozzles,
and any other place deemed desirable by the designer.
Pressurizing and Drain (Dump) Valves
The pressurizing and drain valve prevents flow to thefuel nozzles until sufficient pressure is reached in the main
fuel control. Once pressure is attained, the servo assemblies compute the fuel-flow schedules. It also drains
the fuel manifold at engine shutdown to prevent postshutdown fires. But it keeps the upstream portion of thesystem primed to permit faster starts.
Flow Divider. The pressurizing and dump valve
used on some engines has a somewhat different function.
In addition to the draining or dumping function, this unit
also serves as a flow divider. At the beginning of an engine
start, the fuel control supplies a pressure signal to thepressurizing and dump valve. This causes the valve toclose the manifold drain and open a passage for fuel flow
to the engine. On engine shutdown, fuel flow is cut off
immediately by a valve in the fuel control. The pressure
signal drops, the dump valve opens, and fuel drains fromFM 1-506

4-3

FM 1-506

the manifold. The flow divider allows fuel to flow to theprimary and secondary manifolds depending on fuelpressure. (See Figure 4-3.)
Pressure-Operated Valve. Some manufacturers,
such as Allison and Lycoming, install a pressureoperated valve in the combustion chamber section.
When the pressure in the burners drops below a specified
minimum, usually a few pounds per square inch, this valve
will open and drain any residual fuel remaining after afalse start or normal shutdown. (Refer to Figure 4-4 tosee where this drain valve fits into the system.)

4-4

Fuel Heater
Gas turbine engine fuel systems are very susceptible tothe formation of ice in the fuel filters. When the fuel in theaircraft fuel tanks cools to 32°F or below, residual water m
the fuel tends to freeze when it contacts the filter screen.
A fuel heater operates as a heat exchanger to warm
the fuel. The heater can use engine bleed air, an air-toliquid exchanger, or an engine lubricating oil, a liquid-toliquid exchanger, as a source of heat.

FM 1-506

A fuel heater protects the engine fuel system fromice formation. However, should ice form, the heater canalso be used to thaw ice on the fuel screen.
In some installations the fuel filter is fitted with apressure-drop warning switch which illuminates a warning light on the cockpit instrument panel. If ice begins tocollect on the filter surface, the pressure across the filter
will slowly decrease. When the pressure reaches apredetermined value, the warning light flashes on.
Fuel deicing systems are designed to be used intermittently. The system may be controlled manually by aswitch in the cockpit or automatically using a thermostatic

4-5

FM 1-506

sensing element in the fuel heater to open or close the airor oil shutoff valve. An automatic fuel heater is illustrated
in Figure 4-5.

Simplex Nozzle Figure 4-6(A) illustrates a typicalsimplex nozzle. Its chief disadvantage is that it is unable
to provide a satisfactory spray pattern in bigger engines

Fuel Nozzles
On most gas turbine engines, fuel is introduced into
the combustion chamber through a fuel nozzle. Thisnozzle creates a highly atomized, accurately shaped sprayof fuel for rapid mixing and combustion with the primary
airstream under varying conditions of fuel and airflow.
Most engines use either the single (simplex) or the dual
(duplex) nozzle.

because of the large changes in fuel pressures andairflows.
Duplex Nozzle. The chief advantage of the duplex
nozzle is its ability to provide good fuel atomization andproper spray pattern at all rates of fuel delivery andairflow. At starting and low RPM and at low airflow, thespray angle needs to be fairly wide to increase the chancesof ignition and good mixing of fuel and air. At higher

4-6

FM 1-506

RPM and airflow, a narrow pattern is required to keep
the flame of combustion away from the walls of the
combustion chamber (Figure 4-6{B]). The small fuelflow used in idling is broken up into a fine spray after being
forced through a small outlet formed by the primaryholes. The secondary holes are larger but still provide afine spray at higher RPM because of the higher fuelpressure.

For the duplex nozzle to function, there must be adevice to separate the fuel into low- (primary) and high(secondary) pressure supplies. This flow divider maybe
incorporated in each nozzle, as with the single-entry
duplex type (Figure 4-6[C]), or a single-flow divider may
be used with the entire system (Figure 4-6[D]).
Single-entry duplex nozzles incorporating an internal flow divider require only a single fuel manifold

4-7

FM 1-506(Figure 4-7). Dual-entry fuel nozzles require a double
fuel namifold. Some dual fuel manifolds may not beapparent as such.

Extreme care must be exercised when cleaning,
repairing, or handling the nozzles. Even fingerprints on

The flow divider, whether self-contained in eachnozzle or installed in the manifold, is usually a springloaded valve set to open at a specific fuel pressure. Whenthe pressure is below this value, the flow divider directs
fuel to the primary manifold or nozzle orifice. Pressures
above this value cause the valve to open, and fuel isallowed to flow in both manifolds or nozzle orifices.

the metering parts may produce a fuel flow which is outof tolerance.

Most modern nozzles have their passages drilled at
an angle. The fuel is discharged with a swirling motion toprovide low axial air velocity and high flame speed. In
addition, an air shroud surrounding the nozzle cools thenozzle tip. It also improves combustion by retarding theaccumulation of carbon deposits on the face. The shroud
also provides some air for combustion and helps tocontain the flame in the center of the liner (Figure 4-8).

4-8

FM 1-506

Fuel Shutoff Valves
The engine fuel shutoff valve is installed in the main
fuel supply line or tank outlet to the engine. It is controlled from the pilot’s compartment. A fuel shutoff valve
is usually installed between the fuel control unit and thefuel nozzles. When the throttle is placed in the closed
position, this ensures positive shutoff of fuel to the engine.
T-700/701 FUEL SYSTEM
The fuel system consists of the fuel boost pump, fuelfilter, hydromechanical unit (HMU), and overspeed anddrain valve. Integral to the HMU are the high-pressurevane pump, variable geometry (VG) actuator, and compressor inlet temperature sensor. These fuel systemcomponents are mounted on the accessory gearbox(AGB) and oriented as shown in Figure 4-9. At a 100

percent gas generator speed (Ng) of 44,700 RPM, the fuelboost pump is driven at 10,678 RPM and the HMU at
9947 RPM. This fuel-oil heat exchanger is located in themetered fuel flow path to the engine between the HMU
discharge and the overspeed and drain valve inlet. Except for one external line between the HMU discharge(metered fuel) and the fuel-oil heat exchanger inlet passage, all the fuel transfer is through cored passages within
AGB. This includes the boost pump inlet flange (engineaircraft interface) to the overspeed and drain valve where
the fuel discharges into the primer and the main
manifolds.
A schematic diagram of the fuel system component
arrangement is shown in Figure 4-10. An aircraft fuelsystem is depicted to show the overall aircraft-engine
fuel-handling systems.

4-9

FM 1-506

Requirements
The primary requirements of the fuel system are —To control accurate and stable engine speed forsteady-state operations and to provide transient
control to achieve rapid power changes. Thissystem pumps, filters, and meters fuel in
response to power-available spindle (PAS)
position, load-demand spindle (LDS) position,
sensed engine variables, and torque motor input
from the electrical control unit.
To position the compressor variable stator vanes
throughout the engine operating range toachieve the required airflow and stall margin
performance by the compressor.
To provide a schedule for starting bleed valve.
To provide automatic start schedules from sea
level to 20,000 feet altitude.
To protect the engine against destructive gas
generator and power turbine overspeed.

4-10These requirements must be met over the full engine
operating envelope and environment.
Fuel Boost Pump
The fuel boost pump is mounted on the AGB
forward side (Figure 4-11). It is designed —To provide reliable suction feed from theaircraft fuel boost pump to the engine.
To minimize vulnerability and fire hazard bycontinuous operation with fuel pressure in thelines from tank to engine at a pressure less thanatmospheric.
To provide discharge pressure for high-pressurevane pump minimum inlet pressure requirements.
At 100 percent Ng (44,700 RPM) the fuel boost
pump operates at 10,678 RPM. Pump discharge pressure ranges from 45 to 90 psi at maximum continuous
speed and 20 psi minimum at ground idle speed.

FM 1-506

Pump Inlet
The pump inlet is the engine-airframe fuel interface.
The pump is mechanically designed as a cantilevered
pumping element on a rigid shaft running in two oillubricated sleeve bearings. Oil is supplied from the engine oil system through a face port mating with gearbox
forbearing lubrication. Oil and fuel are separated by two
dynamic carbon seals with a center vent to the engine
overboard drain manifold. Shaft splines are lubricatedby oil mist which is pumped through the splints by radialpumping holes in the pump quill shaft and the mating
engine gearbox shaft.
Pump Elements
The pumping elements are an ejector or jet pump, a
mixed-flow centrifugal inducer, and a radial-flow impeller with a flow path from impeller discharge to provide
ejector pump motive flow. This bypass flow is approximately equal to twice maximum engine flow atpump rated speed.

Fuel Filter
Refer to Figures 4-12 and 4-13 and refer back to 4-1.
The fuel filter provides 30-micron absolute filtration forengine fuel prior to entering the high-pressure valve
pump in the HMU. After passing through a core in theaccessory gearbox, engine fuel flows to the filter throughits inlet port from the engine boost pump discharge. The
flow then is directed across the 30-micron pleated, bariertype filter element (outside to inside direction), out thedischarge port through another gearbox core, and to the
HMU vane pump inlet.
The fuel differential pressure across the element issensed across a magnet assembly piston on both theimpending and actual bypass indicators. At 8-10 psi ofpressure differential, the impending button piston assembly moves, carrying the magnet with it. As the magnetof the impending indicator moves away from the button,
the magnetic field holding the red button in the inward
position is reduced, allowing the button to pop out aided
by a spring.
4-11

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4-12

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The electrical filter bypass sensor (or actual bypass
indicator) is activated by a pressure signal as the bypass
valve opens at 18-22 psi differential pressure. However,in this case the reduced magnetic field allows the button
return spring inside the microswitch to actuate the switch
button through the action of the pivoted stitch lever arm.
This provides a cockpit indication of filter bypass.
Impending indicator mechanisms are locked when
the wire and half-ball element and the magnet piston
assembly move towards the center of the filter. This
causes the half ball to drop and catch on the edge of theramp. This latching prevents the magnet piston assemblyfrom returning to its original position. In the impending
by-pass position, the button can be pushed in its recessed
position but will not remain since the magnet has not
returned to provide the necessary magnetic field. The
filter bypass sensor does not latch.
The impending bypass button cannot be reset until
the filter element and bowl assembly are removed and thereset piston pushed upwards. The reset piston upper
land then trips the half-spherical ball at the end of thelocking wire, allowing the spriig-loaded piston to returnthe magnet piston assembly to its original position.
The filter bypass valve is a poppet type. As the filter
differential pressure reaches 18-22 psi, the bypass valve
opens, and the sensor is activated for remote indicationof filter bypassing. The filter element is disposable when
loaded.
Hydromechanical Unit (HMU)The HMU is mounted on the aft center of the AGB(Figure 4-14). It provides fuel pumping, fuel metering
fuel flow computation, fuel pressurization, and fuelshutoff. It also provides gas generator speed control,
compressor VG scheduling and actuation, and anti-icing
and starting bleed valve actuation. The unit responds to
PAS input for fuel shutoff, start, ground idle, up to maximum permissible gas generator speed, vapor venting, andelectrical control unit override capability. The HMU
also responds to an externally supplied load demand
input via the LDS which is proportional to the power
absorber load. This initially and directly coordinates gas
generator speed and power to closely approximate thepower required for the rotor or shaft power absorber.
The HMU then responds to input from an electricalcontrol unit (ECU) via the HMU torque motor. This
trims gas generator speed as directed for both power
turbine speed control and turbine temperature limiting
for more exact load share control. In the event of an ECU
failure, the HMU has the capability to mechanicallydeactivate the ECU. It also vents the unit case to the
overboard drain in the event of excessive air or vapor at
the inlet because of overtravel in the PAS. The HMU
responds to sensed engine parameters (T2, P3 and N3)
which influence fuel flow and VG position.
The HMU provides three major engine functions:fuel pumping fuel metering and VG positioning.
Fuel Pumping System (High Pressure). The main
fuel pump is housed within the HMU. It is a vane-type,pressure-balanced pump which is self-priming,
contamination-resistant, and interchangeable (at depotlevel maintenance). The gearbox drives the pump splineat 9947 RPM at 100 percent Ng The opposite end of thedrive shaft serves as the rotational input to the HMU.
The vane pump design was specifically developed withtungsten carbide vanes, pressure plates, and cam ring forcontamination resistance.
The vane inlet pressure is internally boosted above
engine boost discharge pressure by a fuel injector andinjector bypass valve. This ensures positive charging of
the vane element under all normal operating conditions
and allows limited operation with a failed engine boost
pump. The injector bypass valve maintains a differentailpressure across the injector of 12.5-135 psi. This results
in a vane pump innlet pressure up to 30 psi above boost
pump discharge.
A spring-loaded, high-pressure relief valve on thevane pump discharge prevents system overpressurization
by limiting vane pump differential pressure to 832 psi.
Fuel Metering System. The HMU fuel metering system controls fuel flow to the engine. The components inthis system are the —High-pressure fuel pump.
Servo supply filter.
Metering valve.
Pressure-regulating valve.
Shutoff valve.
Minimum pressure valve.
High-pressure fuel from the pump passes around theoutside of the servo supply wash filter and goes to themetering valve. The metering valve, in conduction with thepressure-regulating valve, meters required engine flow
proportionally to the metering valve area and bypasses
excess fuel through the PRV back to the pump inlet.
Variable Geometry Positioning System. A stator vane
actuator is an integral part of the HMU. It provides
positioning of the variable stator vanes and the anti-ice

4-13

FM 1-506

and the start bleed valve for efficient engine operation
throughout the engine speed (Ng ) range. A servo piston
within the HMU provides an external output (force andposition) to set the engine VG as a function of Ng and T2
This schedule is contoured on the Ng speed servo 3D
cam. Motion of the schedule-follower lever moves a link
pivoted on the actuator piston. This action strokes adouble-acting, spring-loaded, rotating pilot valve. The
valve then provides high- and low-fuel pressure signals to
opposite sides for the VG actuator piston. Motion of theactuator returns the pilot valve to a null position, which
cm-responds to the input position being scheduled. The
actuator provides minimum force output of 280 pounds
and is positioned as a function of corrected speed.

Overspeed and Drain Valve. The T7-GE-701 valve
controls the sequencing of fuel between the main
manifold and the primer manifold (Figure 4-15). The
valve also provides shutoff, drain, primer manifold purge
functions, and Np overspeed protection. The assembly
consists of the following components
Shutoff and drain valve.
Pressurizing valve.
Wash flow filter.
Bypass valve (Np overspeed).
Fuel servo solenoid valve (Np overspend).

4-14

Engine Start
With the engine shut down, the pressure in theoverspeed and drain valve is equal to ambient pressure,and the inlet and pressurizing valve is spring-loaded
closed. When the engine is cranked over and the PAS isadvanced to the stopcock in the HMU, the HMU delivers,
metered flow to the overspeed drain valve inlet. This
metered flow passes through awash-flow filter, and linepressure builds up until the inlet and pressurizing valve
opens. Fuel boost pressure is directed to the top of a
selector valve. Higher-pressure Wf passes through theopen selector valve and flows to the 12 fuel injectors forengine lightoff and normal operation.

Engine Shutdown
When the PAS is moved to the OFF position, the
HMU is stopcocked, fuel flow to the overspeed and drain
valve is stopped, and the inlet and pressurization valve isforced down by spring action. This opens the main fuelline in the overspeed and drain valve to overboard drainpressure. The residual P3 pressure in the combuster
forces the fuel injectors and main manifold back throughthe overspeed and drain valve and out the overboard
drain line.

FM 1-506

Power Turbine (Ng) Overspeed Protection
The turbine overspeed function is designed to divert
fuel flow from the combustor and flameout the engine toprotect against destructive Np overspeed (Figure 4-16).
When Np reaches 25,000 RPM (119.6 percent), thesolenoid valve in the ODV is energized and opens a ballvalve. This causes a bleed-off of equalizing fuel pressure
on the bottom part of the selector valve. The high pressure W4 causes the selectro valve to move down and all
W4 to bypass back to the HMU inlet. The instantaneous
loss of W4 to the combustor causes the engine toflameout.
Fuel Injectors
Twelve fuel injectors, installed in the midframe,receive fuel from the main fuel manifold and supply it to
the combustion liner swirl cup (swirler) subassemblies(Figure 4-17). The swirlers consist of counter-rotatingflow primary and secondary vanes with venturi sleeves
between the two radial inflow swirler vanes. Fuel fromthe injector is forced into these sleeves, and the counterrotating flow creates shear layers which break up the fuelfilm into finely atomized sprays.
The injector is a simplex nozzle with a single spinchamber fed by two spin holes. The swirled fuel from thespin holes exits through an orifice and discharges onto a
30° primary cone. This cone flares outward at the exit.
The contoured exit provides a good-quality fuel spray atvery low fuel pressure in the starting regime. The primarycone is housed in an air shroud which has six air holes.
These six air holes feed air to augments the atomizationof the fuel. This hydraulic design is simple; contains novalves, flow dividers, or secondary orifices and is purged
by compressor discharge air during shutdown to prevent
plugging.
The injector has a military standard fitting at the inlet
with a last-chance, 0.009-inch diameter mesh screen toprotect the orifice spin slots. A locating pin in the mounting flange assures correct orientation of the injector
within the midframe casing. It is not necessary to remove
the injectors before removing or replacing the combustion liner.
FUEL CONTROL UNIT
Types
Modern fuel controls can be divided into two basic
groups: hydromechanical and electronic.
Hydromechanical controls are used most often. Fuelcontrols are extremely complex devices composed of
speed governors, servo systems and feedback loops, valves,

4-15

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4-16

FM 1-506

metering systems and various sensing mechanisms.
Electronic fuel controls contain thermocouples,
amplifiers, relays, electrical servo system switches, andsolenoids. The discussion of fuel-control theory will address mainly the hydromechanical type. The simplest
control is a plain metering valve to regulate fuel flow tothe engine. This type of control could be installed on anengine used for thrust or gas generation.
Fuel controls may sense some or all of the following
engine operating variables:
Pilots demands.
Compressor inlet temperature.
Compressor discharge pressure.
Burner pressure.
Compressor inlet pressure.
RPM.
Turbine temperature.
Some refinements might include a —Pump to pressurize the fuel.
Shutoff valve to stop fuel flow.
Relief valve to protect the control when the
shutoff valve is closed.
Minimum fuel control adjustment to prevent
complete stoppage of fuel by the metering valve.
Turboprop and Turboshaft Engine Fuel Controls
Like fuel controls for turbojet and turbofan engines,
the fuel control for a turboprop or a turboshaft engine
receives a signal from the pilot for a given level of power.
The control then takes certain variables into consideration. It adjusts the engine fuel flow to provide the desired
power without exceeding the RPM and turbine inlet
temperature limitations of the engine. But the turboprop
or turboshaft engine control system has an additional job
to do that is not shared by its turbojet and turbofan
counterparts. It must control the speed of the propeller
or the free turbine, and it usually governs the pitch angleof the propeller blades.
Many turboprop and turboshaft engines in production today are the free turbine type. Engines of this kindact principally as gas generators to furnish high-velocity
gases that drive a freely rotating turbine mounted in theexhaust gas stream. The free turbine turns the propeller
through a system of reduction gears. If the engine is aturboshaft model, the free turbine rotates (also throughreduction gears) a helicopter rotor or powers a machine.

