Aviation Machinist'

Aviation Machinist’s
Mate 3 & 2
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
About this course:
This is a self-study course. By studying this course, you can improve your professional/military knowledge,
as well as prepare for the Navywide advancement-in-rate examination. It contains subject matter about dayto-day occupational knowledge and skill requirements and includes text, tables, and illustrations to help you
understand the information. An additional important feature of this course is its reference to useful
information in other publications. The well-prepared Sailor will take the time to look up the additional
History of the course:
Sep 1991: Original edition released. Prepared by ADCS(AW) Terence A. Post.
Jan 2004: Administrative update released. Technical content was not reviewed or revised.
Published by
1. Jet Engine Theory and Design ...............................................................................
2. Tools and Hardware ...............................................................................................
3. Aviation Support Equipment..................................................................................
4. Jet Aircraft Fuel and Fuel Systems ........................................................................
5. Jet Aircraft Engine Lubrication Systems ...............................................................
6. Engine and Airframe Related Systems...................................................................
7. Helicopters and Turboshaft Power Plants ..............................................................
8. Turboprop Engines and Propellers.........................................................................
9. Power Plant Troubleshooting .................................................................................
10. Power Plant Inspection, Repair, and Testing .........................................................
I. Glossary .................................................................................................................
After completing this chapter, you will be able to:
State the theory of jet propulsion.
Identify the two types of engine designation
Identify the different types of engines and their
major assemblies.
Identify the common terms and variables
effecting engine performance.
Every rating or specialty has a language of
its own. The Aviation Machinist’s Mate is no
different. To be a good technician, you must learn
and understand the language (terms and theories)
necessary for a thorough understanding of your
specialty. With this basic understanding, you will
develop the skills to recognize, analyze, and
correct problems with jet engines. Without it, you
become a “parts changer” unable to recognize
possible reasons for the problem and analyze
the velocity of the water, giving us the term,
“a jet of water.”
Another example of the theory of jet propulsion is an inflated balloon. With the opening in
the balloon closed (fig. 1-1) there is no action
because the pressure of the gas inside the balloon
is equal in all directions. When you allow the
This chapter explains the basics necessary for
the Aviation Machininst’s Mate to build a strong
foundation. You’ll learn the theory, terms, types
of engines, and major parts of jet engines.
Jet propulsion is the propelling force generated in the direction opposite to the flow
of a mass of gas or liquid under pressure.
The mass escapes through a hole or opening called a jet nozzle. A familiar example
is the nozzle at the end of a fire hose. The
nozzle forms a smaller passageway through
which the water must flow. The nozzle increases
Figure 1-1.-Balloon example of restricting jet propulsion.
opening to release the air (fig. 1-2) the balloon
moves, Its movements appear to be in all directions. Actually, it is always moving in the opposite
direction from the open end where the air is
Let’s look at the balloon example from the
mechanics point of view. Igniting a hydrocarbon
fuel (compound containing only hydrogen and
carbon) and oxygen in a closed container (fig. 1-3)
releases heat. The burning fuel causes the trapped
gases to expand rapidly. Since the force of the
pressure is balanced, the container does not move.
Obviously, propulsion depends solely upon
internal conditions. The container does not “push
against” external air. In fact, a complete vacuum
would produce greater force. This is the basic
operating principle for all jets. The rocket
(propulsion unit) is one of the four main classes
of jet engines.
Before we move on to the physical principles
of jet engines, let’s review the basic principle to
the three other types of jet engines.
Suppose you attach a plain cylinder with open
ends under the wing of an aircraft flying at high
speed. Air enters the front of the duct and leaves
at the rear. Nothing increases the force of flow
through the duct. There is a loss of energy because
of skin friction and airflow disturbances at the
entrance and exit.
If you add heat energy to the air as it passes
through the duct, the air would expand and
increase the jet velocity. (Figure 1-5 shows a duct
When combustion takes place in a container,
the expanding gases rush out at a high velocity
(fig. 1-4). The release of internal pressure
at the nozzle end of the container leaves an
unbalanced pressure at the other end. The
released pressure propels the container (rocket)
in the direction opposite of the exhaust gases.
Figure 1-4.-Principle of jet propulsion.
Figure 1-2.-Balloon example of jet propulsion theory.
Figure 1-5.-Thermal duct with heat added externally to
accelerate the airflow.
Figure 1-3.-Combustion in a closed vessel.
heated externally by burning oil sprays.) The
amount of heat you can add is largely dependent
upon the pressure of the air treated. A simple
method of raising the pressure is to pass the air
through a DIVERGENT entry nozzle. A divergent
entry nozzle decreases the velocity of the air and
increases the pressure. This also provides a forward pressure wall for the jet to react. A
CONVERGENT exit nozzle further increases the
jet velocity. The simple gas unit (fig, 1-6) created
has little practical use because of the following:
The “intermittent impulse” jet engine (fig. 1-8),
known as the aeropulse or pulsejet improves
compression by sacrificing the principle of continuous power generation. The pulsejet is like the
ramjet, but with a series of nonreturn shutter
valves. Fuel injection nozzles located just aft of
the shutter valves provide fuel. As the engine
travels through the air, pressure on the nose opens
the valve and rams air into the duct, mixing air
with fuel. Igniting the combustible mixture creates
a high pressure (from the expanding gases), closing the valves. The violent ejection of the gases
forms a relatively low-pressure area inside the
duct, admitting a fresh charge of air through the
flat spring valves. Because of the temperature of
the duct and the return of part of the flaming
exhaust gases, the rest of the charges burn without
an igniter plug. This operating cycle or pulsations
creates a loud buzzing sound. “Buzz bomb”
described an early application of this unit, the
German V-1 flying bomb.
We learned the basic principle of jet propulsion with the rocket. The ram jet taught us that
adding heat would expand the gases and increase
velocity. It also showed the amount of heat that
is possible to add is dependent upon the amount
1. Air compression depends solely on “ram
2, A limited amount of heat is added.
3. Considerable heat is lost by radiation.
The next step is to improve the method of
adding heat, through internal combustion.
Figure 1-7 shows a divergent-convergent duct.
Fuel is injected and burned, releasing heat directly
into the airstream. This simple “Aero THermO
DYnamic Duct” (ATHODYD) or RAM JET is
used in remotely piloted vehicle (RPV) and cruise
Figure 1-8.-The aeropulse or pulsejet.
Figure 1-6.-A convergent discharge nozzle.
Figure 1-7.-The ramjet engine.
Air from the compressor section goes to the
combustion chamber. This is the area where fuel
and air are mixed and ignited. The burning of this
fuel/air mixture produces hot, expanding gases
that rush into the turbine rotors. The turbine
rotors attach to the same shaft as the compressor
rotors; so the turbine drives the compressor
making the engine self-sustaining. Finally, the
exhaust gases exit the engine (fig. 1-10) as jet
Now that you have a basic understanding of
jet propulsion, let’s look at the physical principles
of jet propulsion.
of air available. The pulsejet proved that the more
air an engine could compress the greater the power
(thrust) it produced. Now we will learn how a gas
turbine engine increases air compression to
develop the tremendous amount of thrust used in
modern aircraft.
The compressor is the greatest single reason
the gas turbine engine runs and produces the
thrust needed in modern aircraft. We will discuss
compressors in greater detail later, but for now
let’s keep with the basics. A basic axial-flow
compressor is a shaft with blades attached in rows,
called a rotor. When the rotor shaft turns, the
blades pull air into the engine. It is directed to
a set of stationary blades, called stator vanes
(fig. 1-9). The stator vanes direct air into the next
set of compressor vanes. Each set of blades and
vanes increases the compression of the air used
for combustion. The greater the number of stages,
the higher the compression ratio.
Physical principles govern the action of
matter, motion, force, and energy. You study the
actions in physics. An English scientist, Sir Isaac
Newton, stated three laws of motion explaining
jet propulsion. Another scientist, Bernoulli,
explains the principle behind the convergent/
divergent ducts we discussed earlier. These laws
and principles have words or terms with specific
meanings. Understanding the exact meaning of
the words is the key to understanding the
principles of physics. So let’s define some of the
basic terms of physics we need for a good
understanding of jet propulsion.
Force is the action or effect on a body that
changes the state of motion of the body. A force
may move a body at rest, or it could increase/
decrease the speed of a body, or change the
direction of motion. The application of a force
does not necessarily result in a change in motion.
A force is any push or pull acting on a body.
Water in a can exerts a force on the sides and
bottom of the can. A tugboat exerts a push or pull
(force) on a barge. A man leaning against a
bulkhead exerts a force on the bulkhead.
Matter is anything that occupies space and has
Mass is the quantity of matter in a body
measured in relation to its inertia. Mass and
weight are similar terms, and they are often
confused with each other. Weight is the common
measurement to determine the quantity of matter
wit h the pull of gravity on it. The following
example will help you understand the difference
between weight and mass. A person weighing 164
Figure 1-9.-Rotor and stator elements of an axial-flow
Figure 1-10.-Effects on airflow through an engine.
pounds on earth weighs 32 pounds on the moon
because of the difference in the pull of gravity.
Yet the mass of the person is exactly the same.
A mathematical formula for mass is as follows:
Mass is equal to the weight of the object divided
by the acceleration due to gravity, or
time. Negative acceleration is commonly called
Slug is an English measurement of mass. A
slug is mass with force and acceleration (due to
gravity) taken into consideration. So, a 1-pound
slug is the mass accelerated by 32 feet per second
per second (32 ft/sec2) when acted upon by a force
of 1 pound.
Acceleration due to gravity is 32.2 feet per second
per second (or feet per second squared).
This means that a free-falling 1-pound object
accelerates at 32.2 feet per second each second that
gravity acts on it. We use the lower case g to
express the acceleration due to gravity.
Standard day is a reference or standard.
Standard day shows conditions at sea level:
barometric pressure—29.92 inches of mercury
(Hg); temperature—59.0°F.
All gas turbine engines are rated with air at
the standard temperature of 59°F. Operation of
engines at a temperature above or below this
temperature will proportionally affect thrust output by as much as 15 or 20 percent. As the
temperature of a slug of air increases, the
molecules move faster. They run into each other
with more impact, and move further apart. This
decreases the density of the air. With the decrease
in density, the weight of the air is less, and the
thrust produced is proportional to the weight of
the slug of air.
Pressure effect is an increase in pressure,
resulting in more molecules per cubic foot, which,
in turn, increases the weight of the slug of air.
The weight of the air affects thrust output.
Ram effect is defined simply as more air
arriving at the engine intake than the engine can
ingest. Ram recovery is the airspeed at which ram
pressure rise is equal to friction pressure loss. This
speed varies with duct design factors. Mach 0.2,
or 150 miles per hour, is a representative reference
number for the beginning of ram effect.
Now let’s apply the terms we learned to the
principles of thrust.
Energy is the capacity for doing work.
Work is done when a force moves a mass
through a distance (work = force x distance). For
example, if you raise a 100-pound weight 10 feet,
1,000 foot pounds of work was done. The amount
of work is the same, regardless of how much time
(rate) is involved.
Power is the rate of doing work, or
In the above example, if the work was done
in 10 seconds, power expended was 100 foot
pounds per second. If it took 5 minutes (300
seconds), the rate of power is 3.3 foot pounds per
second. We often talk of power in terms of
Horsepower is the English measurement for
mechanical power and is 33,000 foot pounds per
minute, or 550 foot pounds per second. (Foot
pound is the energy required to lift a 1-pound
weight 1 foot in height.)
Speed is the distance a body in motion travels
per unit of time. Expressed in terms like miles per
hour (MPH) and feet per second.
Velocity is speed in a given direction. The
symbol V represents the term velocity.
Acceleration (a) is the rate of velocity change.
This definition is not based on distance traveled.
Acceleration is the gain or loss of velocity with
Newton’s first law states “A body (mass) at
rest tends to remain at rest, and a body in motion
turbine. As an AD, you will work with the gas
turbine engines. Under jet turbine engines are four
types: turbojet, turboprop, turbofan, and turboshaft. There are 2 engine designation systems in
use today to identify the different types of gas
turbine engines.
tends to move at a constant speed, in a straight
line unless acted upon by some external force.”