The General Electric T-700 and the Lycoming T-55-L712 fuel control system follow.
T-700 CONTROL SYSTEM
The engine control system incorporates all control
units necessary for complete control of the engine. The
system provides for the more common functions of fuelhandling, computation, compressor bleed and VG control, power modulation for rotor speed control, and overspeed protection. The system also incorporates control
features for torque matching of multiple engine installations and overtemperature protection.
The T-700 control system was designed for simple
operation requiring a low level of pilot attention. The
system performs many of the controlling functions
formerly performed by the pilot. This has been done byproviding —Isochronous power turbine and helicopter rotor(Np/Nr) governing.
Automatic load sharing.
Automatic limiting of power turbine inlet
temperature.
Rapid engine transient response through collective compensation.
Automatic starting
Contingency power.
Basic system operation is governed through the interaction of the electrical and hydromechanical control
units. In general, the HMU provides for gas generator
control in the areas of acceleration limiting, stall andflameout protection, gas generator speed limiting rapid
response to power demands, and VG actuation. The
ECU trims the HMU to satisfy the requirements of theload to maintain rotor speed, regulate load sharing, andlimit engine power turbine inlet temperature.
Component Arrangement and Interconnection
Metering of fuel to the engine and basic engine
control computations are performed in thehydromechanical control unit (Figure 4-18). The electrical and hydromechanical control units compute the fuelquantity to satisfy power requirements of the engine. The
fuel and control system contains the following components:
Hydromechanical control unit which includes a9947 RPM (at 100 percent of Ng) high-pressurevane pump package and a closed-loop VG
servo-acutator. The basic unit meters the required fuel to the engine as a function of Ng T2,
P3, an electrical trimming signal from the
4-17

FM 1-506

engine-mounted electrical control unit, andcockpit inputs of PAS and LDS.
Electrical control unit which resets thehydromechanical control unit within acceptableengine limits to maintain isochronous Np— The required signals to an engine-mounted
recorder to provide a record of engine
operational history.
— The desired cockpit signals needed forproper engine operation.

governing while automatically limiting TGT.
The unit exchanges torque signals to provide
automatic load sharing for multiengine use.
The ECU provides —For the T-700-GE-701/701A engines in a oneengine inoperative (OEI) condition, the automatic
contingency power switch, located in the ECU,
resets the temperature limit amplifier to allow for
more power is required for the operating engine.
Figure 4-19 lists the functional split between the twounits mentioned above.

— A separate signal to the overspeed and drain
valve for NP overspeed protection.

4-18

FM 1-506

Fully integrated operation of the engine results fromthree inputs:
Power-available spindle. This input has afunction similar to the throttle or power lever ofother engine controls. It mechanically stopcocks fuel and sets permissible gas generator
speeds for operating condtions from groundidle through intermediate power. Overtravel
beyond 120° to 130° actuates an electrical unit
interface lockout, used in the event of an electrical unit failure. It vents the hydromechanicalunit ease to overboard drain to facilitate primingof the fuel system.
Load-demand spindle. This input is a function of collective pitch. It provides collective
compensation reduce rotor transient droop.
Rotor speed referenee. A cockpit mounted
electric potentiometer sets desired power turbine speed.
Hydromechanical Unit Operation
The HMU operates as a conventional gas generator
power control when there is no input to the torque motor
and with the load demand spindle set for maximum load
(Figure 4-20). Engine schedules such as acceleration fuelflow and compressor VG position, are controlled by Ng
and T2. The acceleration (or stall) limit is immposed bya maximum can schedule of Wf/P3 versus Ng and T2. A

minimum flow schedule controls engine deceleration.
Maximum and minimum metering valve stops provide
absolute fuel-flow limits. An intermediate topping line
controls maximum Ng speed. The topping line also limits
excessive T4.5 in event of an electrical T4.5 control-system
failure. Since these features are incorporated in the
HMU, any power-available spindle motion will result insafe engine operation and will not cause engine damage.
Except for intentional stopcocking of the control, no
inadvertent shutdowns will occur from PAS motion.
A load-demand signal is introduced to the HMU
through the LDS. When the LDS is reduced for its
maximum setting with a reduction of aircraft rotor collective pitch setting the desired Ng is reset down from theprevailing PAS setting to provide immediate and accurate
gas generator response. This new setting is trimmed by
the ECU to satisfy the NP and load control functions
established by the ECU.
The PAS sets a maximum available gas generator
speed. The pilot usually sets this 120° PAS angle allowing
the gas generator speed to reach a value which gives
intermediate power. Through this collective pitch setting, the pilot positions the LDS to adjust available Ng topower level approximately equal to the rotor loaddemand power. The actual level of horsepower at 120°
PAS angle will be more than required by the helicopter.
This is intentionally done for two reasons:4-19

FM 1-506

Fail-safe to high power. By placing theschedule high, the torque motor is designed toreduce the schedule to the desired level when it
is energized. Because loss of torque motor current causes the schedule to return to thehighest level, this provides failure protection toa high power. A schedule that is biased high due
to electrical failure does not cause power limiting. In the event of an electrical system failure,
it can be manually retarded to a more desirable
level using the PAS. With all engine protection
functions in the HMU, no engine damage orstall can occur during or following any loss ofelectrical signal to the torque motor.

4-20

Power available with one engine out. The
schedule is placed high enough that, in the eventof failure of one engine, the remaining engine’s
gas generator can increase power enough (automatic contingency power) to carry the load at
the given LDS (collective pitch) setting.
Without this arrangement of schedules, a singleengine failure could not be compensated for by
the remaining engine and would result in a significant loss of available power.
Electrical Control Unit Operation
The ECU, powered from an engine-mounted alternator (Figures 4-21, 4-22)—

FM 1-506

4-21

FM 1-506

Compares actual Np to a reference Np to compute a speed error input signal for use in electrical control computation.
Compares actual torque of the engine to torqueof the other engine for load sharing. If theengine torque (Q1) is less than that of the otherengine (Q2), supplies a torque error signal to thecomputer. No error signal appears if Q 1 isgreater then Q2. This arrangement (frequently
reffered to as a match-the-maxixmun principle)
prevents the load-sharing circuit from causing
a good engine to reduce power to match a
power-limited engine. The torque error signal
is added to the Np speed reference signal andcan increase up to 3 percent over the selected
Np.
Compares measured engine TGT to a fixedreference TGT. When (and only when)temperature is above the reference, develops asignal to reduce fuel flow to correct the overtemperature.
Compares the combined speed error and loadsharing signal to the overtemperature signal in
the selector circuit of the ECU. Selects thesignal calling for the lowest Ng and passes it to
the HMU to reduce gas generator power as
required.
Sends an electrical output from the selector
drives to a torque motor in the HMU. The
torque motor can shift the load demandscheduled as if the PAS were moved in a
reduced power demand direction. The torquemotor can retard Ng no further than ground idle
and can only allow power to increase to the PAS
setting. The ECU may be overridden at any
time by advancing PAS beyond the normal(120°) flight position to the 130° stop. PAS may
then be retarded to trim the engine manually.
Note that this has no effect on the Np overspeed
system, which remains fully operational.
Takeoff and Climb
Before takeoff, the PAS lever is advanced fromground idle to 120°, which is the normal flight position.
This allows the rotor head to accelerate to 100 percent
Np/Nr with the collective still in flat pitch. Ng will rise as
the PAS is advanecd but will cut back after the Np governing speed of 100 percent is reached. (Np governing canbe selected at 95-101 percent, but 100 percent is considered standard.)

4-22

To take off, collective pitch is increased, increasing
the torque load on the power turbine. Simultaneously,
the LDS rotates, calling for an increase in gas generator
speed to keep Np/Nr from falling below 100 percent. The
ECUs of each engine perform a fine trim of fuel flow tomatch torques and trim Np/Nr to 100 percent.
As Ng inereases, the HMU schedules the bleed valve
closed and the variable stators open to increase airflow
through the combustor and turbine.
If collective pitch is increased to a very high angle,
TGT may approach the temperature limiting value.
When this occurs, the ECU prevents any further increasein fuel to the engine. If torque load is increased further,
Np/Nr will droop below 100 percent because power turbine governing must be sacrificed to protect the engine
against overtemperature.
Cruise and Descent
At the end of the climb segment, less rotor lift isrequired and collective pitch is recked. The LDS
rotates to reduce fuel flow and Ng; the variable stators willclose slightly to optimize part-power fuel consumption
and preserve stall margin.
Again, ECUs trim fuel flow. Upon entering a descent, the same sequence of events reduces Ng to thepoint that the bleed valve maybe opened. If the collective
pitch is fully lowered, autorotation (power-off descent) isreached, and the torque drops to zero; overrunningclutches in the airframe transmission preclude a negative
torque situation. Once the engines are uncoupled fromthe rotor, Nr is free to accelerate above 100 percent; andboth engines will continue to govern Np at 100 percent.
Nr and Np rotor speeds are held in reserve topick up the rotor load when collective is again increased.
This condition (100 percent Np zero torque) is known as
flight idle.
T-55-L-llE/T55-L-712 FUEL CONTROL SYSTEM
The fuel control for the Lycoming T55-L-712 free
power turbine engine is a hydromechanical type designated by Hamilton Standard as the JFC 31-22. It consistsof the following main units:
Single-element fuel pump.
Gas producer speed governor.
Power turbine speed governor.
Acceleration and deceleration control.
Air bleed signal mechanism.
Maximum flow limiter.
Fuel shutoff valve.
Temperature compensator.

FM 1-506

Figure 4-23 schematically describes the fuel control
system. Fuel control can be divided on a functional basis
into two sections: the flow control section, consisting of
the valving which meters engine fuel flow, and the computer section. The computer section includes the elements which schedule position of the flow control section
metering valves as a function of the control input signals.
The computer section also signals the closure of thecompressor bleed valve.
This computer operates primarily on four engine
parameters gas producer speed (Ng), power turbine
speed (Np), compressor discharge total pressure (Pt3),
and compressor inlet total temperature (Tt2) for a given
power Iever position and power turbine speed-setting
lever position. The ratio of engine fuel flow to compressor discharge total pressure is the manipulated variable
used to control engine power output. This variable isbiased by Tt2 during acceleration and Ng topping.
The position of the engine fuel flow metering valve
is established by a multiplication of two positions representing compressor discharge pressure (P3) and fuel flow(Wf). The (Wf) engine fuel flow is made proportional tothe metering valve position by maintaining a constantdifferential pressure across the valve flow area, which
varies linearly with valve position. This constant pressure
in maintained by the pressure regulating valve. This valve
is positioned as necessaryto maintain a 40-psi differential
across the metering area. To accomplish this, all pump
output not required to operate servos and supply engine
flow requirements is bypassed back to the pump inlet.
The P3 servo (motor) is positioned by directing
compressor discharge (P3) to a bellows, which positions
the servo directly proportional to P3 air pressure. The
servo is connected to the P3 ramp, which moves proportional to servo inputs. The ramp works directly against
the ratio servo and in turn positions the metering valve.
The ratio servo maybe positioned by anyone of four
inputs: the Np governing linkage, the Ng governing
linkage, acceleration limiting linkage, or the maximum
flow limiter. The inputs, which schedule the lowest fuelflow, determine the position of the metering valve. The
ratio servo and the Ng servo provide an input to thecompressor bleed air signal mechanism.

The Np governor linkage input to the ratio servo iscontrolled by the Np servo piston whose position isestablished by a flyball-operated pilot valve. It functions
as a droop-type governor, limiting fuel flow proportionalto speed over its operating range, which is set by thepower turbine speed setting lever position. The Np
governing linkage input to the ratio servo is biased by Ng
speed such that Np governing droop line position increases in speed as Ng speed decreases.
The Ng governing linkage input to the ratio servo iscontrolled by the Ng servo piston whose position is established in the same manner as the Np servo. It functions
as a droop-type governor, limiting the ratio servo inversely proportional to Ng speed over its operating range.
Its operating range (ground idle to maximum speed) is
established by the power and shutoff lever position. This
linkage is also biased by T2 through the speed reset
linkage.
The acceleration limiting linkage input to the ratio
servo is controlled by a contour on the 3D cam. The 3D
cam is positioned longitudinally by the Ng servo piston
and rotated by the T2 servo piston. This provides aunique cam positioned for each set of NG speeds and T2
conditions. The acceleration finding cam contour
schedules the maximum allowable fuel of low P3 ratio at
each speed and temperature setting. The 3D cam also
positions the speed reset linkage through the speed reset
contour and provides an input to the bleed air signal
mechanism.
The position of the T2 servo is established by theliquid-filled bellows which positions the servo directly
proportional to the compressor inlet temperature.
The maximum flow limiter input to the ratio servo
limits the engine shaft torque by limiting fuel flow. Thisis done by positioning the ratio servo as necessary to
reduce fuel flow at a safe value.
The control computer section also provides apneumatic signal to the bleed air acutator which is used
to open or close the bled air during appropriate phasesof engine operation.

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4-24

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CHAPTER 5

LUBRICATING SYSTEMS

Although the oil system of the modern gas turbine
engine is varied in design and plumbing most have units
which perform similar functions. In most cases a pressure
pump or system furnishes oil to the engine to be lubricated and cooled. A scavenging system returns the oil tothe tank for reuse. The problem of overheating is more
severe after the engine has stopped than while it is running. Oil flow which would normally have cooled thebearings has stopped. Heat stored in the. turbine wheel
will raise the bearing temperature much higher than thatreached during operation. Most systems will include aheat exchanger (air or fuel) to cool the oil. Many aredesigned with pressurized sumps. Some incorporate apressurized oil tank. This ensures a constant head pressure to the pressure- lubrication pump to prevent pumpcavitation at high altitude.
Oil consumption in a gas turbine engine is low compared to that in a reciprocating engine of equal power.
Oil consumption on the turbine engine is affected by theefficiency of the seals. However, oil can be lost throughinternal leakage and on some engines by malfunction of
the pressurizing or venting system. Oil scaling is very
important in a jet engine. Any wetting of the blades orvanes by oil vapor will encourage the accumulation ofdust and dirt. A dirty blade or vane represents highfriction-to-airflow. This decreases engine efficiency, andresults in a noticeable decrease in thrust or increase infuel consumption. Since oil consumption is so low, oiltanks can be made relatively small. This causes a
decrease in weight and storage problems. Tanks may
have capacities ranging from l/2 to 8 gallons. Systempressures may vary from 15 psig at idle to 200 psig duringcold starts. Normal operating pressures 0and bulk
temperatures are about 50 to 100 psig and 200 F, respectively.
GENERAL
In general the parts to be lubricated and cooled
include the main bearings and accessory drive gears andthe propeller gearing in the turboprop. This represents
again in gas turbine engine lubrication simplicity over thecomplex oil system of the reciprocating engine. The main

rotating unit can be carried by only a few bearings. In a
piston power plant there are hundreds more moving
parts to be lubricated. On some turbine engines the oil
may also be uscd–
To operate the servo mechanism of some fuelcontrols.
To control the position of the variable area
exhaust-nozzle vanes.
To operate the thrust reverser.
Because each bearing in the engine receives its oilfrom a metered or calibrated orifice, the system isgenerally known as the calibrated type. With a few exceptions the lubricating system used on the modem turbine engine is of the dry-sump variety. However, someturbine engines are equipped with a combination dryand wet-type lubrication system. Wet-sump engines
store the lubricating oil in the engine proper. Dry-sumpengines utilize an external tank usually mounted on ornear the engine. Although this chapter addresses drysump systems, an example of the wet-sump design can beseen in the Solar International T-62 engine. In this engine
the oil reservoir is an integral part of the accessory-drive
gear case. An example of a combination dry- and wetsump lubrication can be found in the Lycoming T-55series engines.
TURBINE ENGINE DRY-SUMP LUBRICATION
In a turbine dry-sump lubrication system, the oil
supply is carried in a tank mounted externally on or near
the engine. With this type of system, a larger oil supply
can be carried and the oil temperature can be controlled
An oil cooler usually is included in a dry-sump oil system(Figure 5-l). This cooler may be air-cooled or fuelcooled. The dry-sump oil system allows the axial-flow
engines to retain their comparatively small diameter.
This is done by designing the oil tank and the oil cooler
to conform to the design of the engine.
The following component descriptions include mostof those found in the various turbine lubrication systems.
However, not all of these components will be found in any
one system.
5-1

FM 1-506

The dry-sump systems use an oil tank which contains
most of the oil supply. However, a small sump usually is
included on the engine to hold a supply of oil for an
emergency system. The dry-sump system usually contains –Oil pump.
Scavenge and pressure inlet strainers.
Scavenge return connection.
Pressure outlet ports.
Oil falter.
Mounting bosses for the oil pressure transmitter.
Temperature bulb connections.

5-2

A typical oil tank is shown in Figure 5-2. It is
designed to furnish a constant supply of oil to the engine.
This is done by a swivel outlet assembly mounted inside
the tank a horizontal baffle mounted in the center of thetank, two flapper check valves mounted on the baffle, anda positive-vent system.
The swivel outlet fitting is controlled by a weighted
end, which is free to swing below the baffle. The flapper
valves in the baffle are normally open. They close only
when the oil in the bottom of the tank rushes to the topof the tank during deceleration. This traps the oil in thebottom of the tank where it is picked up by the swivel

FM 1-506

fitting A sump drain is located in the bottom of the tank.
The airspace is vented at all times.
All oil tanks have expansion space. This allows foroil expansion after heat is absorbed from the bearings andgears and after the oil foams after circulating through thesystem. Some tanks also incorporate a deaerator tray.
The tray separates air from the oil returned to the top of
the tank by the scavenger system. Usually these
deaerators are the “can” type in which oil enters a tangent. The air released is carried out through the ventsystem in the top of the tank. Inmost oil tanks a pressurebuildup is desired within the tank. This assures a positive
flow of oil to the oil pump inlet. This pressure buildup
is made possible by running the vent line through an
adjustable check-relief valve. The check-relief valve
normally is set to relieve at about 4 psi pressure on the oilpump inlet.
There is little need for an oil-dilution system. If the
air temperature is abnorrnally low, the oil may be changed

engines are in operation, there are few engines using a
wet-sump type of oil system.
The components of a wet-sump system are similar
to many of a dry-sump system. The oil reservoir locationis the major difference.
The reservoir for the wet-sump oil system may be theaccessory gear case, which consists of the accessory gearcasing and the front compressor bearing support casing.
Or it may be a sump mounted on the bottom of theaccessory case. Regardless of configuration reservoirs
for wet-sump systems are an integral part of the engine
and contain the bulk of the engine oil supply.
The following components are included in the wetsump reservior:
A bayonet-type gage indicates the oil level in thesump.
Two or more finger strainers (filters) are
inserted in the accessory case for straining

to a lighter grade. Some engines may provide for the
installation of an immersion-type oil heater.

pressure and scavenged oil before it leaves orenters the sump. These strainers aid the main
oil strainer.
A vent or breather equalizes pressure within theaccessory casing.
A magnetic drain plug may be provided to drain
the oil and to trap any ferrous metal particles in

TURBINE ENGINE WET-SUMP LUBRICATION
In some engines the lubrication system is the wetsump type. Because only a few models of centrifugal-flow

5-3

FM 1-506

the oil. This plug should always be examinedclosely during inspections. The presence of
metal particles may indicate gear or bearing
failure.
A temperature bulb and an oil pressure fitting
may be provided.
This system is typical of all engines using a wet-sumplubrication system. The bearing and drive gears in theaccessory drive casing are lubricated by a splash system.
The oil for the remaining points of lubrication leaves thepump under pressure. It passes through a falter to jet
nozzles that direct the oil into the rotor bearings andcouplings. Most wet-sump pressure systems arevariable-pressure systems in which the pump outlet pressure depends on the engine RPM.
The scavenged oil is returned to the reservior (sump)by gravity and pump suction. Oil from the front compressor bearing in the accessory-drive coupling shaft drains
directly into the reservoir. Oil from the turbine coupling
and the remaining rotor shaft bearings drains into a sump.
The oil is then pumped by the scavenge element througha finger screen into the reservoir.
OIL SYSTEM COMPONENTS
The oil system components used on gas turbine engines are–Tanks.
Pressure pumps.
Scavenger pumps.
Filters.
Oil coolers.
Relief valves.
Breathers and pressurizing components.
Pressure and temperature gages and warning
lights.
Temperature-regulating valves.
Oil-jet nozzle.
Fittings, valves, and plumbing.
Chip detectors.
..
Not all of the units will be found in the oil system of
any one engine. But a majority of the parts listed will befound in most engines.
Oil Tanks
Tanks can be either an airframe or enginemanufacturer-supplied unit. Usually constructed ofwelded sheet aluminum or steel, it provides a storage
place for the oil. Inmost engines the tank is pressurized
5-4

to ensure a constant supply of oil to the pressure pump.
The tank can contain–
Venting system.
Deaerator to separate entrained air from the oil.
Oil level transmitter or dipstick.
Rigid or flexible oil pickup.
Coarse mesh screens.
Various oil and air inlets and outlets.
Pressure Pumps
Both gear- and Gerotor-type pumps are used in the
lubricating system of the turbine engine. The gear-typepump consists of a driving and a driven gear. The engineaccessory section drives the rotation of the pump. Rotation causes the oil to pass around the outside of the gears
in pockets formed by the gear teeth and the pump casing.
The pressure developed is proportional to engine RPM
up to the time the relief valve opens. After that any
further increase in engine speed will not result in an oilpressure increase. The relief valve may be located in thepump housing or elsewhere in the pressure system forboth types of pumps.
The Gerotor pump has two moving parts: an innertoothed element meshing with an outer-toothed element.
The inner element has one less tooth than the outer. The
missing tooth provides a chamber to move the fluid fromthe intake to the discharge port. Both elements aremounted eccentrically to one another on the same shaft.
Scavenger Pumps
These pumps are similar to the pressure pumps but
have a much larger total capacity. An engine is generally
provided with several scavenger pumps to drain oil fromvarious parts of the engine. Often one or two of the
scavenger elements are incorporated in the same housingas the pressure pump (Figure 5-3). Different capacities
can be provided for each system despite the common
driving shaft speed. This is accomplished by varying thediameter or thickness of the gears to vary the volume of
the tooth chamber. A vane-type pump may sometimes
be usedOil Filters and Screens or Strainers
To prevent foreign matter from reaching internal
parts of the engine, falters and screens or stainers areprovided in the engine lubricating system. The threebasic types of oil falters for the jet engine are the cartridge
screen-disq and screen (Figures 5-4, 5-5 and 5-6). The
cartridge falter is most commonly used and must be
replaced periodically. The other two can be cleaned andFM 1-506

reused. In the screen-disc falter there are a series ofcircular screen-type falters. Each falter is comprised of
two layers of mesh forming a chamber between meshlayers. The filters are mounted on a common tube andarranged to provide
a space between each circular.
element. Lube oil passes through the circular meshelements and into the chamber between the two layers ofmesh. This chamber is ported to the center of a common
tube which directs oil out of the falter. Screens or strainers
are placed at pressure oil inlets to bearings in the engine.
This aids in preventing foreign matter from reaching thebearings.

5-5

FM 1-506

attract ferrous material (metal with iron as its basic element) which may come from inside the engine. This
ferrous material builds up until it bridges a gap. Whenever there is a requirement, the chip detectors may becollected and analyzed to determine the condition of theengine. Most engines utilize an electrical chip detector,
located in the scavenger pump housing or in the aceessory
gearbox. Should the engine oil become contaminatedwith metal particles, the detector will catch some of them.
This causes the warning light on the caution panel tocome on.
Tubing, Hose, and Fittings
Tubing hose, and fittings are used throughout the
lubricating system. Their purpose is to connect apart intoa system or to connect one part to another to complete asystem.
To allow for oil flow in the event of filter blockage,
all falters incorporate a bypass or relief valve as part of thefalter or in the oil passages. When the pressure differential reaches a specified value (about 15 to 20 psi), the valve
opens and allows oil to bypass the filter. Some falters
incorporate a check valve. This prevents reverse flow or
flow through the system when the engine is stopped
Filtering characteristics vary, but most falters will stop
particles of approximately 50 microns.
Magnetic Chip Detector
One or more magnetic chip detectors are installed
on gas turbine engines. They are used to detect and

5-6

Oil Pressure Indicating System
In a typical engine oil pressure indicating system theindicator receives inlet oil pressure indications from theoil pressure transmitter and provides readings in pounds
per square inch Electrical power for oil pressure indicator and transmitter operation is supplied by the 28volt AC system.
Oil-Pressure-Low Caution Light
Most gas turbine engine lubricating systems incorporate an engine oil-pressure-low caution light warning
device into the system for safety purposes. The light isconnected to a low-pressure switch. When pressure
drops below a safe limit, the switch closes an electricalFM 1-506

circuit causing the caution light to bum. Power is supplied by the 28-volt DC system.
Oil Temperature Indicating System
In a typical engine oil temperature indicating system,
the indicator is connected to and receives temperature
indications from an electrical resistance-type thermocouple or thermobulb. These are located in the pressure pump oil inlet side to the engine. Power to operate
this circuit is supplied by the 28-volt DC system.
Oil Coolers
The oil cooler is used to reduce oil temperature bytransmitting heat from the oil to another fluid usually
fuel. Since the fuel flow through the cooler is muchgreater than the oil flow, the fuel is able to absorb aconsiderable amount of heat. This reduces the size andweight of the cooler. Thermostatic or pressure-sensitive
valves control the oil temperature by determining
whether the oil passes through or bypasses the cooler. Oilcoolers are also cooled by air forced through them by ablower/fan.
Breathers and Pressurizing Systems
Internal oil leakage is kept to a minimum by pressurizing the bearing sump areas with air that is bled off
the compressor (Figure 5-7). The airflow into the sumpminimizes oil leakage across the seals in the reverse
direction.
The oil scavenge pumps exceed the capacity of the
lubrication pressure pump They are capable of handling considerably more oil than actually exists in the
bearing sumps and gearboxes. Because the pumps area
constant-displacement type, they make up for the lack ofoil by pumping air from the sumps. Large quantities of
air are delivered to the oil tank. Sump and tank pressures
are maintained close to one another by a line whichconnects the two. If the sump pressure exceeds the tankpressure, the sump vent check valve opens, allowing theexcess sump air to enter the oil tank. The valve allows
flow only into the tank; oil or tank vapors cannot back up
into the sump areas. Tank pressure is maintained little
above ambient.
The scavenge pumps and sump-vent check valve
funtions result in relatively low air pressure in the sumps
and gearboxes. These low internal sump pressures allow
air to flow across the oil seals into the sumps. This airflow
minimizes lube oil leakage across the seals. For thisreason it is necessary to maintain sump pressures low
enough to ensure seal-air leakage into the sumps. Under
some conditions, the ability of the scavenge pumps topump air forms a pressure low enough to cavitate thepumps or cause the sump to collapse. Under otherconditions, too much air can enter the sump throughworn seals.
If the seal leakage is not sufficient to maintainproper internal pressure, check valves in the sump andtank pressurizing valves open and allow ambient air toenter the system. Inadequate internal sump and gearbox
pressure may be caused by seal leakage. If that occurs,
air flows from the sumps, through the sump-vent checkvalve, the oil tank, the tank and sump pressurizing valvesto the atmosphere. Tank pressure is always maintained a
few pounds above ambient pressure by the sump and tankpressurizing valve.
The following adresses two types of lubrication systems used in the Army today the General Electric T-701
turboshaft engine and the International/Solar T-62-seriesengine.

TYPICAL OIL SYSTEM FOR T-701
The lubrication system in the T-700-GE-701 engine
distributes oil to all lubricated parts (Figure 5-8). In
emergencies it supplies an air-oil mist to the main
shaft bearings in the A- and B-sumps. The system is
5-7

FM 1-506

a self-contained, recirculating dry-sump system. It consists of the following subsystems and components:
Oil supply and scavenge pump.
Seal pressurization and sump venting
Emergency lube system.
Oil filtration and condition monitoring.
Tank and air-oil cooler.
Oil cooler.
Oil pressure monitoring.
Cold oil-relief and cooler-bypas valves.
Chip detector.
Integral accessory gearbox

installed on each side of the tank. A coarse pickup screen
located near the tank bottom keeps sizable debris fromentering the lube supply pump inlet. A drain plug islocated at the bottom of the tank.
Oil from the pickup screen enters a east passage inthe mainframe. It is then conducted to the top of theengine to a point beneath the lube supply pump. A shortconnector tube transfers the oil from the mainframe tothe accessory gearbox pump inlet port. The connector
tube contains a domed, coarse-debris screen. The screen
keeps foreign objects out of the passage when the accessory module is not installed on the mainframe. Oil flows
through the connector tube to the pump inlet. There it

Lube Supply System
The oil tank, integral with the mainframe, holds
approximately 7.3 quarts of oil (Figure 5-9). This is asufficient quantity to lubricate the required engine parts
without an external oil supply. The tank is filled using a
3-inch, gravity-fill port on the right-hand side. Visualindication of oil level is supplied by a fluid level indicator

enters the pump tangentially in alignment with pumprotational direction.
The lube supply pump, a Geroter-type pumping
element assembly,. is comprised of an inner and outerelement. The element assembly is located adjacent to thedrive spline end of the pump. Six scavenge elements arealso located in tandem on the common drive shaft. The

5-8

FM 1-506

stack of pump elements is retained in a cast tubular hosing
having an integral end plate. The complete pump slides
into a precision bore in the gearbox casing. Oil from the
supply pump flows to the lube filter inlet and through theis set for 120-180 psid and reset is 115 psid minimum.
When open, apart of the lube flow is discharged into the
AGB where churning in the gears will assist in reducing
warm-up time.

falter, a 3-micron falter element. Oil flow passes fromoutside to inside of the falter element. It then passes
through the open bore of the bypass valve and into thegearbox outlet passage. Bypass valve opening occurs
when filter differential pressure unseats a spring-loaded
poppet from its scat. The filter bowl contains an impending bypass warning button which will provide an indication for filter servicing. An electrical bypass sensor for
cockpit indication of falter bypass tits into an AGB boss
adjacent to the lube falter. A differential pressure of 60-80
psi across the falter will actuate this sensor.

Oil leaving the falter branches in three directions. It
goes to the to of the emergency oil reservoirs in the Aand B-sumps, the AGB, and C-sump jets.
Scavenge System. After the oil has lubricated andcooled the parts, the scavenge system returns it to the oiltank (Figure 5-10). In addition, fuel-oil and air-oil coolers
and a chip detector are located in the scavenge return
path.
Scavenge Inlet Screens. Each scavenge pump inlet
is fitted with a relatively coarse screen (Figure 5-11). This
screen is designed to protect the pumps from foreign
object damage and to provide for fault isolation.
Scavenge oil (and air) enters the bore of each screen
axially on the open inner end. It exits into a cast annulus
which discharges directly into the scavenge pump inlet.

A spring-loaded poppet-type, cold oil relief valve
is incorporated in this sytem. This valve prevents excessive supply pressure during cold starts when high oil
viscosity creates high line pressures. Cracking pressure

5-9

FM 1-506

These screens may be removed for inspection if chip
generation is suspected
Scavenge pumps. Six scavenge pumps are in linewith the lube supply pump on a common shaft (Figure5-10). Positioning of the pump elements is determinedby these factors:
The lube supply element is placed in the least
vulnerable location and isolated from scavenge
elements atone end
The B-sump element is placed at the other endof the pump to help isolate it from the otherscavenge elements. This element is the only one
with an elevated inlet pressure.
Pump windmilling experience on other engine
scavenge pumps shows that adjacent pumps
tend to cut each other off due to interelement
leaks at very low speed Therefore, the two
A-sump elements are placed adjacent, as are thethree C-sump elements, to reduce the possibiityof both elements in a sump being inoperative
simultaneously.
Porting simplification for the gearbox coring
determines relative positions of A-sump, Bsump, and C-sump elements.

5-10

Scavenge Discharge Passage. The common discharge of all six scavenge pumps is cast into the gearbox
at the top of the pump cavity. Top discharge facilitatespriming by clearing air bubbles and by wetting all pumping elements from the discharge of first pumps to prime.
The discharge cavity is tapered to enlarge as each pumpdischarge enters the flow steam. This keeps discharge
velocity relatively constant. It also tends to avoid air traps
which could short-circuit pumping at windmilling speeds.
This discharge plenum flows into the core to the chip
detector. Flow leaving the chip detector passes to thefuel-oil cooler in series with the air-oil cooler. To promote
faster warm-up and guard against plugged coolers, abypass valve is provided which bypasses both coolers.
Air-oil cooling is an integral part of the mainframe casting. Scavenge oil enters a manifold at the tank top. It thenflows in a serpentine fashion in and out through thehollow scroll vanes and box-sectioned hub. Air for theparticle separator is pulled across the vanes by thescavenge air blower providing the oil cooling process.
Exit from the air-oil cooler is through three holes at the
top of the tank. These outlets disperse the oil over thetank surfaces on both sides to settle in the tank. The oiltank vents to the AGB.

FM 1-506

Emergency Oil System
The T-700-GE-701 engine is designed to have two
oil jets to provide each main bearing with oil for lubricating and cooling (Figure 5-12).
In addition to being designed for normal engine
operation, the system provides for operation if the normal
oil supply from the primary system is interrupted. The
AGB and C-sump components can continue to operate
at least 6 minutes with residual oil present. The No. 4bearing in the B-sump and the bearings and gears in the
A-sump are provided with emergency air-oil mist syterns
located in each sump. The emergeney oil system forms
part of the normal full-time lubrication system and incorporates one full set of main bearing oil jets operating inparallel with the primary jets. The dual-jet system also
provides redundancy to minimize the effect of oil jet
plugging.
A small reservoir, curved to tit the A- and B-sumps,
retains a sufficient amount of oil to provide air-oil mist
when normal lubrication is interrupted The total sump
oil supply is fed into the reservoir at the top. Top feed

prevents reservoir drainage if the supply line is damaged
Primary oil jets, squeeze film damper, and uncritical lube
jets are connected to a standpipe at the top of the tank.
Secondary or emergency jets are similarly connected to
the lowest point in the tank.
Secondary jets are only located at points where
lubrication is vital for short-duration emergeney operation. Each secondary oil jet has a companion air jet or air
source which flows over the end of the oil jet and impinges
on the lubricated part. The air jets aspirate oil mist whennormal oil supply pressure is lost. They are pressurizedfrom the seal pressurization cavities and operate continuously with no valving required.
Component Description
The oil filter (Figure 5-13) consists of threesubassemblies:
Filter element.
Bowl and impending bypass indicator.
Bypass valve and inlet screen.
Filter Element. Media used in this falter are hightemperature materials em-staining organic and inorganic

5-11

FM 1-506

fibers. The layered media are faced on both sides withstainless steel mesh. This mesh provides mechanicalsupport to resist collapse when pressure loads become
high. Pleating of the faced media adds surface area andmechanical rigidity. A perforated steel tube in the bore
also adds rigidity and retains the circular shape of theelement.

5-12

The media and support tube are epoxy-bonded toformed sheet metal and caps. These end caps include an
O-ring groove which seals inlet to outlet leak paths at eachend
Filtration level selected is 100 percent of all particles
three microns or larger and is disposable when saturated

FM 1-506

with debris. Support of the falter element is provided by
the bypass valve on one end and the impending bypass
indicator on the other. The indicator end has a springloaded sleeve which restrains the filter axially.
Bowl and Bypass Indicator. An aluminum bowl
houses the element and contains the impending bypass
indicator at the end. Mounting is horizontal to fit the
space available and provide ready access for servicing
Impending bypass indication is provided by a small unit
which is part of the bowl assembly. The indicator isinstalled from the inside of the bowl. It is retained in placewith an external retaining ring. Basic mechanics ofoperation are as follows:
Different pressure between filter inlet and outlet acts to move a piston against a spring at 44 to
60 psi.
Piston contains a magnet which normally attracts a redbutton assembly and holds it seatedagainst its spring. When the piston moves, thebutton is released. It extends 3/16 inch tovisually indicate an impending bypass condition.
Button is physically reattained from tripping bya cold lockout bimetallic latch if temperature isless than 100 to 130°F. This prevents a false trip
during cold starts.
As the button is released, a small spring-loadedball also moves out of position to latch the button and block reset. The internal piston assembly automatically resets on shutdown;
however, the indicator remains latched out.
After removing the filter element and the bowl
from the gearbox, a spring-loaded sleevearound the indicator moves aft and pulls thepiston assembly to a tripped position. This
causes the button to trip if operation is attempted with no filter in the bowl.
To react the indicator, the bowl is held verticallyso the button latch ball can roll out of the latchedposition. The button is then manually reset.
If the bowl is reassembled with no falter, the indicator
will trip when the temperature exceeds the 100 to 130°F
lock-out level. The internal latch mechanism prevents
resetting the button without disassembling the bowl.
Resetting must be done with the bowl removed from theaccessory gearbox and held verically, button up, torelease the latch.
Oil Filter Bypass Sensor. The oil filter bypass sensor
is a differential-pressure switch which senses fiilter inlet
minus outlet pressure.

The sensor consists of a spring-loaded piston whichmoves aft at high falter differential pressure (60 to 80 psi)and magnetically releases a microswitch lever. The
switch is in a sealed cavity separated from the oil and iswired to a hermetically sealed electrical connector. The
switch connects 28-VDC aircraft power when tripped
and reopens the circuit at 15 psi minimum differential
No latch is used in the sensor so resetting is automatic.
Also, there is no cold lockout. The pilot will be informedof falter bypassing during cold start warm-ups. Sensor
tolerance range is set slightly below the tolerance range
of bypass valve cracking pressure. Therefore, bypassing
will not occur without pilot warning. The impending
bypass indicator will show need-to-change filter elements. This sensor provides backup warning if maintenance action is not taken.
Lubrication and Scavenge Pump. The lube andscavenge pump is a Gerotor-type pump of cartridge
design, located on the forward side of the accessory
gearbox (refer back to Figure 5-10). It fits into a precisionbore in the gearbox casing. The Gerotor-type pump waschosen because of its wear resistance and efficiency.
Gerotor elements are similar to male gear inside a female(internal) gear with one less tooth on the inner member.
The inner Gerotors are keyed to the drive shaft, andthe outer Gerotors are pocketed in individual eccentricrings. AS the assembly rotates, oil is drawn into an expanding cavity between teeth on one side. The oil is
expelled when the cavity contracts approximately 180°
away. Inlet and discharge ports are cast into the port
plates. They are shaped and positioned to fill and empty
at proper timing for maximum volumetric efficiency andresistance to inlet cavitation.
There are seven different elements in the pump fromthe spline end forward. They are the lube supply element,
C-sump cover, C-sump aft, C-sump forward, A-sumpforward, A-sump aft, and B-sump Delta scavenge elements.
The port plate eccentric rings and Gerotors areassembled into a surrounding concentric aluminum
tubular housing The housing maintains all elements inproper alignment. The oil suction and discharge passages from the Gerotors are brought radially through the
housing. They match the appropriate locations of themating passages in the engine gearbox casing. The entire
stack of port plates is retained in the housing with theretaining rings at the spline end. The outermost end of
the housing has an integrally cast cover. The cover bolt
holes are arranged to orient the pump assembly in the
gearbox housing during installation.
5-13

FM 1-506

Cold Oil Relief Valve. The cold oil relief valve
protects the oil supply system from overpressure duringcold starts (refer back to Figure 5-9). It is a conventional
poppet-type valve with a cracking pressure of 120-180 psi.
Valve tolerances are held sufficiently close to achieve thedesired sacking pressure without adjustment shims or
selective fitting of parts.
The valve includes a No. 10-32 threaded hole on theoutside. This allows for the usc of a bolt as a pulling
handle during valve removal from the AGB.
Oil Cooler. The fuel-oil cooler is a tube and shell
design (Figure 5-14). It cools the combined output of the
scavenge discharge oil that is ported through gearboxcored passages to the cooler. The cooler is mounted
adjacent to the fuel-boost pump on the forward side of
the gearbox. Oil and fuel porting enter on the same end
via face porting to the gearbox. Fuel is used as the coolant.
It is provided to the cooler via the boost pump, fuel falter,and hydromechanical control unit. A counterparallel
flow, miltipass cooler design is used to minimize pressure
drop while obtaining maximum cooler effectiveness.
Fuel flows through the tubes, while the oil flows over the
tubes resulting in the counterparallel flow arrangement.