Newton’s second law states “An unbalance of
force on a body tends to produce an acceleration
in the direction of force, and that acceleration,
if any, is directly proportional to the force and
inversely proportional to the mass of the body.”
The law simply stated is “force is proportional
to the product of mass and acceleration.”
Turbojet engines are the basis for all other gas
turbine engines. We already know their cycle. Air
is drawn into the turbojet engine compressed,
mixed with fuel, and burned continuously.
The exhaust product of this burning operates
the turbine, for the compressor, producing
thrust which propels the aircraft. Adding fans,
propellers, or free turbines, change the basic
turbojet into a turboshaft, turboprop, or turbofan.
F = force, in pounds;
M = mass, in slugs; and
A = acceleration, in feet per second per second.
His third law states that “for every acting
force (action) there is an equal and opposite
reacting force (reaction).” We learned this
principle earlier with the rocket.
Turboshaft engines use the free turbine
principle. Exhaust gases drive the gas generator
turbine and a power turbine. The power turbine
(fig. 1-11) drives the helos rotor blades, through
drive shafts and gearboxes.
We learned the four basic types or classes of
jet engines: rocket, pulsejet, ramjet and jet
Figure 1-11.-Turboshaft engine.
Aeronautical (ANA) Bulletin No. 306. The new
system is under MIL-STD-879 and includes all
newly developed gas turbine engines of the Air
Force, Army, and Navy. ANA Bulletin No. 306
will remain in effect until all engines manufactured before the introduction of MIL-STD-879
have had a change to the configuration, the
performance or dropped from service.
Turboprops have a compressor, combustion
chamber (or chambers), turbine, and jet nozzle,
all of which operate in the same manner as their
counterparts in the turbojet. The difference is the
turbine. It sends increased power, generated by
the exhaust gases passing through additional
stages of the turbine, forward to a reduction gearbox (fig. 1-12). A propeller mounted on the gearbox provides thrust for the plane.
The turbofan gas turbine engine is the same
as a turboprop. Except a duct-enclosed, axial-flow
fan (fig. 1-1 3) replaces the gearbox and prop. The
fan is either part of the first-stage compressor
blades or mounted as a separate set of fan blades.
The fan uses 30 to 60 percent of the available
propulsive energy. The propulsive efficiency,
thrust, and specific fuel consumption for the
turbofan engine falls somewhere between those
of the turbojet and the turboprop engines.
The engine designation systems use standard
symbols to represent the type, manufacturer, and
model of aircraft engines used in military.
Knowing the designation systems gives you basic
information about the engine.
Two engine designation systems are in use
today. The old system is under Air Force-Navy
Figure l-13.-Turbofan engine. (A) With fully ducted fan;
(B) with short duct fan.
Figure 1-12.-T56 major engine components.
ANA Bulletin Number 306
that it will not perform satisfactorily under all
operating conditions. The designation applies to
an engine or device that has a specific function
rather than a general one, or that the engine has
only completed a 50-hour qualification test. Upon
satisfactory completion of the 150-hour qualification testing, the engine is approved as a J type.
As stated before, this bulletin is an Air ForceNavy bulletin. It is the older system, and it has
no provisions for Army designations.
Type Symbols
The first part of the designation consists of
a letter (or letters) with a number showing each
basic engine type.
The following are letter-type symbols:
The second part of the designation is a dash
and the manufacturer’s letter(s) symbol.
The following are some prominent engine
R . . . . . . . . . . Internal-combustion, air-cooled,
radial engine (reciprocating)
GE . . . . . . . . . . . . . . . . . . . . .. General Electric
J . . . . . . . . . . . Aviation gas-turbine (turbojet)
A . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. Allison
T . . . . . . . . . . Aviation gas-turbine (turboprop
and turboshaft engines for helicopters)
WE . . . . . . . . . . . . . Westinghouse Electric Co.
. . . . . . . . . . . . . . . . . . . . Pratt and Whitney
TF . . . . . . . . . Turbofan engine
W . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wright
PJ . . . . . . . . . Pulsejet engine
AC . . . . . . . . . . . . . . . . . . . . . .. Allis Chalmers
RJ . . . . . . . . . Ramjet engine
BA . . . . . . . . . . . . . . . . . . . .. Bell Aircraft Co.
RM............ Rocket motors
LA . . . . . . . . . . . . . . .. Lockheed Aircraft Co.
Following the letter symbol(s) is a number.
The Navy or-Air Force assigns numbers with the
letter. These numbers are arbitrary; they do not
represent any characteristics of the units involved.
The Navy uses even numbers beginning with 30.
The Air Force uses odd numbers beginning with
Although an engine may carry an odd or an
even number, it does not restrict the engine to the
service sponsoring or approving them. Engines
selection and use, by the various services, depends
on their satisfactory application for a particular
MD . . . . . . . . . . . . . . . McDonald Aircraft Co.
V . . . . . . . . . . . . . . . . . . . . . . . . . . . .. Packard
The third part of the designation consists of
a dash and a numeral showing the model number.
Model numbers, like type symbols, use odd
numbers for Air Force models and even numbers
for Navy models. Air Force model numbers for
each type of jet engine begin with 1 and continue
with consecutive odd numbers. Navy numbers
begin with 2 and continue with consecutive even
numbers. Model numbers show the service that
approves the model. The Naval Air Systems
Command approves all even model numbers. The
Air Materiel Command approves all odd model
A given engine design has only one type and
model designation for both services. For example,
if the Navy uses an Air Force-approved engine
without any changes, the Navy uses the numbers
assigned. If the Air Force uses a Navy-approved
The letter X or Y before the basic designation
shows an experimental or restricted engine.
The letter X used as a prefix show experimental and service test of a particular engine. After
exhaustive tests confirming the ability of an engine
to perform under all operating conditions, you
remove the X prefix.
The letter Y shows restricted service designation. Its application is self-explanatory in the sense
Type letter symbols:
engine that requires minor changes, the Air Force
uses the Navy-type designation and assigns its own
model designation. It begins with the number 1
and progresses with consecutive odd numbers to
the modified engine. This model number is
actually a modification number that tells which
service made the last production change in it for
a particular aircraft.
J . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbojet
. . . . . . . . . . . . . . . . . . Turboshaft,
F . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbofan
For example, J79-GE-10 breaks down as
Each service consecutively assigns type
numbers with the type letter symbol. Beginning
type numbers for the three armed services are:
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Turbojet
100. . . . . . . . . . . . . . . . . . . . . . . . . Air Force
79 . . . . . . . . . . . . . . . . . . . Air Force developed
400 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Navy
700 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Army
GE . . . . . . . Manufactured by General Electric
10 . . . . . . . . . . . . . . . . . . . . . . . . . . Navy model
Manufacturer’s Symbols
The second part of the designation consists of
a dash and a two-letter symbol showing the
manufacturer, as follows:
T . . . . . . . . . Turbine (turboprop in this case)
AiResearch Division, Garrett Corp. . . . . . GA
56 . . . . . . . . . . . . . . . . . . . . . . . Navy sponsored
Allison Division,
General Motors Corp . . . . . . . . . . . . . . . . ..AD
A . . . . . . . . . . . . . . . . Manufacturer (Allison)
Continental Aviation
and Engineering Corp . . . . . . . . . . . . . . . . ..CA
16 . . . . . . . . . . . . . . . . . . Eighth model (Navy)
General Electric Company . . . . . . . . . . . . . GE
Lycoming Division, Avco Corp. . . . . . . . . LD
The new system applies to all the armed
services—Army, Navy, and Air Force. It consists
of three parts; a type indicator, manufacturer’s
symbol, and a model indicator. All newly
developed gas turbine engines use this designation.
Existing engines change to a new three-digit model
number with major changes in configuration or
design. However, the engine usually keeps the old
two-digit type indicator.
Pratt and Whitney Aircraft Division,
United Aircraft Corp . . . . . . . . . . . . . . . . . .PW
Rolls Royce, Ltd . . . . . . . . . . . . . . . . . . . . . .RR
United Aircraft of Canada, Ltd. . . . . . . . . CP
Curtiss-Wright Corp . . . . . . . . . . . . . . . . . . .WA
Special manufacturer symbols show when two
manufacturers are jointly producing an engine.
In that case, the manufacturer symbol consists of
one letter from the symbol for each manufacturer.
Special designations (X or Y) used in the new
system are the same as the old system.
Type Indicator
NOTE: Manufacturer letter symbols are
not the same as manufacturer code symbols. The code symbols are with the
Federal Supply Code for manufacturers.
The type indicator, the first part of the
designation, consists of the appropriate type
letter symbol together with the type numeral.
The following are examples of the various
Model Indicator
The model indicator consists of a dash and a
model number, or a dash and a model number
with a suffix letter you assign, and a model
number for each configuration of a given engine.
Each service has a block of numbers they use
consecutively, just like the type indicator number.
The following is the beginning number for each
F = Turbofan
401 = Second Navy turbofan in new designation system
PW = Pratt and Whitney Aircraft Division,
United Aircraft Corporation
100 . . . . . . . . . . . . . . . . . . . . . . . . . . .. Air Force
400 = First Navy model of this particular
400 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Navy
Mixed designation
700 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Army
NOTE: Should any service use another’s
designated engine, the designation remains
the same unless there is a model change.
In this case, only the model indicator is
changed, showing the engine has been
416 = Shows the incorporation of configuration and performance changes to the
Table 1-1 lists by engine type some of the
more common designations and their associated
Table 1-1.-Navy Aircraft and Associated Engines
F-5, T-38, T-2C
K F I R (2) ( F - 2 1 A )
A-6E, EA-6B, KA-6D, A-4, TA-4
F/A- 18
F- 14D
A-7E, TA7C
S-3A, US-3A
T-56-A- 14
P-3, C-130, C-2, E-2
T - 3 4( 2 )
H-l J
H-3 , H-2, H-46
drops. Divergent inlet designs change ram air
velocity into high static pressure at the compressor
inlet. Friction due to air passing through the duct
surfaces and bends in the duct cause pressure
drops and differences. Performance achieved
through proper duct design is only half the story.
Careful construction and maintenance is essential
to maintain designed performance. Small amounts
of airflow distortion result in loss of engine
efficiency and unexplainable compressor surges.
This is caused by poor sheet metal work,
protruding rivet heads, or poor welds.
Navy/Marine aircraft. You will notice that
included with the turboprop engines are two
civilian engines, which have the manufacturer’s
There are many different models of jet engines
in the Navy today. Developments over the years
has produced a more efficient engine, both from
a performance and maintenance point of view.
These modern engines are more complex, but they
still operate according to the same basic principles.
This section discusses the major parts found in
various gas turbine engines; their name
(nomenclature), construction, purpose and
operating characteristics. Major components of
all gas turbine engines are basically the same. The
nomenclature of the various engines in current
use, however, may vary slightly because of
differences in manufacturers’ terminology.
Through experience and reading the many and
varied publications, the mechanic recognizes the
engine components, regardless of the terminology
Engine inlet ducts take a variety of shapes,
depending on the position of the engine and
purpose of the aircraft. Two methods of
classifying inlet ducts are as follows:
1. Single entrance and divided entrance
2. Subsonic and supersonic ducts
Single Entrance/Divided Entrance
The simplest type of air ducts are on engines
mounted in pods under the wing. The single
entrance (fig. 1-14, view A) gets maximum ram
pressure through the straight flow. It’s used where
an unobstructed entrance lends itself readily to
a single, short, straight duct.
A turbojet engine consists of the following
sections and systems:
Some aircraft, because of fuselage design or
internal parts, have to use a dual or divided
entrance, as shown in figure 1-14, view B. These
dual entrances are in the wing root, or in scoops
on the side of the fuselage. Either type presents
more problems to the aircraft and engine designers
than the single entrance. Problems are caused by
boundary layer airflow and the difficulty in
obtaining enough entrance area without creating
too much drag.