5-14

Oil Cooler Bypass Valve. Design of the oil cooler
bypass valve is identical to the cold oil relief valve with an
exception (refer back to Figure 5-9). A lighter spring isutilized to obtain a lower cracking pressure of 22-28 psi.
Housing modifications prevent inadvertent interchangewith the cold oil relief valve.
Chip Detector. The chip detector in the common
scavenge line is the engine diagnostic device most likely
to provide first warning of impending part failure (Figure5-15).
The chip detector magnetically attracts electricallyconductive ferrous chips. The chips bridge the gap between the detector’s electrodes and close a circuit inseries with the aircraft cockpit indicator (warning light).
The chip detecting gap has a magnetic field induced in
tapered pose pieces at each end of a cylindrical permanent magnet. A single ferrous chip 0.090 inch in length
or longer will be indicated if magnetically attracted tobridge the pole pieces. The local mangetic field is intenseat the gap and tends to orient particles in the bridging
direction. Smaller particles tend to form chains until thepole pieces are bridgedFM 1-506

Nonconductive particles greater than 0.015 inch aretrapped inside the screen for visual examination. Smaller
particles will be found either in the lube tank or in the lube
supply falter.
The detector housing pushes into the accessory
gearbox It is retained by two captive bolts used in common with other accessories. Self-locking inserts in the
gearbox ensure retention of these bolts if assembly torqueis improperly low.

Venting System
A-Sump. The A-sump centervent handles air-oil
separation and overboard venting from these sources:
A-sump seals and emergency air system.
Scavenge pumped air from the lube tank.
Accessory gearbox vent (no air sources).
B-sump centervent flow which passes throughthe intershaft seal.
Path of this vent is into the bore of the power turbine
shaft and torque-reference tube and out the aft end of theengine through the C-sump cover. The centrifugal air-oil
separator vent holes in the power turbine shaft are located under the forward end of the high-speed shaft.
Windage from PTO gear locknut wrenching slots assists
in turning oil back into the sump. Air from the sump andintershaft seal flows inward radially through these holes
in the power turbine shaft. The air must flow forward in
the annulus between the power turbine shaft and thetorque-reference tube. Movement of air is blocked by astandoff ring on the reference tube OD. The forward
axial passage of the air centrifuges oil droplets outward

to the bore of the power turbine shaft. They either flow
back into the sump at the centervent or at small weep
holes forward of the PT shaft spline. Dried air then exits
through multiple rows of holes in the reference tube andout the aft C-sump cover. Some remaining oil in this air
is spun into the C-Sump if it has condensed in transit. Any
additional accumulated oil is then scavenged through theC-sump cover.
B-Sump. A centervent on the forward side of theNo. 4 bearing accommodates air entering the sump at thelabyrinth seals at each end Two rows of small holes are
drilled in a radially thickened section of the forward seal
runner. Use of many small holes increases the surface
area of metal in contact with exiting oil droplets. These
small holes also reduce effective window area for any
droplets which may have a trajectory aimed directly at theholes. Ater the air is inside these holes, it follows atortuous path through additional rows of holes in theturbine shaft and compressor rear shaft. The air thenenters the annlus between the high- and low-speed
shafts. In doing this, remaining oil is spun back into thesump.
About 70 percent of B-sump centervent flow moves
forward through the bore of the compresor tiebolt andintershaft seal. It exits at the A-sump centervent. Oilweep holes are provided near the aft end of the compressor tiebolt. These weep holes keep oil out of the rotor by
returning it to the sump. A rotor seal is provided hereto
keep any weepage out of the seal air. This airflow keeps
the compressor tiebolt relatively cool and uniformly
clamped.
The remaining 30 percent of B-sump centervent air
joins the inner balance piston seal leakage flow. It exits
aft under the gas generator turbine wheels.
C-Sump. Centerventing the C-sump is a passage
between the aft end of the PT shaft and a stationary
standpipe built into the C-sump cover. Windage at thetorque and speed-sensor teeth and in the annulus between the reference tube and the standpipe will return oil
droplets to the sump. Weep holes are provided throughthe reference tube, shaft, and bearing spacer to allow oilfrom C-or A-sumps to enter the C-sump. C-sump coverscavenging through the C-sump housing removes
remaining oil accumulation from the centerventing
process during locked PT rotor operation and normal
operation.
Oil Tank. After being routed through air-oil cooler
pasages into the oil tank, air from the scavenge pumps
flows down the radial drive shaft passage (Axis A) into
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the A-sump. Centerventing occurs after air enters the
A-sump.
Accessory Gearbox. The accessory gearbox isvented through the Axis A pad via the mainframe oil tankand eventually through the A-sump. The AGB, tank, andA-sump essentially operate at the same pressure levels
since they are interconnected.
LUBRICATION SYSTEM FOR T-62
The lubrication system consists of–
Pump.
Internal oil passages.
Oil filter assembly.
Falter bypass relief valve.
Pressure switch (mounted externally).
Oil jet ring,
Sump.
The oil falter cavity, oil passages, and oil sump arebuilt into the reduction drive housing. Two oil separator
plates are installed on the accessory drive gear. Lubrication system capacity is 3 quarts and is a wet-sump system.
Oil is drawn out of the sump into the pump housing.
The oil is carried between the pump gear teeth and pumphousing wall. It is then forced through drilled passages
to the oil filter housing.
Oil under pump pressure enters the bottom of thefilter housing and passes through the filter element (fromoutside to inside). It then flows out the housing througha passage in the falter element cap. A relief valve in thefilter element cap unseats at a differential pressure of 15

5-16

to 25 psi. This allows oil to flow from outside the filter
element, through a passage in the falter element cap, tothe falter outlet passage. If the filter element becomesclogged, the valve will open and allow oil to bypass thefalter element.
From the filter, oil is forced into a passage to thesystem relief valve and to four oil jets. The oil jet ring,which encircles the high-speed input pinion, contains
three of these jets. It sprays oil to the points where thehigh-speed input pinion meshes with the three planetary
gears. One jet directs a spray between the end of theoutput shaft and the high-speed pinion to create a mist
for lubrication of the rotor shaft bearings. The remaining
gears and bearings are lubricated by air-oil mist created
when oil strikes the planetary gears and high-speed
pinion.
System pressure is maintained at 15 to 25 psi by asystem relief valve. The valve regulates pressure bybypassing excessive pressure directly into the reduction
drive housing. The bypassed oil strikes the inside surfaceof the air inlet housing, aiding in cooling the oil. Bypassed
oil returns to the sump by gravity flow through an opening
in the bottom of the planet carrier.
The normally open contacts of the low oil pressure
switch close on increasing oil pressure at 5 to 7 psi. Whenthe switch contacts close, the low oil pressure cirecuit isdeenergized. At rated engine speed a drop in oil pressure
below 5 to 7 psi will open the low oil pressure switch
contacts. Through electrical circuitry, the drop in oilpressure will also close the main fuel solenoid valve andshut down the engine.

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CHAPTER 6

IGNITION SYSTEMS

Turbine engine ignition systems fall into twogeneral classifications. The induction type produces
high-tension sparks by conventional induction coils.
The capacitor type causes ignition by high-energy andvery high temperature sparks produced by a condenser
discharge. A third kind of ignition system not widelyadopted uses aglow plug.
DESCRIPTION
Igniter-exciter components are contained in hermetically sealed boxes. With a malfunction, it is necessary
to replace the entire exciter. Shielded cable is employedthroughout the system to protect against abrasion andignition system radio interference.
Ignition systems are not only used for engine starting
but also for ignition standby protection. Ignition systems
would be used to relight the engine if an in flight flameout
occurred when operating under potentially unstableflight conditions. Turbine engines sometimes tend toflameout because of an overrich fuel-air ratio caused bya momentary fluctuation of air supply at the compressor
inlet. A change in airflow at the compressor inlet or theentrance to the aircraft inlet duct may cause a condition
fuel control cannot immediately compensate for.
Flameout results. Flying in turbulent air, ingestion of abird, or ingestion of ice broken loose at the engine inlet
may cause such a situation. When one or both of theigniter plugs are operating, the engine will relight automatically after fuel control compensation takes place or
the inlet condition corrects itself. The pilot may not beaware that flameout has occurred. When ignition is used
as a precaution against flameout, the prescribed ignitionuse time limitations must be observed. This prevents
overheating the ignition system components and enhances the life span of the ignition system.
EARLY INDUCTION TYPE
Early turbine engine ignition systems evolved using
the tried principles developed for the reciprocating engine. Some of the early systems employed a vibrator andtransformer combination. This was similar to the booster
coils used for starting reciprocating engines. Other units

substituted a small electric motor driven cam. This
provided the neccessary pulsating magnetic field to theprimary coil of the transformer. Several variations appeared, all using the same basic principle. This principle
was high-voltage induction using a transformer to reachthe necessary voltage capable of causing an arc across thewide-gap jet igniter plug. An interesting variation of thistransformer-type ignition system is the opposite-polarity
system. In this circuit two electrodes extend into the
combustion chamber. Each electrode alternately becomes highly positively and negatively charged, causing avery high potential difference to exist across theelectrodes.
MODERN CAPACITOR TYPE
The high-energy, capacitor-type ignition system has
been universally accepted for gas turbine engines. It
provides both high voltage and exceptionally hot spark
which covers a large area. Excellent chances of igniting
the fuel-air mixture are assured at reasonably high altitudes.
The term “high energy” is used throughout thischapter to describe the capacitor-type ignition system.
Strictly speaking the amount of energy produced is verysmall. The intense spark is obtained by expending a small
amount of electric energy in a very small amount of time.
Energy is the capacity for doing work. It can be
expressed as the product of the electrical power (watt)and time. Gas turbine ignition systems are rated in joules.
The joule is also an expression of electric energy. It isequal to the amount of energy expended in one secondby an electric current of one ampere through a resistanceof one ohm. The relationship between these terms canbe expressed by the formula—w=1
twhere w= watts (power)
j = joulest = time (second)
All other factors being equal, the spark temperature
is determined by the power level reached. A hightemperature spark results from increasing the energy
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level (1) or shortening the duration (t) of the spark.
Increasing the energy level results in a heavier, bulkier
ignition unit. Energy delivered to the sparkplug is about
30 to 40 percent of total energy stored in the capacitor.
Higher erosion rates on the igniter-plug electrodes occur
because of heavy current flowing for a comparatively longtime. Much of the spark is wasted since ignition takes
place in a matter of microseconds (microsec). On theother hand, since heat is lost to the igniter-plug electrodes
and the fuel-air mixture is never completely gaseous, thespark duration cannot be too short.
The relationship between watts and time is shownin the following table. The example is for a 4-joule
ignition unit (4 joules appearing at the plug).
In an actual capacitor-discharge ignition system,
most of the total energy available to the igniter plug isHigh-voltage capacitor ignition system with DC
or AC input.
Low-voltage capacitor ignition system with DC
or AC input.
High-Voltage Capacitor System - DC input(more than 5000 VDC to the plug)
This system is a typical turbojet engine system in use
today. This system can include: two exciter units, twotransformers, two intermediate ignition leads, and two
high-tension leads. Depending on engine configurationa dual ignition is provided on the engine by two separately
mounted exciters or by twin circuits throughout the exciter.
As operation begins, the power source delivers
28VDC (maximum) input to the system. Each triggering

dissipated in 10 to 100 microsec (0.000010 to 0.000100
seconds). The system above would actually deliver
80,000 watts if the spark duration was 50 microsec.

circuit is connected to a spark igniter. The operation
described here takes place in each individual circuit.
Except for the mechanical features of the armature, theoperation is essentially the same in both Units.
As a safety factor, the ignition system is actually adual system designed to fire two igniter plugs.
Before the electrical energy reaches the exciter unit,
it passes through a falter. This filter prevents noise voltage
from being induced into the aircraft electrical system.
The low-voltage input power operates a DC motor, which
drives one multilobe cam and one single-lobe cam. At
the same time, input power is supplied to a set of breaker
points that are actuated by the multilobe cam.
From the breaker points, a rapidly interrupted current is delivered to an automatic transformer. When the
breaker closes, the flow of current through the primarywinding of the transformer establishes a magnetic field
When the breaker opens, the flow of current stops. The
collapse of the field induces a voltage in the secondary
winding of the transformer. This voltage causes a puke ofcurrent to flow into the storage capacitor. The voltage

1
w= t
=

4
0.000050= 80,000

To review, the spark temperature (a function of thewatts value) is the most important characteristic of any
ignition system. All three factors – watts, energy, andtime – must be considered before any ignition systemeffectiveness can be determined
HIGH-ENERGY TYPE
Ignition systems for jet engines are divided into
induction and capacitor discharge types. The capacitor
discharge type can be further divided into two basic
categories:

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flows through the rectifier, which limits the flow to a single
direction. With repeated pulses the storage capacitor
assumes a charge up to a maximum of 4 joules. (One joule
per second equals 1 watt.)
The storage capacitor is connected to the spark
igniter through the triggering transformer and a contactor, normally open. When the charge on the capacitor
builds up, the contactor is closed by the mechanical actionof the single-lobe cam. A portion of the charge flows
through the primary of the triggering transformer and thecapacitor connected in series with it.
This current induces a high voltage in the secondary
which ionizes the gap at the spark igniter. When the spark
igniter is made conductive, the storage capacitor discharges the remainder of its accumulated energy. Thisis done together with the charge from the capacitor inseries with the primary of the triggering transformer.
The spark rate at the spark igniter varies in proportion to the voltage of the DC power supply, which affects
the RPM of the motor. However, since both cams are
geared to the same shaft, the storage capacitor always
accumulates its store of energy from the same number ofpukes before discharge.
The employment of the high-frequency triggeringtransformer, with a low-reactance secondary winding,holds the duration of the discharge to a minimum. This
concentration of maximum energy in minimum timeachieves an optimum spark for ignition. An optimum
spark is capable of blasting carbon deposits and vaporizing globules of fuel.
All high voltage in the triggering circuits is completely isolated from the primary circuits. The complete
exciter is hermetically sealed, protecting all components
from adverse operating conditions and eliminating
flashover at altitude due to pressure change. This also
ensures shielding against leakage of high-frequency voltage interfering with the radio reception.
Two igniter plugs are mounted in the combustion
section outer case. The spark igniters are generally located in two diametrically opposite combustion liners.
The igniters receive the electrical output from the ignition
exciter unit. The igniters discharge the electrical output
from the ignition exciter unit. And they discharge theelectric energy during engine starting to ignite the fuel-air
mixture in the combustion liners.
Typical specifications for this system are as follows
Normal: 24VDC
Input voltage:
Operating limits: 14 to 30 VDC

4 to 8 per second at each plug,depending on input voltage
Designed to fire: 2 igniter plugs
Accumulated
energy3 joulesDuty cycle:
2 minutes ON, 3 minutes 0FF,
2 minutes ON, 23 minutes OFF
High-Voltage Capacitor System-AC Input
Power is supplied to the unit input connector fromthe 115-volt, 400-cycle source in the aircraft. Power is firstled through a filter which blocks conducted noise voltage
from feeding back into the airplane electrical system.
From the filter, the circuit is completed through theprimary of the power transformer to groundIn the secondary of the power transformer, an alternating voltage is generated at a level approximating 1700
volts. During the first half-cycle this follows a circuit
through the doubler capacitor and rectifier A to ground,leaving the capacitor charged. During the second halfcycle when the polarity reverses, this circuit is blocked byrectifier A. The flow of this puke is through ground to
the storage capacitor, rectifier B, resistor, doubler
capacitor, and back to the power transformer.
With each pulse the storage capacitor assumes agreater charge. By virture of the action of the doubler
capacitor, the charge approaches voltage approximately
twice that generated in the power transformer. When thisvoltage reaches the predetermined level calibrated forthe spark gap in the discharge tube X (the control gap),
the gap breaks down. This allows a portion of the accumulated charge to flow through the primary of thehigh-tension transformer and the trigger capacitor inseries with it. This surge of current induces a very highvoltage in the secondary of the high-tension transformer.
This surge is enough to ionize the gap in discharge tube Y.
The storage capacitor immediately discharges theremainder of its accumulated energy through the spark
igniter. This produces a capacitive spark of very highenergy.
The bleeder resistors are provided to dissipate theresidual charge on the trigger capacitor. This is accomplished between the completion of one discharge at
the spark igniter and the succession of the next cycle.
Typical specifications for this system are as follows:
Normal: 115-volts, 400-cycle
Input voltage:AC
Operating limits 90 to 1241 volts
Spark rate:

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Spark rate:
Designed toignite:Accumulated
energy:Duty cycle:

Normal: 1.50 to 275 per sec
Operating limits: 0.75 to
5.00 per second
One spark igniter

14 to 17 joules
2 minutes ON, 3 minutes OFF
2 minutes ON, 23 minutes OFF
Low-Voltage Capacitor System-DC Input(less than 1000 volts to the plug)
The basis of operation on which the low-voltage,high-energy ignition system is built is the self-ionizing
feature of the igniter plug. In the high-voltage system adouble spark is produced. The first part consists of ahigh-voltage component to ionize (make conductive)
the gap between the igniter plug electrodes. The
second high-energy, low-voltage portion follows. The
low-voltage, high-energy spark is similar except thationization is effected by the self-ionizing igniter plug.
The main ignition unit changes the amplitude andthe frequency characteristics of aircraft power into pulsating DC. To do this, the components in the ignition unit
are grouped in stages to falter, amplify, rectify, and store
an electric charge.
The spark plugs used in the ignition system are theshunted-gap type, which are self-ionizing and designed
for low-tension (relatively low voltage) applications.
Although the spark plug fires at relatively low voltage, a high-temperature spark is obtained from the speed
the energy is discharged across the gap. The spark is ofshort duration (40 microsec), but momentarily expends
a great amount of power. Tank capacitor discharge current from the main ignition unit surges to the spark-plug
electrodes. This builds a potential between the centerelectrode and ground electrode. The semiconducting
material shunts the electrodes. When the potential between electrodes reaches approximately 800 volts, it
forces enough current through the semiconductor toionize the air gap between the electrodes. The full-tank
capacitor current arcs instantly across the ionized gap,
emitting a high-energy spark.
COMBINATION OR DUAL DUTY
The operation of this system will not be discussed indetail. However, it is mentioned to make you aware of
the system.

6-4

The ignition system includes one intermittent-duty
exciter, one continuous-duty exciter, one intermediate
voltage lead, and two high-tension leads. It is designed
to tire two spark igniters during ground starts. This isaccomplished by the 20-joule intermittent-duty exciter or
one spark igniter during flight by the 4-joule continuousduty exciter.
When intermittent operation is employed, DC
power is supplied to the input of the intermittent-duty
exciter from the 24-volt aircraft electrical system. It is firstpassed through a radio noise filter to prevent highfrequency feedback. When continuous operation is
employed, power is supplied to the input of thecontinuous-duty exciter from the 115-volt, 400-cycle
AC source in the aircraft.
ELECTRONICThis modified capacity-type system provides ignition for turbojet and turboprop engines. It is required
only for starting the engine. Once combustion begins theflame is continuous. Figure 6-1 shows the components ofatypical electronic ignition system.

The system consists of a dynamotor/regulator/filter
assembly, an exciter unit, two high-tension transformer
units, two high-tension leads, and two igniter plugs. The

FM 1-506

necessary interconnecting cables, leads, control switches,
and associated equipment for operation are used withthese components.
The dynamotor is used to step up the direct currentof the aircraft battery or the external power supply to theoperating voltage of the exciter unit. This voltage is used
to charge two storage capacitors which store the energy
used for ignition.
In this system, the energy required to tire the igniterplug in the engine burner is not stored in an inductor coil
Instead, the energy is stored in capacitors. Each discharge circuit incorporates two storage capacitors. Bothare located in the exciter unit. The voltage across thesecapacitors is stepped up by transformer units. At the
instant of igniter plug firing, the resistance of the gap islowered sufficiently to permit the larger capacitor todischarge across the gap. The discharge of the second
capacitor is of low voltage but very high energy. The
result is a spark of great heat intensity. It is capable
not only of igniting abnormal fuel mixtures but also ofburning away any foreign deposits on the plug electrodes.
The exciter is a dual unit, and it produces sparks ateach of the two igniter plugs. A continuous series ofsparks is produced until the engine starts. The battery
current is then cut off. The plugs do not tire while theengine is operating.
TURBINE ENGINE IGNITERS
Turbine engine igniters come in many sizes andshapes depending on what their function is. The
electrodes of the plugs used with high-energy ignition
systems must accommodate a much higher energy current than the electrodes of conventional sparkplugs. The
high-energy current causes more rapid igniter electrodeerosion than in reciprocating engine sparkplugs. This is
not a problem because of the relatively short time aturbine engine ignition system is in operation. This is oneof the reasons for not operating the gas turbine ignition
system any longer than necessary. Igniter plug gaps arelarge in comparison with those of conventional spark
plugs. The gaps are large because the operating pressure
at which the plug is fired is much lower than that of areciprocating engine.
Most igniter plugs are of the annular-gap type. Constrained gaps are used in some engines. Normally, toprovide an effective spark the annular-gap plug projects
slightly into the combustion chamber liner. The spark of
the constrained-gap plug does not closely follow theface of the plug. Instead, it tends to jump an arc which
carries it beyond the face of the chamber liner. The

constrained-gap plug need not project into the liner. The
result is that the electrode operates at a cooler temperature than the annular-gap plug.
The turbojet ignition system is designed for severe
altitude conditions common to military operation. It israrely taxed to its full capability by transport use.
Flameout is much less common than it was, and flightrelight is not normally required of the ignition system.
Ignition problems in general are minor compared to theconstant attention required by the piston engine system.
Airborne ignition analysis equipment is unnecessary.
Spark igniter plug replacement is greatly minimized.
Only two plugs per engine are used.
The trends taking place in the gas turbine ignitionarea are —Use of AC power inputs, which eliminates thevibrator, a major source of trouble.
Use of solid-state rectifiers.
Use of two discharge tubes, which permits thelevel of stored energy per spark to be moreconsistent throughout the life of the exciter.
Sealed units.
Longer time between overhauls.
The advent of short-range jets, which increased
the ratio of ignition-on time to engine operationand led to development of dual systems.
Two types of ignition systems in Army aircraft today
are General Electric’s T-701 and Lycoming's T-55-L-712
T-701 GENERAL ELECTRIC
The ignition system is an AC-powered, capacitordischarge, low-voltage system. It includes a dual exciterunit mounted on the right-hand side and two igniter plugs(Figure 6-2). The spark rate of each ignition circuit is two
sparks per second minimum; energy at the igniter plugs
is at least 0.25 joules per spark. The exciter is poweredby one winding of the engine alternator and is connected
to it by the yellow harness. The ignition system must be
turned off after starting by shorting the alternator output.
For normal starting the aircraft-ignition circuit is tied inwith the aircraft-starting system to de-energize the ignition system at the starter or dropout speed
The igniter plug is a homogeneous semiconductor,
surface-gap type spark plug using air cooling of the firing
tip (Figure 6-3). It provides a projected electrical discharge for lightoff of the combustor. An engine set
consists of igniters located at the 4 o’clock and 8 o’clock
positions. They are mounted in the midframe using a
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FM 1-506

screw-in boss and extend inward through the outer panelof the combustion liner.
The ignition system operates with a maximum output of 7000 volts. To create a spark across the electrodegap with this voltage, the gap surface is a semiconductormaterial, homogeneous button extending into the tip of
the plug. It is, therefore, capable of coping with erosion
over a long period of operating time. Consistent with thislong life objective, the center electrode is pure tungsten
and the outer electrode is tungsten alloy. Theseelectrodes are nickel-plated to prevent oxidation, aproblem further minimized by tip cooling. Compressor

6-6

discharge air provides cooling air which enters throughsix holes around the tip body and exits through twelve
holes at the tip end.
T-55-L-712 LYCOMING
The high potential ignition pulse is developed by theignition exciter (Figure 6-4). A direct current at 28 volts
is applied to the input of the exciter. Current flows
throught the primary transformer winding the bias coil,and the vibrator points to ground This generates magnetic lines of force which permeate the transformer andbias coil cores, attract the vibrator reed upward, andinterrupt the circuit. As current flow ceases, the lines offorce collapse, and the reed falls back. This closes thecircuit. This cycle repeats at a rate proportional to the
input voltage. The resultant current flows in pulses,
causing magnetic lines of force to build up and collapse
with each pulsation. These lines induce voltage acrossthe secondary. They are transformed to a higher potential by virtue of an increased number of windings comprising the secondary. The diodes rectify the pulsating
current back into direct current to charge the capacitors.
The charge on the capacitors continues to build up at a
rate proportional to input voltage until a potential of 2500
AC volts exists. The calibrated spark gaps ionize at thisvoltage creating an electrical path for the tiring pulse.
The capacitors discharge through this path into the leadcoil assembly for distribution to the spark igniters.

FM 1-506

Radio frequency energy is generated within the exciter during normal operation. An inductive capacitive
filter has been incorporated at the input. This will prevent
energy from being fed back onto the 28-volt input line.
Radio frequency interference on this line could damage
the operation of other electrical accessories. The filter isturned to radio frequencies. It does not offer any appreciable opposition to the flow of 28-volt direct current.
The ignition lead and coil assembly constitutes thehigh potential ignition wiring. This assembly incorporates two coils fed with high voltage from the two

outputs of the ignition exciter. The coil assemblies function as spark splitters distributing high voltage to four
igniter plugs. Each coil assembly has one input and two
outputs. The coil windings forma transformer having a
1:1 ratio. Any current flowing through either winding will
induce a voltage across the other. Even a shortened
igniter plug will not short out the high-voltage ignition
signal. The entire wiring harness is shielded andgrounded at airframe potential to suppress radiation ofradio frequency interference.

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CHAPTER 7

COOLING SYSTEM

The intense heat generated when fuel and air are
burned mandates that some means of cooling be
provided for all internal combustion engines.
Reciprocating engines are cooled either by passing air
over fins attached to the cylinders or by passing a liquid
coolant through jackets surrounding the cylinders.
Cooling is made easier because combustion occurs only
during every fourth stroke of a four-stroke-cycle engine.
In contrast, the burning process in a gas turbine engine iscontinuous, and nearly all of the cooling air must passthrough the inside of the engine. If enough air wereadmitted to the engine to provide an ideal air-fuel ratioof 151, internal temperatures would increase to morethan 40000°F. In practice, a large amount of air exceedingof the ideal ratio is admitted to the engine. This large
surplus of air cools the hot sections of the engine toacceptable temperatures ranging from 1100° to 1500°F.
COOLING OF ENGINE OUTER CASE
Figure 7-1 illustrates the approximate engine outercase (skin) temperatures encountered in a properlycooled dual-axial compressor turbojet engine. Becauseof the effect of cooling, the temperatures of the outsideof the case are considerably less than those encounteredinside the engine. The hottest spot occurs opposite theentrance to the first stage of the turbine. Although thegases have begun to cool a little at this point, the conductivity of the metal in the case carries the heat directly tothe outside skin.
COOLING OF COMBUSTION CHAMBER AND
GAS PRODUCER
The air passing through the engine the combustionchamber burner cans or liners. The cans are constructed
to induce a thin, fast-moving film of air over both the innerand outer surfaces of the can or liner. Can-annular-typeburners are frequently provided with a center tube to leadcooling air into the center of the burner to promote highcombustion efficiency and rapid dilution of the hot combustion gases while minimizing pressure losses. In all
types of gas turbines, large amounts of relatively cool air

7-1

FM 1-506

join and mix with the burned gases aft of the burners tocool the hot gases just before they enter the turbines.
All stage 1 and 2 airfoils in the GE T-701 engine areinternally cooled by means of compressor discharge air(Figures 7-2, 7-3). The stage 1 nozzle leading edge is
conection- and film-cooled with the air exiting through7-2

a shower-head series of holes in the leading edge. Aft ofthe leading edge, film-cooling air exits through convex
and concave side gill holes and trailing edge slots. Cooling air for the stage 2 nozzle is bled from the centrifugal
impeller exit and piped back through the midframe toenter the stage 2 nozzle. It cools by internal convection,

FM 1-506

exiting both by trailing edge holes and by inner bandholes. There, it cools the interstage seal. The turbine
blades are convection-cooled through radial holes in thedovetail. Air exits through tip holes. The stage 1 bladesalso employ trailing edge holes for cooling.
EXTERNAL COOLING
Cooling air inlets are frequently provided around theexterior of the engine to permit entrance of air to cool theturbine case, the bearings, and the turbine nozzle (Figure7-4). In some instances internal air is bled from the engine
compressor section and is vented to the bearings andother parts of the engine. Air vented into or from theengine is ejected into the exhaust stream. When anaccessory case is mounted at the front of the engine, it iscooled by inlet air. When located on the side of theengine, the case is cooled by outside air flowing around it.

The engine exterior and the engine nacelle arecooled by passing air between the case and the shell of
the nacelle. The engine compartment frequently is
divided into two sections. The forward section is built
around the engine air inlet duct; the aft section is built
around the engine. A fume-proof seal is provided
between the two sections. The advantage of such anarrangement is that fumes from possible leaks in the fueland oil lines contained in the forward section cannot
become ignited by contact with the hot sections of theengine. In flight ram air provides ample cooling of thetwo compartments. On the ground air circulation isprovided by the effect of reduced pressure at the rear of
the engine compartment produced by gases flowing fromthe exhaust nozzle.

7-3

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CHAPTER 8

ACCESSORY SYSTEMS

Accessories for gas turbine engines can be divided
into two categories those driven by bleed air taken fromthe compressor section of the engine; those driven
mechanically by an accessory drive shaft and gearbox
connected directly to the turbine shaft. The mechanicalconnection from the turbine shaft may be through anengine-mounted gearbox or through a power takeoff
shaft to a remotely mounted gearbox.
BLEED-AIR-DRIVEN ACCESSORIES
Gas turbines are unique among engines. High-pressure air is available for driving aircraft accessories
by air motors or air turbines. Compressor discharge
air at high pressure and temperature is bled from theengine through ports provided. This air is ducted as asource of power. It operates the air-conditioning units,
hydraulic pumps, thrust reverser actuators, andvarious mechanical actuators in the airplane. Air for
cockpit or cabin pressurization is also provided bybleed air from the engine compressor. On multiengined
aircraft equipped with pneumatic engine starters, one
engine is usually started from an auxiliary power unit
or a ground air source. Air from this operating engine
is bled through a system of ducts in the aircraft, topower the starters of the other engines.
The Pratt and Whitney dual-axial compressor turbine engine is an example that uses bleed air to operateaccessories. The JT3D turbofan engine is used on
aircraft such as the Boeing 707 and B-52 bomber. This
engine also uses a mechanical accessory gearbox. It
usually has three separate bleed air systems: high pressure, low pressure, and overboard. The high- and lowpressure systems are used to drive aircraft, engine
components, and accessories. The overboard is required
to preclude compressor instability.
Compressor bleed air is also used to anti-ice theengine air inlet guide vanes and, frequently, parts of the
air inlet duct. Low-pressure air has a pressure of approximately 50 psi and a temperature of more than 300°F.
This low-pressure air is taken from bleed air parts
compressor mid-section between the low- and high-pressurecompressors. High-pressure bleed air has a pressure ofabout 160 psi and a temperature of more than 650°F when
operating near sea level. This air is taken from the rearof the high-pressure compressor. The air available fordriving accessories and for other purposes in the aircraft
is usually about 3 or 4 percent of the primary engine
airflow. Keep in mind the air under pressure that is
extracted from the engine is not a bonus. Engine output
and fuel consumption are sacrificed.
MECHANICALLY DRIVEN ACCESSORIESThe other method of driving accessories is adirect, mechanical drive operated by gearing from thecompressor-turbine drive shaft. Accessory drives andaccessory mounting pads are provided in an enginemounted, accessory drive gearbox or in a remotely
mounted gearbox. On some turbojet engines, accessory
pads and mechanically powered drives are also provided
in the engine nose section. For dual compressor, axialflow engines, the main accessory drive gearbox usually
receives its power from the high-pressure compressor
drive shaft. Mechanically driven accessories include:
tachometers, generators (alternators), hydraulic pumps,
fuel pumps, oil pumps, fuel controls, starters, and (iisome instances) water pumps.
LYCOMING T-55 ENGINE
The power extraction system transmits power fromthe N1 and N2 systems to the accessory gearbox locatedat the 6 o’clock position on the inlet housing (Figures 8-1,
8-2). Most of these components receive their driving
force from the N1 system. A minimum amount of poweris extracted from the N2 system.
The starter gearbox, mounted at the 12 o’clock position, functions as a centrifuge for air-oil separation during
engine operation. The N1 system provides the driving
force. A single bevel gear is located at the front end of
the compressor shaft; meshes with a planetary geartrain
housed in the inlet housing This gear train transmits N1
power through two drive shafts: one to the starter gearbox, the other to the accessory gearbox to drive the idler
system. A gear located on the output power shaft

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8-2

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interconnects with the 90° angle gear in the output shaft
support housing. The gears drive the N2 section of theaccessory gear train.
The N2 driven overspeed governor in the fuel control is driven from a spur gear. This gear is attached to
the N2 system drive shaft which rotates on a ball bearing
with a pad, located at the 6 o’clock position on the inlet
housing
Customer power extraction is provided through thevertical bevel pinion gear located within the accessory
gear carrier. It is available for use by inserting acustomer-supplied drive shaft in its internal splines.
The accessory drive gearbox supplies mounting pads
for the engine-supporting accessories and provides thetransmission capability to drive these components.
Dimensional standardization permits interchanging
components among engines and eliminates the need foradapters for test equipment.
GENERAL ELECTRIC T-701
The accessory module mounts on the cold section
module at the 12 o’clock position of the main frame(Figures 8-3, 8-4). It includes the accessory drive gearbox (AGB) that is driven by a bevel-gear system from thecompressor rotor via a radial drive shaft. Several accessories are contained in or mounted on the front and rearcasings of the AGB. The rear face provides drive pads
for the engine starter, hydromechanical unit, inlet
separator blower, and a face-ported pad for the overspeed and drain valve. Pads for the alternator and fuelboost pump are on the front face. A cavity is provided
for the lube and scavenge pump and chip detector. Pads
are supplied for the oil cooler, fuel meter, and lube falter.
Cored passages in the AGB housing convey fuel and oil
between components.
Drive pad seals for the starter, hydromechanicalunit, and fuel boost pump drain into a common coredpassage in the AGB housing. It then drains to an external
port on the right-hand side of the mainframe.
ACCESSORY DRIVES
Accessory Drive System
The accessory drive system provides drives for boththe N1 driven accessory gearbox and the N2 driven
overspeed governor and tachometer drive assembly.
Provision is also made within the system for the drive ofcustomer-furnished accessories through the power

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takeoff pad located at the 2 o’clock position on the engine
inlet housing (Figure 8-5).
N1 drive is provided from a pinion gear (9) mounted
on the forward end of the compressor rotor shaft, driving
two bevel gears (10 and 19) located within the accessory
gear carrier.

8-4

The bevel gear located at the 6 o’clock position
within the carrier, being the accessory gearbox drive gear(10), is splined internally to accept the accessory gearbox
shaft (18). This drive shaft connects the gear carrier tothe accessory gearbox through the 90° pinion gear (16)which in turn is splined directly to the starter-generator
drive gear (15). The starter-generator drive gearprovides drive to all subordinate gears located with theaccessory gearbox housing.
The power takeoff drive is provided through thesecond bevel gear (19) located within the accessory gear
carrier and is used to drive airframe accessories.
The N2-driven overspeed governor and tachometer
drive gearbox (1) receives its drive from a spur gear (20)pressed to the power shaft aft of the sun gear. This gear
engages the N2 drive and driven gear package (8) locatedwithin the accessory gear carrier. This package, a seriesof three gears, provides an internal spline drive for thedrive shaft (2) which passes up through the 10 o’clock
inlet housing strut and into the gearbox (l).
The drive shaft then engages the internal splines of
the upper drive gear (3) which provides drive to thetachometer gear (5). This gear meshes directly with anidler gear (6) which in turn transmits the drive to the
combination torquemeter boost pump and overspeed
governor drive gear (7).
Main Accessory Drive Gearbox
The accessory drive gearbox (Figure 8-6) is mounted
at the 6 o’clock position of the engine inlet housing and isdriven through bevel gears from the front end of thecompressor rotor. Drive pads are provided on the rearof the gearbox for the fuel control, the starter-generator,
and the gas producer (nI) tachometer generator. The

FM 1-506

gearbox front side has mounting for the rotary oil pump
and also has an unused drive pad with connection line.
Oil scavenge lines are connected at the right rear on the
gearbox which serves as an oil collector sump, kept
practically empty by the pump. A chip detector plug islocated in the lower right side, and the oil filter is on theleft side. A drain line from the inlet guide vane actuatoris connected to the right side.

Power and Accessory Gearbox
The main power and accsessory drive gear trains are
enclosed in a single gear case. The gear case serves as the
structural support of the engine. All engine components,
including the engine-mounted accessories, are attachedto the case. At 100 percent engine speed, reduction
gearing reduces power turbine speed from 33,290 to 6016
RPM at the output drive pads. The power turbine gear
train has a torquemeter to measure engine output torque.