1. Air entrance section
2. Compressor section
3. Combustion section
4. Turbine section
5. Exhaust section
6. Accessory section
7. Systems necessary for starting, lubrication,
fuel supply, and auxiliary purposes, such as antiicing, cooling, and afterburning
Subsonic/Supersonic Ducts
The air entrance directs incoming air to the
compressor entrance with a minimum of energy
loss. Additionally, it must deliver this air under
all flight conditions with as little turbulence and
pressure variation as possible.
Normally, the engine inlet is part of the airframe. Because of its important part in engine
performance we include it here. Proper duct
design contributes to aircraft performance by
increasing ram recovery and limiting pressure
Modern Navy aircraft capable of supersonic
flight pose another problem to aircraft designers
because the airframe can withstand supersonic
velocity air, but the engine cannot. There are two
methods commonly used to diffuse the intake air
and slow its flow to subsonic speeds during supersonic flight. One is to create a shock wave in the
intake airstream, which will disrupt the flow and
cause a decrease in velocity. The other method
the burners in the quantity and at the pressures
required. A secondary function is to supply
compressor bleed air for various purposes in
the engine and aircraft. The compressor provides
space for mounting accessories and engine
There are two basic types of compressors. The
compressor type is also the engine type, so a
centrifugal-flow compressor is in a centrifugal
engine. Centrifugal-flow compressors have a
compression ratio of 5:1. Present-day axialflow compressors have compression ratios
approaching 15:1 and airflows up to 350 lb.
The addition of a fan raises these values to 25:1
and 1,000 lb/sec.
Centrifugal-Flow Compressors
The single entry centrifugal-flow compressor
(fig. 1-15) consists of an impeller (rotor element),
a diffuser (stator element), and a manifold. The
impeller picks up and accelerates air outward to
the diffuser. The diffuser directs air into the
manifold. The manifold distributes air into the
combustion section. Double entry centrifugal-flow
Figure 1-14.-Air entrance designs. (A) Single entrance;
(B) dual entrance; (C) variable entrance.
is to vary the area, or geometry, of the intake
duct. Navy aircraft use this method, incorporating
movable ramps as shown in figure 1-14, view C,
to change the area and shape of the intake
The primary function of the compressor
is to supply air in enough quantity to satisfy
the requirements of the combustion burners.
Specifically, the compressor increases the air mass
received from the air inlet duct and directs it to
Figure 1-15.-(A) Elements of the centrifugal compressor;
(B) air outlet elbow with turning vanes for reducing air
pressure losses.
cascade effect is of prime importance in determining blade design and placement.
compressors (fig. 1-16) handles the same airflow
with a smaller diameter. Small multi-stage
centrifugal-flow engines used in aircraft (fig. 1-17)
or as APUs take advantage of these features.
The axial-flow compressor has its disadvantages, the most important of which is the stall
problem. If, for some reason, the angle of
attack—the angle at which the airflow strikes the
rotor blades—becomes too low, the pressure
zones, shown in figure 1-18, will be of low value,
and the airflow and compression will be low. If
the angle of attack is high, the pressure zones will
be high, and airflow and compression ratio will
be high.
Axial-Flow Compressors
The term axial flow applies to the axial
(straight-line) flow of air through the compressor
section of the engine. The axial-flow compressor
has two main elements—a rotor and a stator.
Each consecutive pair of rotor and stator blades
makes a pressure stage. The rotor is a shaft with
blades attached to it. These blades impel air rearward in the same manner as a propeller, by reason
of their angle and airfoil contour. The rotor,
turning at high speed, takes in air at the
compressor inlet and impels it through a series of
stages. The action of the rotor increases the
compression of the air. At each stage it accelerates
rearward through several stages. The stator blades
act as diffusers at each stage, partially converting
high velocity to pressure. Maintaining high
efficiency requires small changes in the rate of
diffusion at each stage. The number of stages
depends on the amount of air and total pressure
rise required. The greater the number of stages,
the higher the compression ratio. Most presentday engines use from 10 to 16 stages.
An axial-flow compressor follows the same
rules and limitations of an aircraft wing. The
concept is more complicated than a single airfoil,
because the blades are close together. Each trailing
edge blade affects the next leading edge. This
If the angle of attack is too high, the
compressor will stall. The airflow over the upper
foil surface will become turbulent and destroy the
pressure zones. This will decrease the compression
airflow. The angle of attack will vary with engine
rpm, compressor-inlet temperature, and compressor discharge or burner pressure. Any action
that decreases airflow relative to engine speed will
increase the angle of attack and increase the
tendency to stall. The decrease in airflow may
result from a too-high compressor-discharge
During ground operation of the engine, the
prime action that causes a stall is choking. If there
is a decrease in the engine speed, the compression
ratio will decrease with the lower rotor velocities.
With a decrease in compression, the volume of
air in the rear of the compressor will be greater.
This excess volume of air causes a choking action
in the rear of the compressor with a decrease in
airflow. This, in turn, decreases the air velocity
in the front of the compressor and increases the
tendency to stall. If no corrective action is taken,
the front of the compressor will stall at low engine
Another reason for engine stall is high compressor inlet air temperatures. High-speed aircraft
may experience an inlet air temperature of 250°F
because of ram effect. These high temperatures
cause low compression ratios (due to air density
changes) and will also cause choking in the rear
of the compressor. This choking-stall condition
is the same as the stall condition caused by low
compression ratios due to low engine speeds.
Each stage of a compressor should develop the
same pressure ratio as all other stages. When the
engine slows down or the compressor inlet air
temperature climbs, the front stages supply too
much air for the rear stages to handle, and the
rear stages will choke.
Figure 1-16.-Impeller with inducer vanes as separate pieces.
There are five basic ways manufacturers
can correct this front-end, low-speed, hightemperature stall:
1. Lowering the angle of attack on the front
stages so the high angles at low engine speed are
not stall angles
2. Installing a bleed valve in the middle or rear
of the compressor to bleed air and increase airflow
in the front of the compressor at low engine speeds
3. Splitting the compressor into two rotors
and designing the front rotor rpm to decrease
more than the rear rotor at low speeds, so low
front-rotor speed will equal the low choked
4. Installing variable inlet-guide vanes and
variable stators in the front of the first series of
compressor stages so the angle of attack is
changed at low engine speed
5. Using a variable-area exhaust nozzle to
unload the compressor during acceleration
NOTE: A combination of any of the above
may be used.
Figure 1-18.-The cascade effect.
The stator has rows of blades or vanes
dovetailed into split rings and attached inside an
enclosing case. The stator vanes project radially
toward the rotor axis and fit closely on either side
of each stage of the rotor.
The compressor case, into which the stator
vanes fit, is horizontally divided into halves.
Figure 1-17.-Centrifugal-flow engine (J33).
Either the upper or lower half is removed for
inspection or maintenance of the rotor and stator
The function of the vanes is twofold. They
receive air from the air inlet duct or from each
preceding stage of the compressor. It is delivered
to the next stage or to the burners at a workable
velocity and pressure. They also control the
direction of air to each rotor stage to get the
maximum compressor blade efficiency.
The rotor blades are in front of the inlet guide
vane assembly. The guide vanes impart a swirling
motion to the air entering the compressor in the
direction of engine rotation. This motion
improves the aerodynamic characteristics of the
compressor by reducing the drag on the first-stage
rotor blades. The inlet guide vanes are curved and
airfoil shaped. The vanes are made of steel alloy,
many with a protective coating to prevent erosion.
They are welded to steel inner and outer shrouds.
The variable inlet-guide vanes are fitted and
pinned to spherical bearings that are retained in
the compressor front frame.
At the discharge end of the compressor, the
stator vanes straighten the airflow to cut
turbulence. These are straightening vanes or the
exit guide vanes.
The casings of axial-flow compressors support
the stator vanes and provide the outer wall of the
axial path the air follows. They also tap off
compressor air for various purposes, such as
cockpit pressurization and heating, or fuel tank
pressurization. There are outlet ports for bleeding
off compressor air at different stages, depending
on the pressure or temperature desired. (The
temperature rises proportionately with pressure
The stator vanes are made of steel with
corrosion- and erosion-resistant qualities. Frequently they are enclosed by a band of suitable
material to simplify the fastening problem. The
vanes are welded into the shrouds; then, the outer
shroud is secured to the compressor housing inner
wall by radial retaining screws.
The rotor blades are made of stainless or
semistainless steel. Methods of attaching the
blades in the rotor disc rims vary in different
designs. They commonly fit into discs by either
bulb (fig. 1-19) or fir-tree (fig. 1-20) type roots.
The blades then lock by grub screws, peening,
locking wires, pins, or keys.
Compressor blade tips reduce in thickness by
cutouts, and are referred to as blade “profiles.”
These profiles allow rubbing, when rotor blades
come into contact with the compressor housing
Figure 1-19.-Bulb root-type rotor blades.
Figure 1-20.-Fir-tree root-type rotor blades.
or shroud without serious damage. This condition
may occur if rotor blades become excessively loose
or by reduction of rotor support by a malfunctioning bearing. Even though blade profiles reduce
such chances, occasionally a blade may break
under duress of rubbing and cause considerable
damage to compressor blades and stator vane
The blades vary in length from entry to
discharge. The annular working space (drum to
casing) reduces progressively toward the rear by
the increase in the rotor drum diameter. The
rotor may feature either drum-type or disc-type
The drum-type rotor is machined from a single
aluminum alloy forging. Dovetail grooves are
machined around the circumference of the drum
for blade retention. Provisions for bearing
supports and splined drive shafts are on the front
and rear faces of the drum.
The disc-type rotor consists of separately
machined discs and spacers flanged to fit one
against the other in sequence. The entire assembly
is held together by through-bolts, tie-bolts, or
bolted individually to one another. Blades may
be attached to the disc rim by the dovetail or bulb
design locking feature. Similar provisions to those
on the drum-type assembly are made for bearing
supports and splined drive shafts.
Another method of rotor construction is to
machine the discs individually, and shrink fit the
discs over a steel drive shaft (heating the disc and
freezing the shaft to assemble the rotor).
However, this type of compressor construction
is only satisfactory for compressors where rotor
and centrifugal stresses are relatively 10 w.
The drum and disc-type rotor assemblies are
shown in figures 1-21 and 1-22. Many engine
designs now use combination disc and drum
compressor rotor assemblies due to their splitspool design concept. The F404-GE-400 and the
TF34-GE-400 are examples of the combination
compressor rotor assembly.
The coverage of axial-flow compressors up to
this point has dealt solely with the conventional
single-rotor type. Actually, there are two
configurations of the axial compressor now in use.
The single rotor and the dual rotor, sometimes
referred to as solid spool (fig. 1-23) and split spool
(fig. 1-24).
One version of the solid-spool compressor uses
variable inlet guide vanes. This is the arrangement
found on the J79-GE-10 engine. The engine has
a 17-stage compressor. The angles of the inlet
guide vanes and the first six stages of the stator
vanes are variable. During operation, air enters
the front of the engine. Air is directed into the
compressor at the proper angle by the variable
inlet guide and variable stator vanes. The air
is compressed and forced into the combustion
section. A fuel nozzle extending into each
combustion liner atomizes the fuel for combustion. These variables are controlled in direct
relation with the amount of power the engine
requires to produce the pilot’s power lever
One version of the split-spool compressor is
in Pratt and Whitney’s J52 engine. It uses two
Figure 1-21 .-Drum-type compressor rotor.
Figure 1-23.-Solid-spool compressor—single rotor turbine.
Figure 1-24.-Split-spool compressor—dual rotor turbine.
Figure 1-22.-Disc-type compressor rotor.
4. Directing the hot gases to the turbine
compressors with their respective turbines and
interconnecting shafts that form two independent
rotor systems.
The axial-flow type of engine has definite
advantages. The advent of the split-spool axial
compressor made these advantages even more
positive by offering greater starting flexibility and
improved high-altitude performance.
The advantages of the axial-flow compressor
are as follows:
The location of the combustion section is
directly between the compressor and the turbine
sections. The combustion chambers are arranged
coaxially with the compressor and turbines. The
chambers must be in a through-flow position to
function efficiently.
About one-fourth of the air entering the
combustion chamber area mixes with the fuel for
combustion. This is primary air. The remaining
air (secondary air) serves as flame control.
Keeping the temperature of the heated gases down
to a level at which the liners, turbine nozzles, or
blades will not burn. These basic requirements
apply to all combustion sections.
Another general requirement of combustion
chambers is air pollution emission reduction.
Pollution emissions are particles of matter, such
as smoke, carbon monoxide, partially burned
hydrocarbons, and nitric oxides. In general,
exhaust smoke becomes a problem when
combustors operate at pressure greater than 10
atmospheres and when the fuel-air ratio in the
primary zone of the combustor is rich. For
example, in the idle range of operation, both
smoke particles and partially burned hydrocarbons emit. During the combustion process,
emission levels of nitric oxide increase with
temperature increases to about 2,600°F. At this
temperature, these emission levels begin to taper
off. There is research being conducted to correct
problems, but many new factors may influence
the solution of the pollution problem. Present
approaches to reducing exhaust emissions include:
1. High peak efficiencies
2. Low frontal area for given airflow
3. Straight-through flow, allowing high ram
4. Increased pressure rise by increasing the
number of stages, with negligible losses
The disadvantages of the axial-flow compressor are as follows:
1. Good efficiencies are possible over narrow
rotational-speed range only
2. Difficulty of manufacture and high cost
3. High starting-power requirements
Relative to compressors, the fan of a turbofan
engine should be mentioned now. The fan
accelerates a large mass of air rearward. It requires
relatively low drive power, and has a pressure ratio
of 2 to 1 or less. It can be thought of as a
precompressor, as air enters the compressor inlet
at a pressure about 1.5-2.0 to 1 atmosphere.
1. Cut visible smoke by improving primary
zone fuel-air mixing, but not sacrifice altitude
relight capability.
2. Reduce carbon monoxide and unburned
hydrocarbon emissions. Increase fuel atomization
and optimizing the fuel-air ratio in the primary
combustion zone.
3. Reduce nitric oxide emissions by minimizing the amount of time the fuel-air mixture spends
in the combustor (by using short cans) or lowering
the temperature in the primary zone of
Remember that these approaches are not
solutions. Only ideas to give you some insight to
the problems and possible solutions.
All combustion chambers contain the same
basic elements: a casing, a perforated inner liner,
a fuel injection system, some means for initial
The combustion section provides the means
for and houses the combustion process. Its
function is to raise the temperature of the air
passing through the engine. This process releases
energy contained in the air and fuel. The major
part of this energy drives the compressor. The
remaining energy creates the reaction (or
propulsion) and passes out the rear of the engine
in the form of a high-velocity jet.
The primary function of burning the fuel-air
mixture includes:
1. Providing the means for proper mixing of
the fuel and air to assure good combustion
2. Burning this mixture efficiently
3. Cooling the hot combustion products to a
temperature that the turbine blades can withstand
under operating conditions
ignition, and a fuel drainage system to drain off
unburned fuel after engine shutdown.
The three basic types of combustion chambers
are as follows:
dome. When internal mounting at the liner dome,
the chamber cover is removed for replacement or
maintenance of the nozzle.
The simplex nozzle, with its single orifice, does
not provide a satisfactory spray over a wide range
of operating conditions. Therefore, its use on
current models of jet engines is limited.
The duplex nozzle has good spray characteristics. Its use does require a pressurizing valve
(flow-divider) to divide flow to the primary and
main fuel manifolds. During starting and idling,
the small primary orifice of the duplex nozzle
provides a high degree of atomization under low
pressures. As sufficient pressure builds, the
pressurizing valve opens the main line; the larger
orifice supplies increased fuel in a atomized form.
Newer engines use single-or multiple-unit duplex
nozzles for satisfactory sprays under various
operating conditions.
The cross-ignition tubes are a necessary part
of the can-type combustion chambers. Since each
of the cans is in reality a separate burner, each
operates independently of the other. Combustion
is spread during the initial starting operation by
simply interconnecting all the chambers. As the
flame is started by the spark igniter plugs in the
two lower chambers, it passes through the tubes
and ignites the combustible mixture in the
adjacent chambers. This process, similar to the
action of a pilot light on a gas stove, continues
until all the chambers are ignited. Actually only
a few seconds are needed for this process. Then
the two spark igniters are no longer needed, and
they cut off automatically.
To be sure the can-annular type combustion
chambers have positive ignition during the starting
cycle, two spark igniters are used and located in
the two lower chambers.
Another very important requirement in the
construction of combustion chambers is providing
the means for draining unburned fuel. The
drainage requirement involves many factors, such
as the prevention of residual fuel deposits (gum)
after evaporation in the fuel manifold, nozzles,
and combustion chambers. Also, if fuel is allowed
to accumulate after shutdown, an afterfire could
occur. Another possibility is at the next starting
attempt, the excess fuel in the combustion
chamber could ignite. Tailpipe temperature could
go beyond safe operating limits.
The liners of the can-type combustors
(fig. 1-25) have the usual perforations of various
sizes and shapes. Each hole has a specific purpose
and effect on the flame propagation within the
liner. The air entering the combustion chamber
1. The multiple chamber, or can
2. The annular, or basket
3. The can-annular
Can Type
The can-type combustion chamber is typical
of the type used on axial-flow engines. Can-type
combustion chambers are arranged radially
around the axis of the engine. The amount of
chambers will vary in number. In the past (or
development years) as few as 2 and as many as
16 chambers have been used. The present trend
shows the use of about 8 or 10 combustion
chambers. Figure 1-25 shows the liner of a cantype combustion chamber. These chambers are
numbered in a clockwise direction. As you face
the rear of the engine and look forward, the
number 1 chamber is at the top.
Some provision is made in the combustion
chamber case or in the compressor air outlet elbow
for the installation of a fuel nozzle. The fuel
nozzle delivers the fuel into the liner in a finely
atomized spray. The finer the spray, the more
rapid and efficient the burning process.
The two types of fuel nozzles being used in
the various types of combustion chambers are the
simplex nozzle and the duplex nozzle. The fuel
nozzles are constructed so they can be installed
in various ways. The two methods used most
frequently are external mounting and internal
mounting. In external mounting, a mounting pad
is provided for attachment of the nozzle to the
case or the inlet air elbow, with the nozzle tip
projecting into the chamber liner, usually near the
Figure 1-25.-Cars-type combustion liner.
details of the liner. The liner consists of an
undivided circular shroud extending all the way
around the outside of the turbine shaft housing
(fig. 1-27). The chamber is constructed of one or
more baskets. If two or more chambers are used,
they are placed one outside of the other in the
same radial plane. The double-annular chamber
is shown in figure 1-27.
The combustion chamber housing is made in
three sections. These sections are the inlet, center,
and rear sections.
The inlet section receives the air from the axialflow compressor. This section is a diffuser. It
slows the velocity of the air by providing a larger
area just before the liner area, thus raising air
pressure. Also present is a coarse wire screen,
whose function is to increase turbulence to aid in
fuel atomization.
The center section of the chamber housing
surrounds the liner, providing an outer wall for
the axial path of the air. The center section
provides the mounting pads for the installation
of fuel drain valves. The drain valves drain
residual or accumulated fuel out of the combustion chamber after engine shutdown. This action
prevents afterfires or excessive starting
temperatures during the next start. Located on the
bottom of the housing are the spring-loaded
combustion chamber drain valves. These valves
drain automatically whenever internal chamber
pressures approach atmospheric pressure. This
fuel is drained to an overboard drain compartment in the airframe.
The rear section converges to form a narrow
annulus. This type of construction speeds up
airflow before it enters the turbine section.
Fuel is introduced through a series of nozzles
at the upstream end of the liner. The fuel nozzles
are screwed into fuel manifolds, located within
two concentric fairings. If the chamber liner is of
double-annular construction, there are two fuel
manifolds. Only one manifold would be required
if it were of single-annular construction. The two
concentric fairings that support the fuel manifolds
also perform the function of dividing the entering
airflow into three concentric annular streams. The
outer stream is delivered to the space between the
combustion chamber liner and the chamber
housing. The middle stream is delivered to the
space between the inner and outer sections of the
liner. The inner stream is delivered to the space
between the liner and the rotor shaft housing. The
two concentric fairings are supported by radial
struts in the diffuser section.
is divided by the proper holes. Louvers and slots
divide the main streams—primary and secondary
air. The primary or combustion air is directed
inside the liner at the front end, where it mixes
with the fuel and is burned. Secondary or cooling
air passes between the outer casing and the liner
and joins the combustion gases through larger
holes toward the rear of the liner. Combustion
gases are cooled from about 3,500°F to about
1,500°F forward of the turbine. Holes are
provided to aid in atomization of the fuel. These
holes are located around the fuel nozzle in the
dome or inlet end of the liner. Louvers are also
provided along the axial length of the liners to
direct a cooling layer of air along the inside wall
of the liner. This layer of air controls the flame
pattern by keeping it centered in the liner. This
air layer prevents the 3,000°F temperatures of the
combusting gases from burning the liner walls.
Figure 1-26 shows the flow of air through the
louvers in the double-annular type of combustion
Annular or Basket Type
The annular combustion chamber, the type
usually found in axial-flow engines, consists
basically of a housing and a liner, similar to the
can type. The difference lies in the construction
Figure 1-26.-Components and airflow of a double-annular
combustion chamber.
Figure 1-27.-Double-annular combustion chambers.
The spark igniter plugs of the annular
combustion chamber are the same basic type used in
the can combustion chambers. 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.
Figure 1-26 shows a cross section of the parts of
the double-annular combustion section quite clearly.
Also included in the figure is the flow of air from the
axial-flow compressor discharge through the liner on
the turbine section. Primary air and secondary air
are indicated by arrows.
The rotor shaft housing, shown in figure 1-28,
provides the inner wall for the axial path of the air.
This housing unit is attached to the rear face of the
inner ring of the diffuser section. The housing is
supported in the rear by components of the turbine
section, not shown in figure 1-28.
Can-Annular Type
The can-annular type combustion chamber is a
development by Pratt and Whitney for use in
the axis in this instance being the rotor shaft
housing. Figure 1-28 shows this arrangement to
The combustion chambers are enclosed by a
removable steel shroud, which covers the entire
burner section. This feature makes the burners
readily available for any required maintenance.
The burners are interconnected by projecting
flame tubes, which help the engine-starting
process in the can-type combustion chamber.
These flame tubes perform a function identical
with those previously discussed, the only
difference being in construction details.
Figure 1-28 also reveals that each of the
combustion chambers contains a central bulletshaped perforated liner. The size and shape of the
holes are predetermined to admit the correct
quantity of air at the velocity and angle required
to control the flame pattern. Cutouts are provided
in two of the bottom chambers for installation
of the spark igniters. Notice in figure 1-28 how the
combustion chambers are supported at the aft end
by matching outlet ducts in the turbine nozzle
Figure 1-28 shows how the forward face of the
chambers presents size apertures that align with
the six fuel nozzles of the corresponding fuel
their J57 axial-flow turbojet engine. Since this
engine features the split-spool compressor, it
requires combustion chambers capable of meeting
the stringent requirements of maximum strength,
limited length, and high overall efficiency. These
attributes are necessary because of the high air
pressures and velocities present in a split-spool
compressor, along with the shaft length
The split-spool compressor requires two
concentric shafts joining the turbine stages to their
respective compressors. The front compressor,
joined to the rear turbine stages, requires the
longest shaft. This shaft is inside the other. A
limitation of diameter is imposed, so that the
distance between the front compressor and the
rear turbine must be limited if critical shaft lengths
are to be avoided. High torque is present if there
is a long shaft of small diameter.
Since the compressor and turbine are not
susceptible to shortening, shaft length limitation
is absorbed by developing a new type of burner.
The burner designers had to develop a design that
would give the desired performance in much less
relative linear distance.
The can-annular combustion chambers are
arranged radially around the axis of the engine;
Figure 1-28.-Can-annular combustion chamber components and arrangement.
nozzle cluster. These nozzles are the dual-orifice
(duplex) type that require the use of a flow-divider
(pressurizing valve), as mentioned above in the
can-type combustion chamber. Preswirl vanes are
located around each of the nozzles for imparting
a swirling motion to the fuel spray. This results
in better atomization of the fuel, thus better
burning and efficiency.