Accessories driven by the power turbine gear train are
the power turbine tachometer-generator (N2 and thepower turbine governor. The gas producer gear train
drives the oil pump, fuel pump, gas producer fuel control,
and tachometer-generator (N1). The gearbox has a spareaccessory mounting pad which is driven by the gas producer
gear train. During starting the starter-generator cranks
the engine through the gas producer gear train. Aftercompletion of the starting cycle, the starter-generator
functions as a generator.
STARTERS
Gas turbine engines are started by rotating the compressor. In the case of dual axial-flow compressor engines, the high-pressure compressor is usually the only
one rotated by the starter. First, compressor accelerates
to provide sufficient air under pressure to support combustion in the burners. Second, after fuel is introduced
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and the engine is fired, starter continues to assist theengine above the self-sustaining speed. The torque must
be in excess of that required to overcome rotor inertia
and engine friction and air loads.
Basic types of starters developed for gas turbine
engines are air turbine (pneumatic), electric motor,
hydraulic, combustion, and cartridge pnuematic. Air
turbine starters are the most commonly used. Smaller
engines generally use electric starters. Hydraulic starters
are frequently found in helicopters and some marine gas
turbine installations. Combustion and cartridgepneumatic starters are generally special application
devices that may require a self-contained starting system.
An impingement starting system is sometimes used An
impingement starter consists of simple jets of compressed
air piped to the inside of the compressor or turbine case.
The jet air blast is directed onto the compressor or
turbine rotor blades and causes them to rotate.
As the starter accelerates the compressor sufficiently to establish airflow through the engine, the ignition andthen the fuel are turned on. The exact sequence of thestarting procedure is important. There must be sufficient
airflow through the engine to support combustion when
the fuel-air mixture is ignited The fuel flow rate will not
be sufficient to enable the engine to accelerate until after
the self-sustaining or self-accelerating speed has beenreached If assistance from the starter were cutoff below
the self-sustaining speed, the engine would either fail toaccelerate to idle speed or decelerate. Deceleration
occurs because sufficient energy was not produced tosustain rotation or acceleration during the initial phase of
the starting cycle. The starter must continue to assist theengine above the self-staining speed to avoid a delay inthe starting cycle. This would result in a hot or hung(false) start, or a combination of both. In a hot start, theengine lights up, but the exhaust gas temperature exceedsthat allowed for an engine start. In hung or false start theengine lights up normally but, instead of increasing to idle
speed, the RPM remains at some lower value.
At the proper points in the sequence, the starter andthe ignition cut off. The higher the RPM before thestarter cuts out, the shorter the total time required for theengine to attain idle RPM. The engine and starter work
together to furnish the torque necessary for engine acceleration.
The most important requirement of a starter is to
produce sufficient torque to start the engine properly.
Engines must be rotated and accelerated above a certain
minimum rate if consistently good starts are to be

8-6

achieved. The torque characteristics of an acceptable
starter must be well above the required minimum.
Air Turbine
Air-turbine starters (which are also called
pneumatic starters) are used more than any other for
starting jet aircraft engines, particularly larger engines(Figure 8-7). A small geared air turbine is attached tothe engine starter pad located at the accessory drive
gearbox. Air-turbine starters must receive compressed
air from an external power source. A compressor
mounted on a ground unit or onboard the aircraft is one
such source. A small turbine engine usually drives theseunits. On multiengine aircraft, air is often bled from thefirst engine started and used to operate the starters forthe remaining engines. With an air-turbine starter, the air
supply must be of sufficient volume and pressure to meet
starter requirements. Otherwise, the starter torque may
not produce consistently successful starts within an acceptable time limit. When bled air from another operating engine is used, the engine being used for a compressed
air supply must be turning over fast enough to ensure
adequate air pressure to the starter of the engine being
startedStill another form of pneumatic starter is when thestarter itself is part of the ground power unit. The end of
a flexible shaft from the starter is placed in a connection
box on the engine to turn the compressor. When theengine starts, the flexible shaft is removed This typestarter might be used on small gas turbine engines. It
might also be used on a gas turbine engine for a missileor where airborne weight must be held to the minimum.

FM 1-506

Combustion
A combustion starter is essentially a small turboshaft gas turbine engine (Figure 8-8). Its chief advantages are quick starts because of the high torqueproduced and portability. The complete starting system
may be carried aboard the aircraft. This starter obtains
its power from hot expanding gases. These gases are
generated in the starter combustion chamber. They
occur by burning a combustible mixture either fuel, air,or monopropellant. A monopropellant is a speciallycompounded solid or liquid slow-burning substance.
The quantity of fuel or monopropellant required by
the starter is proportional to the length of time the starteris operating. Aircraft weight and size limitations requirethat the quantity of fuel or monoporpellant be held to aminimum in an airborne, combustion-starter system.
The starter burning time must be held to a minimum.
This is determined by the starter torque and the starter
cutout speed. The starter cutout speed depends on the
self-sustaining speed of the engine. The self-sustaining
speed determines, to a large extent, the amount of fuel ormonopropellant required. Ordinarily, sufficient fuel ormonopropellant to provide a minimum of two starts will
be carried aboard the aircraft. Some combustion starters
operate as simple air turbines. When an outside air
source from a ground unit or another engine which isalready started is connected to the starter, the combustion starter functions as a pneumatic starter.

Cartridge Pneumatic
A cartridge-pneumatic starter is a combustion
starter that operates from monopropellant contained ina cartridge. Prior to starting, the monopropellant
cartridge is inserted in the starter. Like a regular

combustion starter, a cartridge starter fires electricallyfrom a switch in the aircraft. Cartridge-pneumatic
starters can be operated as pneumatic starters whenprovided with outside air.
Electric Starting
Direct cranking and starter generator are the twoelectric starting systems for gas turbine aircraft (Figure
8-9).

Direct-cranking electric starting systems are
similar to those used on reciprocating engines.
Starter-generator starting systems are also similar to
direct-cranking electrical systems. Electrically, thetwo systems may be identical. But the starter generator
is permanently engaged with the engine shaft throughthe necessary drive gears. The direct-cranking startermust use some means of disengaging the starter from theshaft after the engine has started.
Direct-Cranking Gas Turbine Starters. On some gas
turbine engines, no overload release clutch or gear reduction mechanism is used. This is because of the low-torqueand high-speed requirement for starting. A reduced
voltage mechanism is used, however, in the starting assembly during starting.
Starter-Generator Starting System. Many gas turbinc aircraft are equipped with starter-generator systems. These starting systems use a combination
starter-generator which operates as a starter motor to

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drive the engine during starting. After the engine reaches
a self-sustaining speed, the starting system operates as a
generator to supply the electrical system power.
The starter-generator unit is basically a shunt generator with an additional heavy series winding. This series

8-8

winding is electrically connected to produce a strong field
and a resulting high torque for starting. Start-generator
units are desirable from an economical standpoint. One
unit can perform the function of both starter and generator. Additionally, the total weight of starting systemcomponents is reduced, and fewer spare parts requiredFM 1-506

CHAPTER 9

INSTRUMENT SYSTEMSEngine instruments are mounted on the instrument
panel in the pilot’s compartment. They are usuallygrouped together. They aid the pilot in monitoring theengine, rotor, or propeller RPM pressures temperatures and general engine performance. Refer to theengine’s maintenance manual for maintenance, operational limitations, and color code markings. Instruments
today read in vertical scales, digital readouts, and in somecases heads up display (HUD). The HUD systemprojects instrument images on the windshield in front of
the pilots. In this way, pilots do not have to divert theireyes from the view in front of and around them. The
vertical scale offers similar safety conditions. The pilots
have the ability to quickly glance at the instrument for areading. This chapter will discuss the most commonlyused instruments and the newer vertical scale (VIDS).
VERTICAL INSTRUMENT DISPLAY SYSTEM(VIDS)The VIDS (Figure 9-1) consists of a vertical strip
central display unit (CDU), two vertical strip pilot display
units (PDU), and two signal data converters (SDC).
These readings are shown by ascending and descending
columns of multicolored lights (red yellow, and green)measured against vertical scales. The scales operate as
segments and light in normal progression. They remainon as the received signal level increases. They go off innormal progression as the received signal level decreases.
When scales with red- and amber-coded segments below
green-coded receive a signal level of zero (bottom scale),
the segments light in normal progression and remain on.
When the first segment above the red or amber range
goes on, all red-or amber-coded segments go off. Thesesegments remain off until the received signal level indicates a reading at or within the red or amber range. At
that time all red- or amber-coded segments on the scale
display goon or off in normal progression. For an increasing indication when a scale with side arrows lights,
the corresponding side arrow also lights. As the segments
go on, the corresponding side arrows also go on, one at atime. Only the side arrow associated with the highest
percent indication of the corresponding scale will be on.

For a decreasing indication, scales with side arrows will
operate in the same manner. Only the side arrow associated with the highest percent indication of the corresponding scale will be on. The CDU and PDUs
contain photocells that automatically adjust indicatorlighting to ambient light. If either photocell fails, the lights
on the vertical scales or the PDUs and the CDU go off.
The DIM knob on the CDU has an override capabilty
which allows the pilot to manually set the display lightlevel.
TACHOMETER SYSTEM
The tachometer provides the pilot with a continuous
indication of engine RPM. A variety or combination ofsystems may be used on gas turbine engines. Gas
producer to gas generator tachometers, turbine and rotor
tachometers, and N-I and N-II tachometer are some of
the tachometer systems used (Figure 9-2).
Engine rotor RPM can be sensed by a mechanicallydriven tachometer generator, mechanically driven permanent magnet, or a pulse pickup which senses passing
compressor, fan blades, or gear teeth. The output or
signal from any of the sensors is directed to an appropriate indicator in the cockpit. The indicator iscalibrated to read directly in percent RPM. Dual axialflow compressor engines are usually provided with twotachometers: one indicates low-pressure compressor
speed (Nl), and the other indicates high-pressure compressor speed (N2). For the dual-axial and centrifugal
compressors the second tachometer indicates power
turbine speed (N2).
For most axial-flow compressor engines, thetachometer monitors RPM during engine start and indicates overspeed condition should one occur. Use of thetachometer for setting thrust is not recommended on
axial-flow compressor engines. The low-pressure compressor (Nl) tachometer on engine pressure ratio(EPR)-controlled dual-compressor engines may be used
as an approximate reference to set engine thrust in transient and certain other flight conditions. When the N1
tachometer sets engine thrust, the thrust setting should

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be more accurately adjusted by engine pressure ratio assoon as possible. On single-compressor, axial-flow engines, engine speed should not be used primarily to set orcheck engine thrust.
The General Electric T-700/701 system eliminatestach generators. Two Np sensors are located in the exhaust frame. One sensor extends through the 1:30 o’clock
strut and the other through the 10:30 o’clock strut. The
power turbine shaft is equipped with two pairs of teeth
which induce electrical pulses in the Np sensors. These
teeth permit measurement of the torsion or twist of the
shaft, which is proportional to output torque. The sensors are identical and interchangeable but serve different
functions.
The left-hand (10:30 o’clock) sensor provides an Np
signal to the ECU. This signal is used in the Np governing
circuitry and for generation of a cockpit tachometersignal. The right-hand (1:30 o’clock) sensor feeds the
ECU torque computation circuit and the Np overspeed
protection system. The sensors contain a permanent

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magnet and wire coil. They produce a pulse of current
each time a shaft or reference tooth passes (Figure 9-3).
TORQUEMETER INDICATING SYSTEM
Only a small part of the propulsive force produced
by a turboprop is caused by jet thrust. Neither turbine

specialized oil pressure tap on the engine inlet housing
and to a vent connection on the accessory drive gearbox
The 28-volt AC system operates the electrical circuit.
The Lycoming (T-55-L-712) torquemeter, a refined
torque measuring system, comprises five components
and interconnecting wiring. Two of these, the power

discharge pressure nor engine pressure ratio is used toindicate power produced by the engine. Instead a torquemeter measures the power level that the engine
develops on the ground and in flight. Consult the engine
maintenance manual for a description of torquemetersystem functions. In most systems, however, torquemeteroil pressure is used to actuate a torque indicating instrument in the aircraft. The torquemeter instrument
portrays torquemeter oil pressure (which is proportionalto engine power) in pounds per square inch. Some
torquemeter instruments are calibrated to read poundfeet of torque. Some may read shaft horsepower directly(Figure 9-4).
One torquemeter indicating system sometimes
called a torque pressure indicating system, contains apressure-indicator-type instrument. It provides continuous reading of engine output shaft torque supplied byan electrical transmitter mounted on the engine inlet
section. The transmitter is connected by hoses to a

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supply and indicator, are airframe-supplied components.
They may be replaced anytime without recalibration.
The remaining three (power output shaft, head assembly,
and junction box) are only replaced as a precalibrated set.
When torque is imposed on the engine power output
shaft, tension and compression stresses change the magnetic reluctance of the shaft. This system measures thechange of magnetic reluctance because of the torsionallyinduced forces (Figure 9-5).
The output shaft has welded to it a ferromagnetic
sleeve which serves as a transformer core in the headassembly. A nonrotating transformer, the head assembly
contains the primary coils which receive alternating current required to power the transformer from the power
supply. It also contains six sets of secondary coils orientedooat 90 to each and 45 to the shaft axis. They pick up
induced current and register the differential in current on
the dial face of the indicator. In operation, the headassembly surrounds the ferromagnetic sleeve on the shaft
which acts as the transformer core.
The junction box acts as the system conditioner andincludes key calibration controls. The potentiometers
standardize the signal level for indicator display andbalance out effects of magnetic nonhomogeneities. The
box also rectifies the secondary voltages from the headassembly.
The cockpit-mounted indicator reads the engineapplied torque in percent. A network located within theindicator computes the algebraic difference between thetwo secondary-induced voltages coming from the junction box.
The power supply is an airframe-mounted unit
which converts the airframe 28VDC signal to a constant
70VAC 2K Hz signal for system operation.
The General Electric T-700/701 torque-sensing system is a reference shaft that is pinned to the front end of
the drive shaft and extends to the aft end. It is free torotate relative to the drive shaft. The relative rotation isdue to output torque. The resultant phase angle between
the drive shaft teeth and reference teeth is electricallysensed by a pickup sensing the two teeth on the drive shaft
plus two reference teeth. The electrical signal is conditioned in the electrical control unit, which provides a DC
voltage proportional to torque (Figure 9-5). An intermediate power (1690 SHP), the output torque is 410
pound-feet, and the twist of the shaft is 7.4°.
9-4

EXHAUST GAS TEMPERATURE INDICATINGSYSTEM
Turbine engines may be instrumented for exhaustgas temperature indication at locations before, between,
or behind the turbine stages. Exhaust gas temperature isan engine operating limit and is used to monitor themechanical integrity of the turbines. It also checks engine
operating conditions. Actually, the temperature at the
turbine inlet is the important consideration. This is themost critical of all engine variables. However, it is impractical to measure turbine inlet temperature in mostengines. Consequently, thermocouples are inserted at
the turbine discharge. This provides a relative indicationof temperature at the inlet. The temperature at this point
is much lower than that at the inlet. It enables the pilot
to maintain surveillance over engine internal operation
conditions.
Several thermocouples are usually used. They arespaced at intervals around the perimeter of the engine
exhaust duct near the turbine exit. The exhaust gas
temperature indicator in the aircraft shows the averagetemperature measured by the individual thermocouples.
The readings of several thermocouples are usually obtained individually during ground engine maintenance bya selective switch. The spread between the lowest andthe highest thermocouple reading is useful in maintenance. It serves to indicate the presence of hot or cold
spots in the engine exhaust pattern which may indicate aproblem inside the engine. The importance of exhaustgas temperature cannot be overemphasized. Two systems that are in use on Army helicopters follow.
Thermoelectrical Systems (T-55-L-11E/T-55-L-712)
The chromel-alumel, thermoelectric (MGT)
measuring system is independent of all other engine
electrical wiring (Figure 9-6). The engine components
are five dual-probe segments; each probe is connected
externally so each segment can be operationally checkedindependently and replaced individually. The 10 thermocouple probes protrude into the gas stream at thepower turbine entrance (station 45). The probes react
to variations in temperature by developing a proportional
electromotive force across the chromel-alumel junction.
This potential difference results in meter deflectionsof the cockpit indicator. The cockpit indicator iscalibrated to read temperature in degrees centigrade.
In addition, this system incorporates two buss-bars
located on the aft fire shield. This allows for connectionof the five MGT segments to the single MGT harness.

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T-701 General Electric Thermocouple
The thermocouple harness is a one-piece assembly.
It consists of seven single-immersion, equally spaced
thermocouples for measuring power turbine inlet gas
temperature (Figure 9-7). The thermoelements are
made from special tolerance, oxidation-resistant,chromel-alumel wire. Each thermocouple junction issealed within a Hastalloy X sheath. The thermoelement
for the junction of each probe is continuous without joints
or splices from the junction to the harness output connector junction box where all outputs are parallel. The
harness output connector is hermetically sealed, has twoalumel contacts, and has two chromel contacts. The
average output of the seven probes provides the temperature signal to the ECU via the yellow cable. From the
ECU, the signal is relayed to the cockpit TGT indicatorand to the history recorder.
EXHAUST GAS TEMPERATURE INDICATORSThe indicator (Figure 9-8) in a typical exhaustgas temperature indicating system operates onelectrical potential from an engine thermocouple
harness. The probes are mounted in the aft sectionof the engine exhaust diffuser, The thermocouple
converts heat into electricity. The exhaust gas
temperature indicator (thermocouple-thermometer
9-6

indicator) is actually a sensitive millivoltmeter
calibrated in degrees centigrade. Its D‘Arsonval
movement is activated by an electrical force generated by
its relative thermocouple. The indicator circuit is entirely
independent of any other electrical power source. It
includes a coil resistor which provides instrument calibration.
FUEL-FLOW-RATE INDICATING SYSTEM
The fuel-flow-rate indicating system measures therate of fuel flow consumed by the engine. A typicalsystem consists of an indicator (dual indicator if a two
engine system), a fuel-flow meter, and a fuel transmitter,
which is an integral part of the fuel-flow meter located on
the engine. The fuel-flow-rate indicator, driven by adifferential autosyn, registers the rate of fuel flow to theengine in hundreds of pounds per hour. A four-digit
subtracting counter on the indicator can be set to readthe total pounds of fuel remaining in the aircraft. The
fuel-flow-meter transmitter is installed on a mount on theengine. The flow meter consists of a metering chamber
and a differential autosyn. As fuel travels from the maintank or valve manifold to the engine fuel control, it passes
through the metering chamber, moving a pivoted vane.
The pivot shaft of the vane is coupled to the rotor shaft ofthe differential autosyn. When fuel flows through themetering chambers, the differential autosyn rotates. ThisFM 1-506

autosyn rotation, in turn, causes the indictor autosyn torotate. This resuks in an indication of engine fuel consumption.
FUEL FLOW INDICATOR
The fuel flow indicator shows fuel flow in pounds (orkilograms) per hour to the fuel nozzles (Figure 9-9). Fuelflow is monitored for in-flight fuel consumption, checking
engine performance, and in-flight cruise control. The
relationship of abnormal fuel flow to other instrument
readings provides one of the best indicators for probablecause of engine malfunction.
FUEL PRESSURE INDICATING SYSTEM
A typical fuel pressure indicating system provides
continuous reading of fuel pressure (psi) in the main fuelsupply line from boost pumps in the tanks by an electricaltransmitter. The transmitter is connected to a tap on thevalve manifold. All fuel supply lines join the top electrical9-7

FM 1-506

transmitter to deliver fuel to the engine through the fuelcontrol inlet hose. Electricity is supplied to the transmitter by the 28-volt AC system.
FUEL INLET PRESSURE INDICATOR
Fuel system characteristics frequently make it advisable to monitor the fuel pump inlet pressure (Figure9-10). In case of fuel flow stoppage in flight, the sourceof difficulty should be located quickly. This determines
whether trouble developed in the engine or in the aircraft
fuel system. Prompt corrective action may then be taken.
In addition, the fuel pump inlet pressure indicates possible cavitation at the fuel pump inlet in flight. It will also
show whether or not the fuel system is operating properly
during engine ground checks.

ENGINE OIL PRESSURE INDICATINGSYSTEM
A typical engine oil pressure indicating systemprovides continuous reading of engine oil pump pressurein psi to the indicator. This is provided by an electricaltransmitter mounted on the engine inlet section. The
transmitter is connected to the 28-volt AC electricalsystem and by a hose to a pressure tap on the engine oil
falter housing.
OIL PRESSURE INDICATOR
To guard against engine failure resulting from inadequate lubrication and cooling of engine parts, oil supply
to critical areas must be monitored (Figure 9-11). The oilpressure indicator shows the pressure relayed by the oilpressure transmitter. On most installations, the oil pressure transmitter takes breather pressure into consideration, relaying the true pressure drop across the oil jets inthe oil system.