The swirl vanes perform two important
functions imperative to proper flame propagation:
mechanical energy to drive the compressor and
necessary accessories. This is the sole purpose of
the turbine. This function absorbs about 60 to 80
percent of the total pressure energy from the
exhaust gases. The exact amount of energy
absorption at the turbine is determined by the load
the turbine is driving. The compressor size, type,
accessories, and a propeller and its reduction gears
if the engine is a turbo-propeller type also effect
The turbine section of a turbojet engine is
located aft, or downstream, of the combustion
chamber section. Specifically, it is directly behind
the combustion chamber outlet.
The turbine assembly consists of two basic
elements, the stator and the rotor, as does the
compressor unit. These two elements are shown
in figures 1-30 and 1-31.
The stator element is known by a variety of
names. Turbine nozzle vanes, turbine guide vanes,
1. High flame speed: Better mixing of air and
fuel, ensuring spontaneous burning.
2. Swirling prevents the flame from moving
rapidly rearward.
The swirl vanes greatly aid flame propagation,
since a high degree of turbulence in the early
combustion and cooling stages is desirable. The
vigorous mechanical mixing of the fuel vapor with
the primary air is necessary, since mixing by
diffusion alone is too slow. This same mechanical
mixing is also established by other means, such
as placing coarse screens in the diffuser outlet,
as is the case in most axial-flow engines.
The can-annular combustion chambers also
must have the required fuel drain valves, located
in two or more of the bottom chambers, thereby
assuring proper drainage and eliminating the
possibility of residual fuel burning during the next
start cycle.
The flow of air through the holes and louvers
of the can-annular chambers is the same as the
flow through other types of burners. Figure 1-29
shows the flow of combustion air, metal-cooling
air, and the diluent or gas-cooling air. Pay
particular attention to the direction of airflow,
indicated by the arrows.
The turbine transforms a portion of the kinetic
(velocity) energy of the exhaust gases into
Figure 1-30.-Stator element of the turbine assembly.
Figure 1-29.-Airflow through a can-annular chamber.
Figure 1-31.-Rotor element of the turbine assembly.
and nozzle diaphragm are three of the most
commonly used. The turbine nozzle vanes are
located directly aft of the combustion chambers
and immediately forward of the turbine wheel.
The function of the turbine nozzle is twofold.
First, after the combustion chamber has introduced the heat energy into the mass airflow and
delivered it evenly to the turbine nozzle, it
becomes the job of the nozzle to prepare the mass
flow for harnessing of power through the turbine
rotor. The stationary vanes of the turbine nozzle
are contoured and set at such an angle that they
form small nozzles. They discharge the gas as
extremely high-speed jets. Thus, the nozzle
converts a varying portion of the heat and
pressure energy to velocity energy. It can then be
converted to mechanical energy through the rotor
The second purpose of the turbine nozzle is
to deflect the gases to a specific angle in the direction of turbine wheel rotation. Since the gas flow
from the nozzle must enter the turbine blade
passageway while the turbine is still rotating, it
is essential to aim the gas in the general direction
of turbine rotation.
The elements of the turbine nozzle assembly
consist of an inner shroud and an outer shroud,
between the nozzle vanes. The number of vanes
employed varies with different types and sizes of
engines. Views A and B of figure 1-32 show typical
turbine nozzles featuring loose and welded vane
fits, respectively.
Figure 1-32.-Turbine nozzle vane assembly. (A) With loose
fitting vanes; (B) with welded vanes.
vanes could fall out of the shrouds as the shrouds
are removed.
Another method of thermal expansion
construction is to fit the vanes into inner and outer
shrouds. However, this method welds or rivets the
vanes into position (fig. 1-32, view B). Some
means provide for the inevitable thermal
expansion; therefore, either the inner or the outer
shroud ring cuts into segments. These saw
cuts dividing the segments will allow enough
expansion to prevent stress and warping of the
The vanes of the turbine nozzle are assembled
between the outer and inner shrouds or rings in
a variety of ways. Although the actual elements
may vary slightly in their configuration and
construction features, there is one characteristic
peculiar to all turbine nozzles; that is, the nozzle
vanes are constructed to allow for thermal expansion. Otherwise there would be severe distortion
or warping of the metal parts because of rapid
temperature variances.
The rotor element of the turbine section
consists essentially of a shaft and a wheel
(fig. 1-31).
The expansion feature of the turbine nozzle
is accomplished by one of several methods. One
method has the vanes assembled loosely in the
supporting inner and outer shrouds (fig. 1-32,
view A). Each of the vanes fits into a contoured
slot in the shrouds. They conform with the airfoil shape of the vanes. These slots are slightly
larger than the vanes to give a loose fit. The inner
and outer shrouds are encased by an inner and
an outer support ring, which give increased
strength and rigidity. These supports also help
remove the nozzle vanes as a unit. Otherwise, the
The following brief discussion of impulse and
reaction turbines should help clarify their
function. The turbine blades are of two basic
types—impulse and reaction. Most aircraft
engines use a blade with both impulse and reaction
sections. The impulse is usually at the base of the
The impulse turbine can be defined as a
turbine that derives its rotation from the weight
and velocity of the air striking its blades. The
reaction turbine derives its rotation from the air
pressure across its blades, as in an airfoil. See
figure 1-33.
The impulse-reaction turbine combines the
rotational forces described in the two previous
turbines. It derives its rotation from the weight
of air striking the turbine blades and the airfoil
reaction of air passing over the blade’s surface.
The turbine wheel is a dynamically balanced
unit with blades attached to a rotating disc. The
disc is attached to the main power-transmitting
shaft of the engine. The jet gases leaving the
turbine nozzle vanes act on the blades of the
turbine wheel, causing the assembly to rotate at
a very high speed. The high rotational speed
causes heavy centrifugal loads on the turbine
wheel. The elevated temperatures result in a lowering of the strength of the material. The engine
speed and temperature must be controlled to keep
turbine operation within safe limits.
The turbine wheel, without blades, is known
as a turbine disc. The disc acts as an anchoring
part for the turbine blades. Since the disc is
attached to the rotor shaft, the exhaust gas energy
extracted by the blades is imparted to the shaft.
The disc rim exposes the hot gases passing
through the blades and absorbs considerable heat
from these gases. In addition, the rim also absorbs
heat from the turbine buckets by conduction.
Hence, disc rim temperature slopes are quite high
and well above the temperatures of the more
remote inner portion of the disc. As a result of
these temperature slopes, thermal stresses are
added to the stresses due to rotation.
There are various methods provided to relieve,
at least partially, these stresses. One such method
is the incorporation of an auxiliary fan somewhere
ahead of the disc. Usually rotor-shaft driven, it
forces cooling air back into the face of the disc.
Another method of relieving the thermal
stresses of the disc follows as incidental to blade
installation. The disc rims are notched to conform
with the blade root design. The disc is made
adaptable for retaining the turbine blades. At the
same time, space is provided by the notches for
thermal expansion of the disc.
The turbine shaft, shown in figure 1-31, is
made from low-alloy steel. It must be capable of
absorbing high torque loads, such as when a heavy
axial-flow compressor is started.
The methods of connecting the shaft to the
turbine disc vary. One method used is welding.
The shaft is welded to the disc, which has a butt
or protrusion provided for the joint. Another
method is by bolting. This method requires that
the shaft have a hub that matches a machined
surface on the disc face. The bolts are then
inserted through holes in the shaft hub and
anchored in tapped holes in the disc. Of the two
methods, bolting is more common.
To join the turbine shaft to the compressor
rotor hub, make a splined cut on the forward end
of the shaft. The spline fits into a coupling device
between the compressor and turbine shafts. If a
coupling is not used, the splined end of the
turbine shaft may fit into a splined recess in the
compressor rotor hub. The axial compressor
engine may use either of these methods.
There are various ways of attaching turbine
blades or buckets, some similar to compressor
blade attachment. The most satisfactory method
used is the fir-tree design, shown in figure 1-34.
The blades are retained in their respective
grooves by a variety of methods; some of the more
common ones are peening, welding, locking tabs,
Figure 1-34.-Turbine blade with fir-tree design root and
tab lock method retention.
Figure 1-33 .-Impulse and reaction blades.
elements necessary for its retention. This method
is shown in figure 1-36. The blade root has a stop
on one end of the root. The blade is inserted and
moves in one direction only, while on the opposite
end of the blade is a tang. This tang is peened over
to secure the blade in the disc.
Turbine blades may be either forged or cast,
depending on the composition of the alloys. Most
blades are precision cast and finish-ground to the
desired shape.
Most turbines in use are open at the outer
perimeter of the blades; there is a second type
called the shrouded turbine. The shrouded
turbine blades, in effect, form a band around the
outer perimeter of the turbine wheel. This
improves efficiency and vibration characteristics
and permits lighter stage weights; on the other
hand, it limits turbine speed and requires more
blades (fig. 1-37).
In turbine rotor construction, it may be
necessary to use turbines of more than one stage.
A single turbine wheel often cannot absorb
enough power from the exhaust gases to drive the
parts dependent on the turbine for rotative power.
In a turbojet engine, these parts are the compressor and engine-driven accessories. In the
turboprop engine, these parts are the propeller and
its reduction gearing.
A turbine stage consists of a row of stationary
vanes or nozzles, followed by a row of rotating
and riveting. Figure 1-35 shows a typical turbine
wheel using riveting for blade retention.
A method of blade retention used quite
frequently is peening, and it applies in various
ways. Two of the most common applications of
peening are described in the following paragraphs.
One method of peening requires that a small
notch be ground in the edge of the blade fir-tree
root before blade installation. The blade inserts
into the disc. The notch is filled with the disc
metal, which is “flowed” into it through a small
punch mark made in the disc, adjacent to the
notch. The tool used for this job is similar to a
center punch, and is usually manufactured locally.
Another method of peening is to construct the
blades root in such a way as to contain all the
Figure 1-35.-Riveting method of turbine blade retention.
Figure 1-37.-Shrouded turbine blades.
Figure 1-36.-Turbine bucket, featuring peening method of
blade retention.
Figure 1-38.-Multi-rotor turbine.
blades. Some models of turboprop engines use as
many as five turbine stages. You should remember
that regardless of the number of wheels necessary
for driving engine parts, there is always a turbine
nozzle in front of each wheel.
The occasional use of more than one turbine
wheel is necessary in cases of heavy rotational
loads. Heavy loads that require multiple-stage
turbine wheels often make it advantageous to
use multiple rotors. Shafts are bolted to the
appropriate turbine on one end and at the other
end to the unit requiring the rotative power.
Typical of this situation are split compressors or
propellers, or a gas generator for helicopters, In
each of these situations, the turbine for each of
the rotors may have one or more stages.
In the single-rotor turbine, shown in figure
1-23, the power is developed by one rotor. All
engine-driven parts are driven by this single wheel,
This arrangement uses engines where the need for
low weight and compactness predominates. The
single-rotor turbine may be either single or
multiple stage.
In the multiple-rotor turbine, the power is
developed by two or more rotors. It is possible
for each turbine rotor to drive a separate part of
the engine. For example, a triple-rotor turbine
may be so arranged that the first turbine
drives the rear half of the compressor and the
accessories. The second turbine drives the front
half of the compressor, and the third turbine
furnishes power to a propeller (fig. 1-38).
The turbine rotor arrangement for a dualrotor turbine, such as required for a split-spool
compressor, is similar to the arrangement in figure
1-38. The difference is in the use of the third
turbine for a propeller,
The remaining elements concerning the turbine
is the turbine casing or housing. The turbine
casing encloses the turbine wheel and the nozzle
vane assembly. It gives either direct or indirect
support to the stator elements of the turbine
The exhaust section of the turbojet engine is
made up of several parts, each of which has its
individual functions. Although the parts have
individual purposes, they also have one common
function. They must direct the flow of hot gases
rearward in such a manner as to prevent
turbulence, while causing a high final or exit
velocity to the gases.
In performing the various functions, each of
the parts affects the flow of gases in different
ways, as described in the following paragraphs.
The exhaust section is directly behind the
turbine section. It ends with the ejection of gas
at the rear in the form of a high-velocity jet.
The parts of the exhaust section include the
exhaust cone, tailpipe (if required), and the
exhaust, or jet nozzle, Each of these parts is
discussed individually so the exhaust section will
be quite familiar to you.
The exhaust cone collects the exhaust gases
discharged from the turbine assembly and
gradually converts them into a solid jet. During
this operation, the velocity of the gases will
decrease slightly, and the pressure will increase.
This is caused by the diverging passage between
the outer duct and the inner cone. The annular
area between the two units increases rearward
(fig. 1-40).
Figure 1-39.-Turbine casing assembly.
Figure 1-40.-Exhaust collector and welded support status.
section. It always has flanges to provide for the
front and rear bolting of the assembly to the
combustion chamber housing and the exhaust
cone assembly, respectively. Figure 1-39 shows a
turbine casing.
The elements of the exhaust cone assembly
consist of an outer shell or duct, an inner cone,
and three or four radial hollow struts or fins. Tie
rods aid the struts in supporting the inner cone
from the outer duct.
The outer shell or duct, made of stainless steel,
is attached to the rear flange of the turbine case.
This element collects and delivers the exhaust
gases. The gases flow either directly or through
a tailpipe to the jet nozzle, depending on whether
or not a tailpipe is required. There is no need for
a tailpipe in some engines. For instance, the
required engine-installation space in wing roots,
pods, or wings is short and requires very little
tailpipe. In which case the exhaust duct and exhaust nozzle will suffice. The construction of the
duct includes such features as a predetermined
number of thermocouple bosses for installing
exhaust gas temperature thermocouples. Also,
there must be insertion holes for the supporting
tie rods. In some cases, there are no requirements
for tie rods for supporting the inner cone from
the outer duct. If such is the case, the hollow struts
provide the sole support of the inner cone; the
struts being spot-welded in position to the inside
surface of the duct and to the inner cone,
The radial struts actually have a twofold
function. They not only support the inner cone
in the exhaust duct, they also perform the
important function of straightening the swirling
exhaust gases, which otherwise would leave the
turbine at an angle of about 45 degrees. If tie rods
are required for inner cone support, these struts
also form fairings around the rods.
The centrally located inner cone fits rather
closely against the rear face of the turbine disc.
This fit prevents turbulence of the gases as they
leave the turbine wheel. The cone is supported by
the radial struts, which are usually vertical and
horizontal in relation to the normal position of
the engine. In some configurations, there is a small
hole located in the exit tip of the cone (fig. 1-40).
This hole allows cooling air to circulate from the
exit end of the cone. The pressure of the gases
is relatively high in the interior of the cone and
against the face of the turbine wheel. The flow
of air is positive, since the air pressure at the
turbine wheel is relatively low due to rotation of
the wheel, thus air circulation is assured. The gases
for cooling the turbine wheel return to the path
of flow by passing through the clearance between
the disc and the cone. The clearance between the
turbine disc and the inner cone must be checked
periodically since the higher pressures aft tend to
push the inner cone against the turbine wheel.
The exhaust cone assembly is the terminating
part of the basic engine. The remaining parts (the
tailpipe and jet nozzle) are usually considered airframe parts.
The tailpipe pipes the exhaust gases out of the
airframe. Actually, the tailpipe imposes a penalty
on the operating efficiency of the engine in the
form of heat and duct (friction) losses. These
losses materially affect the final velocity of the
exhaust gases and, hence, the thrust.
Figure 1-41.-Exhaust system insulation blanket.
the gases pass the turbine, the reduction of heat
losses improves engine performance by retaining
the maximum permissible temperatures, resulting
in maximum velocity in the jet. A typical insulation blanket and the temperatures at the various
locations in the exhaust section are shown in
figure 1-42. This blanket contains fiber glass as
the low-conductance material and aluminum foil
as the radiation shield. The blanket is covered to
prevent its becoming soaked with oil. The heat
shroud type of configuration consists of a stainless
steel envelope enclosing the exhaust system
(fig. 1-43).
The tailpipe ends in a jet nozzle, located
just forward of the end of the fuselage. Most
installations employ a single direct exhaust to get
the advantages of low weight, simplicity, and
minimum duct losses.
The construction of the tailpipe is semiflexible.
Again, the need for this feature is dependent on
its length. On an extremely long tailpipe, a bellows
arrangement allows movement both in installation
and maintenance and in thermal expansion. This
cuts stress and warping, which would otherwise
be present.
The heat radiation from the exhaust cone and
tailpipe could conceivably injure the airframe
parts surrounding these units. For this reason
some means of insulation had to be devised. There
are several suitable methods for protection of the
fuselage structure; two of the most common are
insulation blankets and shrouds.
An insulation blanket type of configuration,
shown in figures 1-41 and 1-42, consists of several
layers of aluminum foil, each separated by a layer
of bronze screening or some other suitable
material. Although these blankets protect the
fuselage from heat radiation, they primarily
reduce heat losses from the exhaust system. Since
engine temperature limits are of little concern after
Figure 1-43.-Exhaust system shroud.
Figure 1-42.-Insulation blanket, with the temperatures that would be obtained at the various locations shown.
NOTE: The exhaust or jet nozzle
gives to the exhaust gases the allimportant final boost in velocity.
The jet nozzle, and the tailpipe, is
not a part of the basic power
plant, but is supplied as a part of
airframe. The nozzle attaches to
the rear of the tailpipe if there is a
need. It is attached to the rear
flange of the exhaust duct if a
tailpipe is not necessary. There a re
basically two types of jet nozzles—
fixed-area and variable-area.
There are two types of jet nozzle design. They
are the converging design, used on most freed-area
nozzles for subsonic velocities, and the convergingdiverging design, for supersonic gas velocities.
The fixed-area type is the simpler of the two jet
nozzles, since there are no moving parts. It is
attached to either a tailpipe or exhaust cone, and any
adjustment in nozzle area is mechanical.
Adjustments in a fixed-area nozzle are
sometimes necessary because the size of the exit
orifice will directly affect the operating temperature
of the engine. There are several ways to adjust a
fixed-area nozzle. One method is to trim or cut away
strips from the conical section of the exhaust nozzle.
Provided, of course, the temperature was too high. If
the inlet temperature is too low, a nozzle of less area
is used to replace the inadequate one.
Another method of reducing the nozzle area is to
use inserts. The inserts fit inside a joggled retainer
held in place by two screws.
Figure 1-44.-Segment-type nozzle assembly.
Figure 1-45.-Variable exhaust nozzle assembly.
hinged to the rear of the tailpipe. The secondary
nozzle is secured by a stationary supporting shroud,
on which the pivot points for the flap-operated
mechanisms are located. The flaps are slotted to
permit thermal expansion and are mounted onto the
tailpipe. The flaps are controlled by four synchronized
Usually a set of 10 inserts of various curvatures
is provided with each aircraft. The different size
inserts allow a total change of 10 square inches in
nozzle area in 1-inch increments. Thus, through
experience, a mechanic can run the engine at
maximum speed with one combination of inserts.
Check the temperature, and substitute another
combination to make up a temperature deficiency or
remedy an excess temperature situation.
NOTE: All advanced-technology engines
now used by the Navy have state-ofthe-art electronic parts that eliminate
the need for physically changing the
exhaust nozzle area.
The accessory section of the turbojet engine has
various functions. The primary function is to provide
space for the mounting of accessories necessary for
the operation and control of the engine. It also
includes accessories concerned with the aircraft, such
as electric generators and fluid power pumps. The
secondary purpose includes acting as an oil reservoir,
oil sump, and providing for and housing of accessory
drive gears and reduction gears.
The arrangement and driving of accessories have
always been major problems on gas turbine engines.
Driven accessories are mounted on common pads
either ahead of or next to the compressor section.
Figure 1-46 shows the accessory arrangement of
an axial-flow engine.
The variable-area nozzle at the exhaust exit is
automatic. A very important use of this type of nozzle
is to increase the exit area during afterburning.
The segment-type nozzle area opens and closes
by individual overlapping sliding segments. See
figure 1-44.
Some engines use the inner- and outer-flap
variable nozzle assembly, shown in figure 1-45. The
assembly has an internal primary nozzle with
sectional flaps and an external secondary nozzle with
sectional flaps. The flaps of the primary nozzle are
The parts of an axial-flow engine accessory
section are the accessory gearbox and a power
takeoff assembly. These units contain the
necessary drive shafts and reduction gears. Views
A and B of figure 1-47 show the location of the
accessory gearbox.
The accessory gearbox and the power takeoff
are located near each other. There are two
factors that affect the location of gearboxes in
general. These factors are engine diameter and
engine installation.
Designers strive to reduce engine diameter to
make the engine more streamlined, thereby
increasing performance by reducing drag. Also,
engine installation in a particular aircraft may
dictate the location or rearrangement of the
accessory gearboxes.
The accessories on engines are the fuel control
with its governing device, the high-pressure fuel
pump(s), and a breather screen or other means
for venting the oil system. Other parts are oil
sump, oil pressure and scavenge pumps, auxiliary
fuel pump, starting fuel pump, and other
accessories, including starter, generator, and
tachometer. Although these accessories are
essential, the particular combination of enginedriven accessories depends upon the use for which
the engine is designed.
The accessories mentioned above (except
starters) are of the engine-driven type. There are
the nondriven-type accessories such as booster
coils or ignition exciters, fuel and oil filters,
barometric units, drip valves, compressor bleed
valves, and relief valves.
The afterburner increases or boosts the normal
thrust rating of a gas turbine engine. There are
times when the maximum normal thrust of an
engine is not enough. For instance, it is
conceivable that although the in-flight requirements are met satisfactorily by an engine of
moderate size, the aircraft still may not have
good takeoff performance. With afterburning,
maximum thrust is obtained without sacrificing
the economy of the small basic gas turbine.
Increased thrust is required for takeoff, emergencies, and combat conditions.
The afterburner duct replaces the usual aircraft tailpipe. Actually, it is more like a converted
tailpipe. It functions as the engine tailpipe during
nonafterburning (cold) operation and is also the
main working element of the afterburner. The
entire afterburner is projected from the engine.
It is supported only at the exhaust end where it
is bolted to the engine.
The essential working element of the afterburner is an afterburner duct. A flameholder or
diffuser and a variable-area exhaust nozzle are the
other parts (fig. 1-45).
The afterburner duct is the main working
element of the afterburner. It’s designed so
that the normal pressure relationship between
the air entering the main engine turbine and
the air leaving the turbine is not upset. Since
the duct acts as a burner, the inlet air velocity
must be sufficiently low to support stable
combustion and to avoid excessive pressure
losses. For these purposes a diffuser is located
between the turbine outlet and the tailpipe
burner inlet. Thus, the burner section of the
duct can reduce gas velocities so they do
not exceed the flame propagation rate. Otherwise, the flame could not get a foothold,
because the onrushing turbine exhaust would
simply push the burning mixture right out
the exhaust nozzle. In addition to the diffuser,
some mechanical mixing of the fuel and air
is necessary. Mixing by diffusion is too slow
a process to be an aid in forming a combustible
The flameholders provide local turbulence and
reduce velocity, which aids combustion stability.
The flameholders are located downstream from
the fuel-injection nozzles, thereby allowing time
Figure 1-47.-Accessory gearbox. (A) Mounted beneath the
compressor; (B) mounted beneath the front bearing support.
engine is subjected to additional burning after the
basic cycle is completed. The unburned oxygen
in the air used for cooling the exhaust gases is the
afterburner air supply. Since no cooling is
required past the turbine assembly, all or any part
of this air may be burned for augmentation.
Additional energy is given to the exhaust gases
by burning additional fuel sprayed in the exhaust
stream aft of the turbine.
for proper mixing of the fuel and air before
reaching the burner area. The flameholders are
circular, concentrically mounted, and supported
in position by tie rods that project through the
wall of the duct. Two of the tie rods are in a
horizontal position, and the remaining two are
The afterburner, tailpipe burner, or reheater
get their names because the air going through the
After completing this chapter, you will be able to:
Identify the purpose and procedures of the
tool control program.