ENGINE OIL TEMPERATURE INDICATINGSYSTEM
In a typical engine oil temperature indicating system,
the indicator is electrically connected to the 28-volt DC
system. A electrical resistance-type thermobulb installed in the engine oil pump housing measures thetemperature of oil entering that unit. The temperature
reading is transmitted to the indicator in degrees centigrade. Two dissembler metals heat to electricity.
OIL-IN TEMPERATURE INDICATOR
The ability of the engine oil to lubricate and cool isa function of the oil temperature and the amount of oil

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supplied to the critical areas. An oil inlet temperature
indicator is frequently provided to show the oil temperature as it enters the engine-bearing compartments. Oil
inlet temperature also indicates proper operation of theengine oil cooler (Figure 9-12).

established on a thrust-measuring ground test standby
varying the exhaust nozzle area of the engine to achieve
the desired EGT.

AIR TEMPERATURE INDICATOR
The air temperature indications (Figure 9-13) currently used in aircraft are free air temperature (FAT),
outside air temperature (OAT), ram air temperature(RAT), total air temperature (TAT), and static air
temperature (SAT). Regardless of which temperature is
instrumented in a specific aircraft model, the flight
manual shows how to use it, along with applicable chartsor tables, to set the EPR values which provide rated thrustlevels. The EPR setting varies with the thrust level
desired and with the true TAT existing at the front of theengine (Tt2). Some aircraft have instrumentation whichindicates Tt2 values that may be used without correction
to determine EPR settings.
ENGINE THRUST INDICATION
The way a pilot sets and monitors the thrust
produced by the engine has been mentioned before(Figure 9-14). Thrust indication is discussed in detailbelow.
On some engines, engine RPM and exhaust gas
temperature (EGT) are used to indicate and set thrust.
On such engines, the pilot receives the full rated thrust of
the engine for takeoff at 100 percent RPM and a specified
EGT. The specified EGT at 100 percent RPM is

On centrifugal compressor engines, notably the
military J48, thrust is indicated by RPM alone. Full rated
thrust for takeoff is obtained when the tachometer reads
200 percent. The J48 has a fixed nozzle area which isestablished at manufacture. While there is an EGT limit
for takeoff and for the other engine ratings, a J48 will
normally operate at EGT below the applicable limit for
the thrust rating used by the pilot. If the EGT reaches theallowable limit, the engine deteriorates or malfunctions.
Most afterburning and non-afterburning turbojets
and turbofans with single or dual axial-flow compressors
use engine ratio to measure engine thrust. EPR indicators compare total turbine discharge pressure to total
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air pressure entering the compressor. EPR then indicates the ratio of these pressures. Engines instrumented
for EPR have a fixed exhaust nozzle area. Some military
afterburning engines have two fixed areas. One is used
for non-afterburning operation. A variable nozzle area
is used for some afterburning engines, but it varies only
while in afterburning. On both afterburning and nonafterburning engines, RPM and EGT may vary when theaircraft throttle is adjusted to obtain desired engine
thrust.
Some military afterburning models have exhaust
nozzles which are scheduled to vary the exhaust area
when the engine is running. Consequently, these engines
cannot be operated to EPR settings. They must be
controlled by throttle position with various engine
parameters checked to assure correct thrust output.
For engines with a fixed nozzle area, actual exhaustgas temperatures obtained during operation are usually
below prescribed limits. It is permissible for an engine tooperate at the temperature limit for any given thrust
rating. However, an engine that does may have a problem
which causes it to run abnormally hot.
With the exception noted in the use of thetachometer, engine RPM is considered a very inadequate
parameter for setting and checking engine thrust on
single and dual axial-flow compressor turbojet and turbofan engines having fixed exhaust nozzle areas. When
RPM is used as the controlling variable on such engines,
complications arise. The most important are –RPM does not provide an accurate means ofdetermining if the complete engine is functioning properly. High- pressure RPM on dual
axial-flow compressor engines and the RPM of
the whole compressor on single axial-flow compressor engines is governed by fuel control. For
example, RPM alone will not enable an engine
operator to detect a damaged or dirty compressor. RPM carefully used in conjunction withother engine variables such as fuel flow, exhaustgas temperature, and engine pressure ratio
allow for detection.
RPM for any given thrust condition will varyslightly among individual engines, depending
upon the engine trim speed. Engines aretrimmed by a fuel control adjustment toproduce full rated thrust at a fixed-throttle position on a standard day. The RPM variationmust be taken into account when RPM is used
to measure thrust being developed by the engine. This causes a complication which cannot

9-10

be tolerated when precise thrust settings arenecessary during flight.
On dual axial-flow compressor engines, RPM
variation of one percent causes approximately
four percent variation in thrust at the higher
thrust settings for the low-pressure compressor
rotor (N1) and five percent variation for thehigh-pressure compressor rotor (N2). One percent variation in turbine discharge pressure orengine pressure ratio results in only one andone-half percent variation in thrust. The five
percent variation in thrust for one percent variation in RPM also holds true for single axial-flow
compressor engines.
RPM does not vary in direct proportion tothrust produced by the engine over the entire
thrust range.
For these reasons, either turbine discharge pressure
or engine pressure ratio must be used as the engine
variable to indicate thrust on axial-flow compressor engines with fixed area exhaust nozzles. The use of either
is simpler undermost conditions and is more accurate.
MEASURING POWER AND THRUST FOR
TURBOPROPS
For engines other than those with fully variable exhaust nozzles, turbine discharge pressure, engine pressure, or engine pressure ratio can be used with goodresults to indicate or set engine thrust. They vary proportionally to the thrust the engine is developing. Most
turbojet and turbofan aircraft today are instrumented forengine pressure ratio. This is the parameter generally
used to set or measure engine thrust during takeoff, climb,
and cruise. For very accurate thrust measurement, suchas during ground trimming of an engine, turbine discharge pressure is often used to measure thrust. In suchcases, it is common practice to temporarily connect aturbine discharge pressure indicator to the engine.
ENGINE PRESSURE RATIO INDICATINGSYSTEM
In a typical engine pressure ratio indicating system,
the indicator is a dual-synchro instrument. The systemshows a constant reading of engine performance. This isdone by computing the ratio between the gas generator
discharge pressure and the inlet pressure of the engine.
These ratios are then transmitted to an indicator (bothindicators if a two-engine aircraft). Sample pressures aretaken from engine gas producer or gas generator ports
and from the pilot pressure system. The system includes
a transducer. The transducer includes a mountingFM 1-506

bracket and a transmitter unit. The transmitter unit
contains a multicell, diaphram-actuated computer; gear
train; two-phase motor; and transmitting synchros. The
indicators are graduated from 1.0 to 2.5 EPR units with
0.1 EPR markings and a vernier dial graduated in 0.01
EPR markings. An adjustable pointer on the face of theindicator is set to the maximum EPR for the ambient
temperature to indicate possible engine overspeed. The
maximum EPR for given temperatures is listed on the
instrument panel. The AC Power system supplies electrical power through the EPR circuit breaker.
TURBINE DISCHARGE PRESSURE
INDICATOR
This instrument indicates the internal engine pressure upstream of the jet nozzle, immediately aft of the laststage of the turbine (Pt5 to Pt7). It indicates pressure
available across the nozzle to generate thrust. Turbine
discharge pressure must be used in conjunction with T12
and P12.
ENGINE PRESSURE RATIO INDICATOR
This instrument (Figure 9-15) indicates the enginepressure ratio as a measure of thrust developed by theengine. This is the ratio of the turbine discharge totalpressure to the equivalent of the compressor inlet totalpressure (Pt5/Pt2 or Pt7/Pt2). Values for Pt2 must be
corrected for inlet duct loss on the engine pressure ratio

curves or charts by the aircraft manufacturer. Therefore,both for static (takeoff) and flight use, the actual value for
Pt2 will vary among different aircraft types and models
because of installation effects. However, the relation of
Pt2 at the engine face to Pam plus Pr is equivalent to totalpressure at or near the compressor inlet when airborne.
It is not advisable to instrument the compressor inlet
when airborne. It is not advisable to instrument thecompressor inlet directly for Pt2. The Pt2 sensor for thepressure ratio indicator may be placed at some otherlocation on the aircraft, preferably as near the engine air
inlet as possible. After appropriate corrections are madeto the in-flight charts in the flight of operation manual,
any rated thrust or percent of rated thrust may be set withthe aircraft throttle (as a function of the TAT or Tt2).
Pressure ratio between the pressure at the engine air
inlet and the discharge pressure at the jet nozzle indicatesthrust developed by the engine. Turbine discharge pressure alone is not an accurate indication of engine output.
Compressor inlet pressure (Pt2) must be taken into account on curves or charts when only turbine dischargepressure is instrumented.
For static engine operation, this is usually accomplished by showing barometric pressure (corrected
for inlet duct loss) rather than Pt2 values on the curves orcharts. In flight, curves or tables usually show airspeed
and altitude. This eliminates the need for actually

9-11

FM 1-506

delineating Pt2 values in the operating data. Enginepressure ratio indicators have the Pt2 value introduced
into the system. This factor is automatically taken into
account on the observed instrument reading. Except for

9-12

an indicator to measure engine thrust, the above represents the minimum adequate instrumentation for controlof the engine. Some installations may have additional
imtruments.

FM 1-506

GLOSSARYABBREVIATIONS AND ACRONYMS
ADU
abs
AC
AGB
API
atm
BDC
BHP
BTU
c
cal
CDU
cyl
db
DC
ECM
ECU
EDM
EGT
EPR
ESHP
F
FAT
FOD
ft-lb
fps
GE
hd
Hg
HP
HMU
hr
HUD
Hz
in
lb
LDS
LH
man
MGT
microsec
min
MPH
N
NATO

auxiliary drive unit
absolute
alternating current
accessory gearbox
American Petroleum Instituteatmospherebottom dead center
brake horsepower
British thermal unit
Celsiuscalibrated
central display unit
cylinderdecibeldirect current
electrochemical machining
electrical control unit
electric discharge machining
exhaust gas temperatureengine pressure ratio
equivalent shaft horsepower
Fahrenheit
free air temperatureforeign object damage
foot-poundfoot per second
General Electricheadmercury
horsepowerhydromechanical unit
hour
heads up display
Hertz
inch
pound
load-demand spindle
left-hand
manifold
measured gas temperature
microsecond
minutemiles per hournorthNorth Atlantic Treaty Organization

Glossary-1

FM 1-506

OAT
OD
ODV
OEI
PAS
PDU
PRT
PRV
psi
psig
PT
PTO
RAT
res
RH
RPM
s
SAT
SDC
SHP
TAT
TC
TCP
TDC
TEL
temp
TGT
TRQ
VG
VIDS
VS
wt

outside air temperature
outside diameter
overspeed and drain valve
one engine inoperative
power-available spindle
pilot display unit
power recovery turbinepressure regulating valve
pounds per square inch
pounds per square inch gage
power turbine
power takeoff
ram air temperature
resistanceright-handrevolutions per minutesouth
static air temperature
signal data converter
shaft horsepower
total air temperature
top center
tricresyl phosphate
top dead centertetraethyl leadtemperature
turbine gas temperature
torquevariable geometry
vertical instrument display systemversusweightDEFINITIONS
Acceleration lag – in the turbine engine, delay between the time instant power is requested and when power isavailable. The time it takes the engine to accelerate and give the required power increase.
Aerodynamic drag – force which thrust must overcome to move an aircraft forward. Design can lesson
aerodynamic drag through streamlining. Drag increases with increased speed.
Airbleed actuator – device that operates the interstage bleed system, to improve compressor acceleration
characteristics by unloading small amounts of compressed air.
Air density – total mass of air per given volume, the weight of a given volume of air. Air is denser at lower
altitude, at lower temperature, and lower humidity.
Airfoil profile – outline of an airfoil section.
Airfoil section – cross section of an airfoil parallel to a specific reference plane.
Air-fuel ratio – Ideal mixture of 15 parts of air to 1 part of fuel by weight; the mixture to be burned in the
combustion chamber.

Glossary-2

FM 1-506

Air inlet – large, smooth aluminum or magnesium duct at mouth of engine which conducts air into thecompressor with minimum turbulence and restriction.
Aluminum alloy – any of a variety of alloys formed by combining commercially pure aluminum with other metals
or materials such as copper, silicon, manganese, magnesium, and zinc. These alloying agents maybe added
singly or in combination to produce a metal with desired characteristics or strength, heat resistance, corrosion
resistance, and the like. Both wrought- and cast-aluminum alloys are widely used in the manufacture of rotor
blades and aircraft propellers.
Ambient – condition of atmosphere existing around the engine, such as ambient pressure or temperature.
Annular combustion chamber – two-part combustion chamber made up of an annular liner and a housing
assembly. The compressed air goes into a ring-shaped space formed by the annular liner around the turbine
shaft rather than into individual combustion chambers. The space between the outer liner wall and the housing
assembly allows the flow of cooling air. Used with axial-flow and dual compressors.
Annular reverse-flow engine – type of gas turbine engine most commonly used in Army aircraft. Airflow
direction is reversed in the combustion area.
Anti-icing system – device that supplies hot air under pressure to prevent icing of the inlet housing areas andinlet guide vanes. Hot scavenged oil is also circulated through internal passages in the walls and struts.
Army Spectrometric Oil Analysis Program (ASOAP) – periodic oil analysis for microscopic metal particles.
This takes place at an oil analysis laboratory.
Atmospheric pressure – barometric pressure exerted by the atmosphere as a result of gravitational attraction
above the point in question.
Atomizer – nozzle that creates minute particles and accurately shaped spray of fuel suitable for rapid mixing
and combustion.
Axial-flow compressor — one in which the air is compressed parallel to the axis of the engine. It is made up ofa series of alternating rotor and stator vane stages.
Bending – combination of tension and compression. The inside curve is under compression; the outside is under
tension.
Bleed system – device that unloads small amounts of air to relieve pressure.
Boss – raised rim around a hole; e.g., axle hole in a wheel. Circular projection on a casting, usually serving as
the seat for a bolt head or nut.
Brayton cycle – constant pressure cycle, with four basic operations which it accomplishes simultaneously andcontinuously for an uninterrupted flow of power. The turbine engine operates on this cycle.
Camber – the curvature of an airfoil’s surfaces from the chord line. It maybe positive, negative, or zero.
Can-annular combustion chamber – one with characteristics of both the annular types. It has an outer shelland a number of individual cylindrical liners.
Can combustion chamber – one made up of individual combustion chambers in which the air from thecompressor enters each individual chamber through the adapter.
Center of pressure – aerodynamic center of an airfoil; the point where all forces act.
Centrifugal axial-flow compressor– combination of the centrifugal-flow and the axial-flow compressor. It
usually consists of a five- or seven-stage axial-flow compressor and one centrifugal-flow compressor. Also called
the dual compressor.

Glossary-3

FM 1-506

Centrifugal-flow compressor – one with an impeller (rotor), stator, and compressor manifold. The rotor revolves
at high speed drawing air into the blades. Centrifugal lforce accelerates the air, and it moves through the stator and throughthe manifold.
Centrifugal twisting movement – force that tends to streamline rotating blades with the plane of rotation.
Choked nozzle – a nozzle whose flow rate has reached the speed of sound.
Chord line – imaginary line drawn between the leading and trailing edges of an airfoil.
Combustion – process of burning the fuel-air mixture in a gas turbine engine.
Combustion chamber – part of a turbine engine in which the propulsive power is developed by combustion of
the injected fuel and the expansive force of the resulting gases.
Combustion chamber liner – engine part usually constructed of welded high-nickel steel, subjected to flame ofextremely high temperature. It is behind the compressor and receives the compressed air which is mixed with
fuel and ignited. The combustor is where the combustion takes place.
Compressibility effects – the phenomenon encountered at extremely high speeds (near the speed of sound)when air ceases to flow smoothly over the wings (or blade) and piles up against the leading edge, causing extreme
buffeting and other effects.
Compressor – that section of an engine that produces an increase in air pressure. It is made up of rotating andstationary vane assemblies. It is the gas producer, or it maybe thought of as an air pump.
Compressor rotor – impeller, may be thought of as an air pump. It accelerates the air rearward into the firststage vane assemblies.
Compressor stall – separation of the airflow from the suction surface of the freed or rotating blades of acompressor. Any degree of stall reduces airflow.
Concave – pressure side of an airfoil.
Conduction – transfer of heat through material by communication of kinetic energy from particle to particle
rather than by a flow of heated material.
Convergent area – place where the cross-sectional area of a duct becomes smaller.
Convergent exhaust duct – duct used on fixed-wing aircraft; formed by tapering toward the rear of the duct.
Convex – suction side of an airfoil.
Crossover tube – duct carrying flame to the individual cylindrical liners of the can-annular combustion chamber.
Diffuser – aft structural member of an engine. It receives high velocity air from the centrifugal impeller anddecreases velocity and increases air pressure. In the combustor, a diffuser forms a divergent flow path for theexhaust gases.
Diffusion – process by which gases intermingle as the result of their spontaneous movement caused by thermal
agitation.
Directional references – specific definitions of terms referring to gas turbine engines to identify front and rear,right and left, bottom and top.
Divergent area – place where air flows from a smaller into a larger area.
Divergent exhaust duct – used on helicopter. Device to diffuse the exhaust gases rearward and to eliminate
thrust.
Drag parasite – drag caused by any member or structure which does not contribute to lift, such as engine
cowlings.

Glossary-4

FM 1-506

Drag profile – friction resistance produced by a member moving through the air; in simple terms, the “stickness”
of air against the surface of an airfoil.
Dry-cleaning solvent – cleaning compound that maybe used for all metal parts.
Dry-sump engine – one in which the oil is stored separate from the engine.
Dual compressor – see centrifugal-flow, axial-flow compressor.
Duplex nozzle – dual-orifice channel through which highly atomized and accurately shaped sprays of fuel go
into the combustion chamber.
Dynamic load – load on an aircraft due to a dynamic force.
Effective angle of attack – that part of a given angle of attack that lies between the chord of an airfoil and a line
representing the resultant velocity of the disturber airflow.
End play – longitudinal back-and-forth play of a shaft.
Engine airflow path – route of the airflow through the engine.
Engine oil pressure indicating system – device that gives continuous readings of engine oil pump pressure inpsi.
Engine oil temperature indicating system – device electrically connected to the 28 VDC system which transmits
temperature readings to the indicator in degrees centigrade.
Engine speed notation – the capital letter N, which represents the rotational speed of the engine. When anumber is placed after the N (as in “N1”), it indicates a specific system on the engine.
Engine stations – specific locations on the engine designating temperature or pressure-measuring locations.
For example, T3 means the third temperature pickup on the engine.
Engine surge – result of compressor stall. The complete engine in stall.
Exhaust – hot gases discharged from the engine through the exhaust diffuser section.
Exhaust diffuser – section composed of an inner and outer housing, separated by hollow struts across theexhaust passage. It forms a divergent flow path for the exhaust gases.
Exhaust gas temperature indicator – sensitive millivoltmeter calibrated in degree centigrade, activated by anelectrical force generated by its thermocouple.
Feedback – relay through the controls of aerodynamic forces exerted on the control surfaces and felt by thepilot.
Fir tree mount – manner of attaching the blades to the disk in the turbine rotor assembly. The root of the blade
where it is attached to the disk is shaped like a fir tree.
Foreign object – any object such as a tool, piece of equipment, engine part (nut, bolt, lockwire) that could in
any way damage the engine.
Foreign object damage – commonly called FOD, harm or destruction to the turbine engine caused by foreign
objects sucked into the inlet area of the engine with the required air.
Forged – shaped by hammering. Only the malleable metals are worked successfully. The application of heatincreases plasticity.
Free-power turbine engine – the turbine engine used by the Army. Sixty percent of the energy produced by
combustion is extracted by the gas producer turbine to drive the gas producer rotor. The rest of the energy isconverted to shaft horsepower to drive the output shaft of the engine.
Frictional loss – resistance to the relative motion of air flowing along a duct.