Recognize the use and safety procedures for
special tools (torque wrenches and micrometers).
Recognize the use and safety procedures for
common hand tools.
Recognize the selection, identification, and
proper use of different aircraft hardware.
Each year, thousands of dollars of damage to
aircraft engines and serious injuries to personnel
result from the wrong use of tools and hardware.
Using the wrong type of hardware or improper
safetying of hardware may cause flight controls
to jam or come loose. Engines experience foreign
object damage (FOD) because of the improper use
of tools and hardware. Tools are found in aircraft fuel cells and engine bays even though a Tool
Control Program guards against these mistakes.
With so much at stake, we must continually
emphasize basic tool and hardware use.
can’t drive spikes with a tack hammer; yet, some
may try. It is your responsibility to keep your tools
in good condition and ready for use. A screwdriver that has been damaged by use as a chisel
or pry bar has no place in a mechanic’s toolbox.
Damaged tools may cause damage to parts or
injury to the worker.
The material in this chapter emphasizes
information from other training manuals and
should be studied with them. For a complete
description of different tools and their use,
refer to Tools and Their Uses, N A V E D TRA 10085 (series).
You must have a well-rounded knowledge of
many different types of tools. You must know the
purpose for which the tools were designed. An
important attribute that a mechanic must have is
the ability to use the right tool for the job. The
correct tool must be used whether that job is a
minor adjustment or a major engine overhaul.
The Tool Control Program was established to
reduce FOD-related mishaps. The program is
intended to ensure tool accountability both before
and following the performance of aircraft-related
maintenance tasks.
The CNO is the overall sponsor. NAVAIRSYSCOM is the responsible agent for development and issuance of a Tool Control Plan for
each type of aircraft and engine. This TCP is a
guide to aviation maintenance activities in the
You have often heard that a mechanic is only
as good as his tools. That is a half-truth. The
mechanic must not only have the correct tools but
must know the proper use of these tools. You
implementation of their own TCP. The following information is part of each TCP:
inventory as well as a separate listing of tools in
calibration or requiring replacement.
1. An allowance list for tool containers
The AD who has custody of a toolbox must
prevent the loss of the tools or the toolbox
through neglect or misuse. Although hand tools
are normally classified as consumable items, they
are expensive and must be paid for when lost or
damaged. OPNAVINST 4790.2 (series) outlines
the policies and procedures for control of hand
tools. Usually, your activity will have a local MI
concerning the inventory interval and methods for
reporting lost or damaged tools.
2. A standard tool list and layout diagram for
each container
3. Procurement information necessary to
procure tool containers and other associated
Aircraft Controlling Custodians (ACCs) are
required to implement each TCP after it has been
formulated and released. Each ACC sets forth a
specific policy by means of instructions. Examples
of these publications are the CNALINST 4790.16
(series), CNAPINST 4790.18 (series), and NAVAIRINST 10290.2 (series) instructions. The Naval
Air Engineering Center will process revisions to
the tool allowance list, as well as error lists, Each
local command, ship, and squadron should
prepare a local command maintenance instruction
(MI) to assign the responsibilities for the TCP.
The material control officer is responsible for
ensuring that tools are procured and issued on a
controlled basis. Some commands may establish
a tool control center under the material control
officer. In activities operating aboard ships,
where a tool control center is not practical, the
commanding officer designates, in writing, a tool
control coordinator.
NOTE: Broken or damaged tools can
damage equipment, hardware, and parts.
They can also cause personal injury to the
worker or others.
All personnel have specific responsibilities
under the TCP. All tool containers should have
a lock and key as part of their inventory. When
work is to be completed away from the work
spaces, complete tool containers should be taken
to the job. If you need more tools than the tool
container contains, tool tags may be used to check
out additional tools. These tools come from other
tool containers in the work center or from another
work center. The following is a list of some of
your responsibilities under the work center Tool
Control Program:
1. Upon task assignment, note the number of
the tool container on copy 1 of the VIDS/MAF.
Place this note to the left of the Accumulated
Work Hours section. Conduct a sight inventory
before beginning each task. All shortages must
be noted. Every measure must be taken to make
sure missing tools do not become a cause of FOD:
Perform inventories before a shift change, when
work stoppage occurs, and after maintenance has
been completed. Perform an inventory before
conducting an operational systems check on the
The TCP contains the listing of each tool
container. The TCP acts as an inventory for each
type, model, and series of aircraft and equipment
worked on. The container exterior will clearly
identify the work center, tool container number,
and organization. The tools in each container will
have the work center code, organization code, and
container number etched onto them. Special
accountability procedures will be established
locally for those tools not suitable for etching.
Drill bits (too hard), jeweler’s screwdrivers
(too small), and beryllium hand tools (dust is
hazardous to personnel) are not suitable for
2. After you account for all tools and
complete all maintenance actions, the work center
supervisor signs the VIDS/MAF.
Tool pouches are to be considered as tool
containers, and most are manufactured locally.
The position for each tool in the container will
be silhouetted against a contrasting background.
The silhouetted tool outline will highlight
each tool location within the container. Those
containers not silhouetted will contain a diagram
of the tool locations. Containers will include an
3. If any tool is found to be missing during
the required inventories, conduct an immediate
search. The search should occur before reporting
the work completed or signing off the VIDS/
MAF. If the tool cannot be located, notify
maintenance control to ensure that the aircraft or
equipment is not released.
If the tool cannot be located, the person doing
the investigation will sign a lost tool statement and
the VIDS/MAF. The statement indicates that a
lost tool investigation was conducted and that the
tool was not found. After the investigation, follow
the normal VIDS/MAF completion process.
In this chapter the term common hand tools
is used to refer to small, nonpowered hand tools
that are common to the AD rating. This term
includes such common tools as hammers, socket
sets, wrenches, screwdrivers, and pliers.
Figure 2-1.-Hammers.
BALL PEEN HAMMER.— The ball peen
hammer is sometimes referred to as a machinist’s
hammer. It is a hard-faced hammer made of
forged tool steel.
Hammers are dangerous tools when used
carelessly and without consideration, Practice will
help the inexperienced to learn how to use a
hammer properly. Hold the handle near the end
with your fingers underneath and your thumb
along the side or on top of the handle. Your
thumb should rest on the handle and never overlap
your fingers. Oils on the face of the hammer will
cause it to glance off the work. Wipe the oil off
with a rag and rub the face with coarse sandpaper
or emery cloth. Never use a hammer that has a
loose head or cracked handle. Most hammer
accidents are caused by a loose head or a slippery
handle. So remember these tips when using any
kind of striking tool. Tighten the loose hammerhead by driving a wedge in the end of the handle.
The wedge spreads the handle tightly inside the
head. Do not strike a hardened steel surface with
a steel hammer. Small pieces of steel may break
and injure someone or damage the work. Use a
soft hammer in striking hardened steel or highly.
polished stock. If a soft hammer is not available,
use a piece of copper, brass, lead, or wood to
protect the hardened steel. It is permissible to
strike a punch or chisel directly with the ball-peen
hammer because the steel in the heads of punches
and chisels is slightly softer than that of the
There are various types of hammers, all of
which are used to apply a striking force where the
force of the hand alone is insufficient. Each of
these hammers has a head and a handle, even
though these parts differ greatly from hammer to
hammer. So that you may have a better idea of
their differences and uses, let’s consider the types
of hammers used most frequently. See figure 2-1.
The flat end of the head is called the face. This
end is used for most hammering jobs. The other
end of the hammer is called the peen. The peen
end is smaller in diameter than the face and is
useful for striking areas that are too small for the
face to enter.
Ball peen hammers are made in different
weights, usually 4, 6, 8, and 12 ounces and
1, 1 1/2, and 2 pounds. For most work, a
1 l/2-pound or a 12-ounce hammer will do.
MALLET.— A mallet is a soft-faced hammer.
Mallets are made with brass, lead, tightly rolled
strips of rawhide, and plastic heads, Sometimes
the plastic head has a lead core for added weight.
Plastic mallets similar to the one shown in
figure 2-1 are the type normally found in your
toolbox. The weight of the plastic head may range
from a few ounces to a few pounds. Use the
plastic mallet for straightening thin sheet ducting
or when installing clamps.
Socket Sets
The socket set is one of the most versatile tools
in the toolbox. Basically, it consists of a handle
and a socket-type wrench that can be attached
to the handle. A complete socket wrench set
consists of several types of handles along with bar
made in deep lengths to fit over spark plugs and
long bolt ends.
There are four types of handles used with these
sockets. See figure 2-2. Each type has special
advantages, and the good mechanic chooses the
one best suited to the job at hand. The square
driving lug on the socket wrench handles has a
spring-loaded ball that fits into a recess in the
socket receptacle. The tool design holds the
assembly together. This mated ball-recess feature
prevents the parts of the wrench from falling apart
during normal usage, but a slight pull disassembles
any wrench connection.
extensions, universals, adapters, and a variety of
sockets. See figure 2-2.
SOCKETS.— A socket has an opening cut in
one end to fit a drive on a detachable handle. The
handle drive is usually square. On the other end
of the socket is a 6-point or 12-point opening very
much like the opening in the box-end wrench. The
12-point socket needs to be swung only half as
far as the 6-point socket before it may be lifted
and fitted on the nut for a new grip. It can be used
in closer quarters where there is less room to move
the handle. Most sockets have 12 points. Use the
6-point socket with nuts made of stainless steel.
Stainless steel is a harder metal than that of the
wrench. Extensive use of a 12-point socket on such
nuts or bolts would cause excessive wear on the
12 points. The socket might fail to hold. By contrast, because of the greater holding surface, a
6-point socket holds the stainless steel nut better.
The 6-point socket offers less chance for wear of
the wrench.
RATCHET HANDLE.— The ratchet handle
has a reversing lever that operates a pawl (or dog)
inside the head of the tool. Pulling the handle in
one direction causes the pawl to engage in the
ratchet teeth and to turn the socket. Moving in
the opposite direction causes the pawl to slide over
the teeth, permitting the handle to back up
without moving the socket. This feature allows
rapid turning of the nut or bolt after each partial
turn of the handle. With the reversing lever in one
position, the handle can be used for tightening.
In the other position, it can be used for loosening.
Sockets are classified for size according to two
factors. One is the drive size or square opening,
which fits on the square drive of the handle. The
other is the size of the opening in the opposite end,
which fits the nut or bolt. The standard toolbox
has sockets that have 1/4- and 3/8-inch-square
drivers. The openings that fit the bolt or nut are
graduated in 1/16-inch sizes. Sockets are also
HINGED HANDLE.— The hinged handle is
also very convenient. To loosen a tight nut, swing
the handle at right angles to the socket. This gives
the greatest possible leverage. After loosening the
Figure 2-2.-Typical socket wrench set.
nut to the point where it turns easily, move the
handle into the vertical position, and then turn
the handle with your fingers.
the sliding bar on the T-handle, the head can be
positioned at either the end or the center of the
sliding bar. Select the position that is needed for
the job at hand.
SPEED HANDLE.— After the nuts are
loosened with the sliding bar handle or the ratchet
handle, the speed handle will help remove the nuts
in a hurry. In many instances, the speed handle
is not strong enough to be used for breaking loose
or tightening. Use the speed socket wrench
carefully to avoid damaging the nut threads.
Figure 2-3 .-Screwdriver adapter.
ACCESSORIES.— Several accessory items
complete the socket wrench set. Extension bars
of different lengths are made to extend the handles
to any length needed. A universal joint allows the
nut to be turned with the wrench handle at an
angle. A universal socket is also available, and
universal socket joints, bar extensions, and
universal sockets in combination with appropriate
handles make it possible to form a variety of tools
that will reach otherwise inaccessible nuts and
Another accessory item that comes in handy
is an adapter, which allows you to use a handle
having one size drive with a socket having a
different size drive. For example, a 3/8- by
1/4-inch adapter would make it possible to turn
all 1/4-inch-square drive sockets with any
3/8-inch-square drive handle.
There are special sockets that are used to adapt
various types of screwdriver bits to a speed handle.
See figure 2-3. This socket-type screwdriver is used
to remove recessed head screws from access panels
on equipment.