Glossary-5

FM 1-506

Frontal area – front part of a gas turbine engine, smaller than that of a reciprocating engine, therefore producing
less drag.
Front of engine – end from which power is extracted. An exception is the T73 engine on the CH-54, in which
the power is extracted at the end where the exhaust gas is expelled.
Fuel-air ratio – see air-fuel ratio.
Fuel atomizer – see atomizer.
Fuel controls – devices to control fuel flow. They are usually hydromechanical and include speed governors,
servo systems, valves, metering systems, and sensing pickups.
Fuel divider – device that meters fuel to the engine nozzles according to a predetermined schedule of secondary
flow versus primary flow.
Fuel nozzle – device to inject fuel into the combustion chamber in a highly atomized and accurately shaped
spray.
Fuel pressure indicating system – device that gives continuous readings in psi of fuel pressure in the main fuelsupply line.
Gas producer – the compressor in a free-power turbine engine.
Gas turbine engine – aircraft power plant that is adaptable for both airplanes and helicopters.
Gerotor pump – modified gear-type pump with two moving parts, an inner toothed element and an outer toothed
element. The inner one has one less tooth than the outer.
Glow plug – device that consists of a heating element in a short conventional-looking spark plug.
Ground angle – angle between the wing chord and the horizontal plane when the airplane is at rest on the
ground.
Heat exchanger – fuel-oil cooler, to help cool the oil. The exchanger is a cylindrical oil chamber surrounded bya jacket through which the fuel passes. Heat from the oil is transferred to the fuel by conduction.
Hot start – overtemperature during starting.
Hung start – failure to reach normal idling RPM during starting.
Igniter plugs – spark plugs which function only during starting and cut out of the circuit as soon as combustionis self-supporting.
Imbalance – uneven distribution of weight resulting in rotating parts being out of balance. Measured ininch-grams or inch-ounces.
Impeller rotor – rotor in a compressor that revolves at high speed, drawing air into the blades.
Induced angle of attack – that part of any given angle of attack in excess of the effective angle of attack.
Induced drag – the part of the total dragon an airplane induced by the airflow about the lifting surfaces.
Inlet guide vanes – devices positioned by the inlet guide vane actuator pilot valve. They are located in front of
the first compressor rotor, and they control the angle of incidence of the inlet air, thus ensuring a compressor
surge margin.
Inlet housing assembly – forward structural support of the engine.
Jam acceleration – rapid movement of the power lever, calling for maximum rate of rotor-speed increase.
Jetcal analyzer – device used to check the exhaust gas temperature during periodic maintenance inspections orwhen abnormally high or low temperatures are noted.

Glossary-6

FM 1-506

Jet propulsion – propulsion of a body by means of a jet or stream of gas, obtaining all or most if its thrust byreaction to the ejection of the combustion products (gas).
Jet silencer – a device used to reduce and change the lower-frequency sound waves emitting from the engine’s
exhaust nozzle, and thus reducing the noise factor.
Joule – unit of energy or work used in rating gas turbine ignition systems. A joule is equal to the amount ofenergy expended in one second by an electric current of one ampere through a resistance of one ohm.
Labyrinth seal – device for preventing leakage of gas on the gas generator shaft in a turbine. A labyrinth consistsof a series of projections on the rotating element running in close contact with grooves on the stationary element.
Maintenance allocation chart – chart in a -20 TM that assigns maintenance tasks to the lowest level capable ofdoing them, based on experience, skills, tools, and time available.
Mean aerodynamic chord – chord of an assumed rectangular airfoil representing the mean chord of an actual
airfoil.
Metal fatigue – weakening of metal due to microscopic changes in molecular structures caused by vibration or
exposure.
Micron – one millionth of a meter.
N1 system – gas producer.
N2 system – power turbine and shaft.
Nacelle – an enclosed shelter (cowling) on an aircraft for a power plant.
Nozzle – channel through which gas is conveyed to the rotor vanes of a turbine. Its purpose is to convert pressure
into velocity.
Orifice – opening having a closed perimeter through which a fluid may discharge. It may be open to theatmosphere, or it may be partially or completely submerged in the discharged fluid.
Otto cycle – a constant volume cycle, with four distinct operations performed intermittently. Reciprocating
engines operate on this cycle.
Outside air temperature – commonly abbreviated as OAT, the temperature of the air outside the engine.
Overspeed – RPM in excess of design; 100 percent RPM.
Overspeed governor, N2 – gearbox mounted on engine inlet housing and driven from the power shaft.
Overspeed governor, fuel control – part of the torquemeter system, an individual pumping unit which, with thetachometer drive assembly, sets the torquemeter oil pressure.
Overtemperature – temperature in excess of maximum allowable design temperature at the turbine exit.
PD 680 – cleaning solvent for exterior of engine and its attached components.
Planform – form or shape of an object, as of an airfoil, as seen in plan view or from above.
Power – the rate of doing work; work per unit of time.
Power-to-weight ratio – relationship between power and weight. Turbine engines produce more power forweight than reciprocating engines.
Power turbine (N2) – turbine that is free and independent of the gas producer system. It develops rotational
shaft power.
Pressure oil system – method of supplying oil under pressure to engine parts.
Pressure pumps – devices to put oil into the system.

Glossary-7

FM 1-506

Pressurizing and drain dump valve– device to prevent flow of fuel to the nozzle until enough pressure is built
up in the fuel control. One also drains the fuel manifold at engine shutdown and traps fuel in the upper portionof the system to keep the fuel control primed for faster starts.
Primary air – air that mixes with fuel in the combustion chamber to form a combustible mixture. The ratio is
15 parts of air to 1 part of fuel.
Radial inflow turbine – type of turbine made by some manufacturers, not used in any Army aircraft today, eventhough it is rugged and simple, relatively inexpensive, and easy to manufacture. Similar in design andconstruction to the centrifugal-flow compressor.
Ram – the amount of pressure buildup above ambient pressure at the engine’s compressor inlet due to theforward motion of the engine through the air (initial momentum of the air).
Ram air pressure – free stream air pressure, provided by the forward motion of the engine,
Ram ratio – the ratio of ram pressure to ambient pressure.
Ram recovery – the ability of an engine’s air inlet duct to take advantage of ram pressure.
Rear of engine – end of engine from which exhaust gas is expelled.
Reverse flow – change in direction of airflow in the combustion chamber of a gas turbine engine.
Rotational direction – direction of movement of the rotating part, determined by viewing the engine from therear.
Rotational speed – speed at which a propeller, rotor, or some other rotating part rotates; measured inrevolutions per minute.
Scavenge oil system – method of returning oil from the engine to the oil tank for cooling and reuse.
Scavenger pumps – those that drain oil from the sumps at various parts of the engine, return it through the oilcooler, and back to the oil tank.
Scoring – multiple scratches, usually parallel and resulting from the same cause.
Secondary air – large surplus of air that cools the hot sections of a gas turbine engine to lower temperatures.
Shaft horsepower (SHP) – energy used to drive the compressor and accessories in a turbine engine.
Shot peening – process used to work-harden metals.
Shroud – device used with turbine rotor to prevent blade tip losses and excessive vibrations. The shrouded
blades can be thinner than unprotected ones.
Simplex nozzles – single-orifice channels through which highly atomized and accurately shaped sprays of fuelgo into the combustion chamber.
Solvent immersion – cleaning method in which parts are immersed in solvent to remove carbon, gum, grease,and other surface contaminants.
Span – dimension of an airfoil from tip to tip or from root to tip.
Spar – any principal structural member in an airfoil running from tip to tip or root to tip.
Specific heat – the ratio of the thermal capacity of a substance to the thermal capacity of water.
Speed governor – device to relieve the pilot from resetting the power lever when outside air temperature andpressure change. Consists of flyweights balanced by a spring.
Splines – teeth in a gear.

Glossary-8

FM 1-506

Stable operation – condition where no appreciable fluctuation, intentional or unintentional, is occurring to anyof the engine’s variables such as RPM, temperature, or pressure.
Standard day conditions – 59ºF, sea level barometric pressure (29.92 inches of mercury).
Static-balanced surface – surface that is in balance about its hinge axis.
Stator – part of assembly that remains stationary with respect to a rotating part. Stator vanes are a stationary
set of airfoils in a compressor.
Stress – resultant condition of strain or pressure.
Tachometer generator – device that supplies power at a frequency proportional to the driven speed which drives
the synchronous motors in the indicator.
TBO – time between overhauls. This is established by the Army and the engine manufacturer.
Test cell – building, usually concrete, that contains both a control room and an engine room, used for testing
engines. The test cell is at the manufacturer’s; a mobile engine-test unit is used in the field.
Thermodynamic cycle – succession of processes which involve changes in temperature, pressure, and density
in which the substance acts as a means of transformation of energy. See Otto and Brayton cycles.
Thrust – a reaction force in pounds.
Thrust, gross – the thrust developed by the engine, not taking into consideration any presence of initial-air-mass
momentum.
Thrust, net – the effective thrust developed by the engine during flight, taking into consideration the initialmomentum of the air mass prior to entering the influence of the engine.
Thrust, reverser – a device used to partially reverse the flow of the engine’s nozzle discharge gases and thus
create a thrust force in the rearward direction.
Thrust, specific fuel consumption – the fuel that the engine must burn per hour to generate 1 pound of thrust.
Thrust, static – same as gross thrust without any initial airmass momentum present due to the engine’s static
condition.
Torque – a force, multiplied by its lever arm, acting at right angles to an axis.
Torquemeter – hydromechanical torque-measuring device located in the reduction-gear section of the inlet
housing. The measurement is read as torque oil pressure in psi.
Torquemeter indicating system – pressure indicator for continuous readings of engine output-shaft torque.
Transient conditions – conditions which may occur briefly while accelerating or decelerating or while passing
through a specific range of engine operation.
Turbine nozzle – stationary nozzle which discharges a jet of gas against the blades on the periphery of a turbine
wheel.
Turbine rotor – rotating portion of a turbine engine. It is made of specially alloyed steel because of severe
centrifugal loads, the result of high rotational speeds.
Turbine section – part of the turbine engine that extracts the kinetic energy of the expanding gases andtransforms it into shaft horsepower.
Turbojet – engine most commonly used in high-speed, high-altitude aircraft.
Vapor blasting – abrasive method used to clean combustor parts. Not to be used on ceramic, magnesium,
painted, or aluminum surfaces.
Vapor decreasing – cleaning method used on unpainted metal parts or aluminum-painted steel parts.
Glossary-9

FM 1-506

Vaporizing tubes – devices used instead of fuel nozzles in a T53-L-11 engine.
Variable inlet guide vanes – devices located in front of the first compressor rotor to guide the angle of incidenceof the inlet air to the first compressor rotor.
Vermatherm element – device which senses outlet fuel temperature and closes the core valve and opens the
bypass valve.
Vibration meter – device for measuring vibrations.

Glossary-10

FM 1-506

REFERENCES

SOURCES USED
These are the sources quoted or paraphrased in this publication.
Technical Bulletin (TB)
55-9150-200-24. Engine and Transmission Oils, Fuels and Additives for Army Aircraft. 30 July 1976.
Technical Manuals (TMs)55-1500-204-25/1. General Aircraft Maintenance Manual. 6 April 1970.
55-1500-345-23. Revision, Painting and Marking Army Aircraft. 12 June 1986.
55-1510-201-20. Organizational Maintenance Manual: Army Models U-8D, U-8G, RU-8D and U-8F
Aircraft. 9 January 1%9.
55-151O-2O4-23-1. Unit and Intermediate Direct Support Maintenance Manual: Army Models OV-1B/OV-lC
Aircraft. 15 November 1979.
55-1510-204-23-2. Aviation Unit and Intermediate Unit Maintenance Manual: Army Models: OV-lB/OV-lC
Aircraft. 15 November 1979.
55-151O-2O4-23-3. Aviation Unit and Intermediate Unit Maintenance Manual: Army Models: OV-lB/OV-lC
Aircraft. 15 November 1979.
55-1510-204-23-4. Aviation Unit and Intermediate Maintenance Manual: Army Models OV-lB/OV-lC
Aircraft. 15 November 1979.
55-1510-209-23-1. Aviation Unit and Aviation Intermediate Maintenance Manual For: Army Models U-21A,
RU-21A, and RU-21D. 31 December 1976.
55-1510-209-23-2. Aviation Unit and Aviation Intermediate Maintenance Manual: Army Models U-21A,
RU-21A and RU-21D. 31 December 1976.
55-1520-210-23-l. Aviation Unit and Intermediate Maintenance Instructions: Army Model UH-lH/V/EH-WX
30 September 1987.
55-1520-210-23-2. Aviation Unit and Intermediate Maintenance Instructions: Army Model UH-lH/V/EH-lH/X
Helicopters. 30 September 1988.
55-152021023-3. Aviation Unit and Intermediate Maintenance Instructions for Army Model UH-1H/V/EH-1
Helicopters. 30 September 1987.
55-1520-227-23-1,-3, -4,-5. Aviation Unit and Aviation Intermediate Maintenance Manual, Army Model
CH-47B and CH-47C Helicopters. 16 August 1978.
55-1520-227-23-2. Aviation Unit and Aviation Intermediate Maintenance Manual: Army Model Ch-47B
and CH-47C Helicopters. 16 August 1978.
55-1520-228-23-1,-2. Aviation Unit and Intermediate Maintenance Manual for Army Model OH-58A andOH-58C Helicopters. 28 February 1989.
55-1520-234-23-1. Aviation Unit and Aviation Intermediate Maintenance Manual Helicopter, Attack AH-lS(MOD). 30 September 1976.
55-1520-234-23-2. Aviation Unit and Intermediate Maintenance Manual Helicopter, Attack AH-lS(MOD). 30 September 1976.
55-1520-236-23-1. Aviation Unit and Intermediate Maintenance Manual: Army Model AH-lS (PROD),
AH-lS (ECAS) and AH-1S (Modernized Cobra) Helicopters. 8 May 1980.

References-1

FM 1-506

55-1520-236-23-2. Aviation Unit and Intermediate Maintenance Manual: Army Model AH-1P (PROD),
AH-lE (ECAS), and AH-lF (Modernized Cobra) Helicopters. 8 May 1980.
55-1520-236-23-3. Aviation Unit and Intermediate Maintenance Manual: Army Model AH-IS (PROD),
AH-lS (ECAS) and AH-lS (Modernized Cobra) Helicopters. 8 May 1980.
55-1520-236-23-4. Aviation Unit and Intermediate Maintenance Manual for Army Model AH-lS (PROD),
AH-lS (ECAS), AH-lS (Modernized Cobra, Helicopters. 8 May 1980.
55-1520-237-23-2. Aviation Unit and Aviation Intermediate Maintenance Manual Aircraft General Informs
tion Manual UH-60A and EH-60A Helicopters. 29 December 1978.
55-1520-237-23-6. Aviation Unit and Intermediate Maintenance Power Plant, Fuel, and Related Systems
Maintenance Task Manual UH-60A and EH-60A Helicopters. 29 December 1978.
55-2840-229-23-1. Aviation Unit and Intermediate Maintenance Manual for Engine Assembly, Model
T53-L-llC (NSN 2840-00-102-3%7), Part No. 1-000-080-12; Model T53-L-llD (2840-00-102-3968), Part No.
l-000-080-13 Model T53-L-13B (2840-00-134-4803), Part No. 1-000-060-10 and Model T53-L-703 (2840-00621-1860), Part No. 1-000-060-23 [TO 2J-T53-16-1]. 9 March 1981.
55-2840-229-23-2. Aviation Unit and Aviation Intermediate Maintenance Manual for Engine Assembly,
Model T53-L-llC (NSN 2840-00-102-3%7), Part No. l-000-080-12; Model T53-L-llD (2840-00-102-3968),
Part No. 1-000-080-13; Model T53-L-13B (2840-00-134-4803), Part No. 1-000-060-10 and Model T-53-L-703(2840-00-621-1860) Part No. 1-000-060-23 [TO 2J-T53-16-2]. 9 March 1981.
55-2840-231-23. Aviation Unit and Aviation Intermediate Maintenance Manual for Engine Assemblies,
Models T63-A-5A (NSN 2840-00-923-6023) and T63-A-700 (%40-00-179-5536). 27 February 1981.
55-2840-234-24/2. Organization, Direct Support and General Support Maintenance Manual: Engine, Shaft
Turbine (Model T55-L-llA) Part No. 2-001-020-05 (NSN 2840-00-428- 6382); (Model T55-L-llA-SA),PN 2-001-020-09 (2840-00-398-9671) (Model T55-L-llD), Part No. 2-001-020-14 (2840-00-000-0048).
6 August 1971.
55-2840-241-23. Aviation Unit and Aviation Intermediate Maintenance Manual: Engine, Aircraft, Gas Turbine
Model T63-A-720, Part No. 6887191 (NSN 2840-01-013-1339). 2 November 1977.
55-2840-248-23. Aviation Unit and Intermediate Maintenance Manual for Engine, Aircraft, Turboshaft,
Model T700-GE-700 and T700-GE-701 (TO 2J-T700-6). 28 April 1982.
55-2840-249-23. Aviation Unit and Aviation Intermediate Maintenance Manual: Engine Assembly, Model
T55-L-7 (NSN 2840-00-987-9717), T55-L-7B (2840-00-950-6875) and T55-L-7C (2840-00-9374480). 8 July 1980.
READINGS RECOMMENDED
These readings contain relevant supplemental information.
Army Regulations (ARs)310-25. Dictionary of United States Army Terms (Short Title: AD). 15 October 1983.310-50. Authorized
Abbreviations and Brevity Codes. 15 November 1985.
DA Pamphlet (DA Pam)
738-751. Functional Users Manual for the Army Maintenance Management System - Aviation (TAMMSA). 15 January 1988.
Field Manuals (FMs)
1-203. Fundamentals of Flight. 3 October 1988.1-563. Fundamentals of Airframe Maintenance. 24 June
1985.1-511. Army Aircraft Quality Control and Technical Inspection. 29 May 1987.10-68. Aircraft Refueling.
29 May 1987.
Technical Bulletins (TBs)
43-0197. Instructions for Safe Handling, Maintenance, Storage and Disposal of Radioactive Items Managed
by US Army Armament Materiel Readiness Command. 12 March 1982.

References-2

FM 1-506

750-103. Maintenance and Care of Common Type Machine Tools and Shop Equipment Before, During
and After Operation. 19 October 1%5.
Technical Manuals (TMs)9-243. Use and Care of Hand Tools and Measuring Tools. 12 December 1983.38-750-1. The Army
Maintenance Management System (TAMMS) Field Command Procedures. 29 December 1978.
Miscellaneous
MIL-T-5624. (CHAPTER 1). ASTM-D-1655. American Society for Testing Materials Specification,
Other Publications
Airframe and Powerplant Mechanics Airframe Handbook, EA-AC-65-15A. US Department of Transportation,Federal Aviation Administration, 1972.
Airframe and Powerplant Mechanics General Handbook, AC-65-9. Department of Transportation, Federal
Aviation Administration, 1970.
Airframe and Powerplant Mechanics Powerplant Handbook, EA-AC-65-12A. US Department of Transpor
taion, Federal Aviation Administration, 1976.
Pratt and Whitney Operating Instruction 200, The Aircraft Gas Turbine Engine and Its Operation Pratt and
Whitney. Traeger, Irwin E. Aircraft Gas Turbine Engine Technology. New York McGraw-Hill.

References-3

FM 1-506

Index-1

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Index-2

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Index-3

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Index-4

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Index-5

FM 1-506
30 NOVEMBER 1990

By Order of the Secretary of the Army:
CARL E. VUONOGeneral, United States Army
Chief of Staff
Official:
THOMAS F. SIKORABrigadier General, United States Army
The Adjutant General

DISTRIBUTION:
Active Army, USAR, and ARNG: To be distributed in accordance with DA Form 12-11-E, requirements for
FM 1-506, The Fundamentals of Aircraft Power Plants and Propellers (Qty rqr block no. 0729).

★ U.S. GOVERNMENT PRINTING OFFICE: 1994-300-421/02261