Figure 2-4 .-Combination wrench.
BOX-END WRENCH.— The box-end fits
completely around the nut or bolt head. The
box-end is usually constructed with 12 points. The
advantage of the 12-point construction is that the
wrench will operate between obstructions where
space for the swing angle is limited. A very short
swing of the handle will turn the nut far enough
to allow the wrench to be lifted and the next set
of points to be fitted to the corners of the nut.
It is possible to use this wrench in places where
the swing angle is limited to as little as 30 degrees.
The box-end portion of the wrench is designed
with an offset in the handle. Notice in figure 2-4
how the 15-degree offset will allow clearance over
nearby parts. One of the best features of the boxend is that there is little or no chance that the
wrench will slip off the nut or bolt. However,
there is the disadvantage of slow work with the
box-end of the combination wrench. Each time
the wrench is backed off, it has to be lifted up
and refitted to the head of the work. To save time,
use the nonslip box-end of the wrench to break
loose tight bolts or to snug up work after the bolt
has been seated with a faster type of wrench that
might slip under stress.
Combination Wrenches
Most toolboxes contain a set of combination
wrenches. As shown in figure 2-4, the combination wrench has an open-end wrench on one end
and a box-end (of the same size) on the other end.
For speed and light stress operations, use the
open-end. Then switch to the box-end for safety
under stress. For ease of explanation, each end
of the wrench is discussed separately. Adjustable
wrenches are also briefly discussed.
OPEN-END WRENCH.— The jaws of the
open-end portion of the combination wrench are
machined at 15 degrees from parallel in respect
to the center line of the handle. This permits the
use of the wrench where there is room to make
only a part of a complete turn. If the wrench is
turned over after the first swing, it will fit on the
same flats and turn the nut farther. After two
swings on the wrench, the nut is turned far enough
so that a new set of flats are in position for the
Use the open end of the combination wrench
on tubing nuts and in cramped places. Sometimes
the cramped space is too small for a socket or boxend to be slipped over the nut or bolt head. When
using any open-end type wrench, always make
sure the wrench fits the nut or bolt head, and pull
on the wrench—never push. Pushing a wrench is
dangerous. The threads could break loose unexpectedly and cause damage to adjacent equipment or injury to the person using the wrench.
wrenches are not intended to replace open-end or
box-end wrenches, but they are useful in working in restricted areas. In addition, they can be
adjusted to fit odd size nuts. Adjustable wrenches
are not intended for standard use but rather for
emergency use. The wrenches were not built for
use on extremely hard-to-turn items. As shown
in figure 2-5, adjustable wrenches have a fixed jaw
(A) and an adjustable jaw (B), which is adjusted
by a thumbscrew (C). By turning the thumbscrew,
the jaw opening may be adjusted to fit various
sizes of nuts. The size of the wrenches ranges from
4 to 18 inches in length. The maximum jaw openings vary in direct proportion to the length of the
Adjustable wrenches are often called “knuckle
busters” because mechanics frequently suffer the
consequences of improper usage of these tools.
There are four simple steps to follow in using
these wrenches. First, choose one of the correct
size. Do not pick a large 12-inch wrench and
adjust the jaw for use on a 3/8-inch nut. This
could result in a broken bolt and a bloody hand.
Second, be sure the jaws of the correct size wrench
are adjusted to fit snugly on the nut. Third,
position the wrench around the nut until the nut
is all the way into the throat of the jaws. If not
used in this manner, the result is apt to be as
bloody as before. Fourth, pull the handle toward
the side having the adjustable jaw. This will
prevent the adjustable jaw from springing open
and slipping off the nut. If the location of the
Figure 2-5.-Adjustable wrenches.
work will not allow all four steps to be followed,
select another type of wrench for the job.
Adjustable wrenches should be cleaned in a
solvent, and a light oil applied to the thumbscrew
and the sides of the adjustable jaw. They should
also be inspected often for cracks, which might
result in failure of the wrench.
Two basic types of screwdriver blades are
used: the common blade for use on conventional
slotted screws, and a crosspoint blade for use on
the recessed head Phillips or Reed and Prince type
of screws. See figure 2-6. The common and crosspoint blade types are used in the design of various
special screwdrivers, some of which are also
shown in figure 2-6.
COMMON SCREWDRIVERS.— The combination length of the shank and blade identifies
the size of common screwdrivers. They vary from
2 1/2 to 12 inches. The diameter of the shank and
the width and thickness of the blade tip, which
fits the screw slot, are in proportion to the length
of the shank. The blade is hardened to prevent
it from being damaged when it is used on screws.
It can easily be chipped or blunted when used for
other purposes. The blade of a poor quality
screwdriver will result in mutilation of the
screwdriver may be used where there is not
sufficient vertical space for a standard
screwdriver. See figure 2-6. Offset screwdrivers
are constructed with one blade forged in-line with
and another blade forged at right angles to the
shank handle. Both blades are bent 90 degrees to
the shank handle. By alternating ends, you can
seat or loosen most screws even when the swinging
space is very restricted. Offset screwdrivers are
made for both standard and recessed head screws.
Many different types of pliers are in use today.
Some of these are the vise grip, the channel-lock,
the duckbill, the needle nose, the diagonal, and
the safety wire pliers.
VISE GRIP PLIERS.— Vice grip pliers can be
used a number of ways. See figure 2-7. These
pliers can be adjusted to various jaw openings by
turning the knurled adjusting screw at the end of
the handle. Vise grips can be clamped and locked
in position by pulling the lever toward the handle.
They may be used as a clamp, portable vise, and
for many other uses where a locking, plier-type
jaw may be employed.
Figure 2-6.-Typical screwdrivers.
screwdriver will sometimes become damaged even
when being used properly. Do not used damaged
When using any type of screwdriver, do
not hold the work in your hand. If the
point slips, it can cause a bad cut. When
removing a screw from an assembly that
is not stationary, hold the work on a solid
surface, in a vise, or with some other
holding tool. Never get any part of your
body in front of the screwdriver point. This
precaution is a good safety rule for any
sharp-pointed tool.
Vise grip pliers should be used with care
since the teeth in the jaws tend to damage
the object on which they are clamped. They
should not be used on nuts, bolts, tube
fittings, or other objects that must be
pliers can be easily identified by the extra-long
are two types of crosspoint screwdrivers in
common use: the Reed and Prince and the
Phillips. The Reed and Prince screwdrivers and
Phillips screwdrivers are not interchangeable.
Always use a Reed and Prince screwdriver with
Reed and Prince screws, and a Phillips screwdriver
with Phillips screws. The use of the wrong
Figure 2-7.-Vice grip pliers.
handles. See figure 2-8. These pliers are very
powerful gripping tools. The inner surfaces of the
jaws consist of a series of coarse teeth formed by
deep grooves. This construction makes a surface
usable for grasping cylindrical objects. Channellock pliers have grooves on one jaw and lands on
the other. The adjustment is effected by changing
the position of the grooves and lands. Channellock pliers are less likely to slip from the adjustment setting when gripping an object. Use the
channel-lock pliers where it is impossible to use
a more adapted wrench or holding device. Many
nuts and bolts and surrounding parts have been
damaged by improper use of channel-lock pliers.
Figure 2-9.-Pliers (A) duckblll; (B) needle-nose.
DUCKBILL PLIERS.— Duckbill pliers have
long wide jaws and slender handles. Duckbills are
used in confined areas where the fingers cannot
be used. The jaw faces of the pliers are scored to
aid in holding an item securely. See figure 2-9,
view A.
DIAGONAL PLIERS.— Diagonal cutting
pliers are an important tool for you to use. They
are used for cutting small wire and cotter pins,
and so forth. Since they are small, they should
not be used to cut large wire or heavy material.
The pliers will be damaged by such use and will
not be effective to cut what they were designed
to cut. They can also be used to remove small
cotter pins where a new pin is to be used when
the work is finished. This is done by gripping the
pin near the hinge of the pliers and lifting up on
the handles, releasing the pin, getting a new grip,
and repeating until the pin is removed.
The inner jaw surface is a diagonal straight
cutting edge offset approximately 15 degrees. This
design permits cutting objects flush with the
surface. The diagonal cutting pliers are not
designed to hold objects. To use enough force to
hold an object, the pliers will cut or deform the
object. The sizes of the diagonal pliers are
indicated by the overall length of the pliers.
pliers are used in the same manner as duckbill
pliers. See figure 2-9, view B, There is a difference
in the design of the jaws, Needle-nose jaws are
tapered to a point, which makes them adapted to
installing and removing small cotter pins. The
pliers have serrations at the nose end and a side
cutter near the throat. Use needle-nose pliers to
hold small items steady, to cut and bend wire, or
to do numerous other jobs that are too intricate
or too difficult to be done by hand alone.
NOTE: Duckbill and needle-nose pliers
are especially delicate. Care should be
exercised when using these pliers, to
prevent springing, breaking, or chipping
the jaws. Once these pliers are damaged,
they are practically useless.
SAFETY-WIRE PLIERS.— When installing
equipment, you may need to lockwire (usually
referred to as safety wire) designated parts of the
installation. The process of lockwiring can be
accomplished faster and neater with the use of
special pliers. See figure 2-10.
Safety-wire pliers are three-way pliers—they
hold, twist, and cut. They are designed to make
a uniform twist and to reduce the time required
in twisting the safety wire.
To operate, grasp the wire between the two
diagonal jaws of the pliers. As the handles are
squeezed together, the thumb and forefinger
brings the outer (locking) sleeve into the locked
position. A pull on the knob of the pliers can
make a uniform twist. The spiral rod may be
pushed back into the pliers without unlocking
Figure 2-8.-Channel-lock pliers.
Figure 2-10.-Safety-wire pliers.
Torque wrenches and micrometers are special
them, and by again pulling on the knob, you can
make a tighter twist. The wire should be installed
snugly but not so tight that the wire is overstressed. A squeeze on the handles unlocks the
pliers, and the wire can then be cut to the proper
length with the cutting jaws.
TORQUE WRENCHES.— There are times
when, for engineering reasons, a definite pressure
must be applied to a nut, bolt, screw, or other
fastener. In such cases a torque wrench must be
used. The torque wrench is a precision tool
consisting of a torque-indicating handle and
appropriate adapter or attachments. Use the
wrench to measure the amount of turning or
twisting force applied to a nut, bolt, or screw.
The three most common torque wrenches are
the deflecting beam, dial indicating, and
micrometer setting types. See figure 2-11. When
using the deflecting beam and the dial-indicating
torque wrenches, the torque is read visually on
a dial or scale mounted on the handle of the
Special Tools
Special tools are normally maintained in a
central toolroom and signed out when needed. A
tool falls into the special category for one of the
following five main reasons.
1. It is an item of special support equipment.
These tools are designed, manufactured, and
issued for supporting or maintaining one particular model of aircraft, engine, or support
2. It is a seldom used tool. When needed, its
use is essential in aircraft maintenance. Most of
the time it would not be required and would just
take up room in the toolbox.
3. It is a high-cost item. A central location is
necessary to permit better use or for security.
4. The large size or awkward shape of the tool
makes it difficult, if not impossible, to put in a
5. It is a instrument type of tool that requires
A wide variety of special tools are furnished
by the manufacturers of the support equipment,
engines, and related equipment. These special
tools are listed in the Allowance List Registers
published by the Aviation Supply Office. Their
use is explained in the manual that covers the
specific support equipment, engine, or item of
equipment for which they were designed. Special
tools that you use frequently may be kept in your
toolbox if permitted by your tool control plan.
Figure 2-11.-Torque wrenches.
wrench. These torque wrenches are all used in a
very similar manner.
To use the micrometer setting type, unlock the
grip and adjust the handle to the desired setting
on the micrometer scale. Relock the grip. Install
the required socket or adapter to the square drive
of the handle. Place the wrench assembly on the
nut or bolt and pull in a clockwise direction with
a smooth steady motion. A fast or jerky motion
will result in an improperly torqued unit. When
the torque applied reaches the torque value
indicated on the handle setting, the handle will
automatically release or “break” and move freely
for a short distance. The release and free travel
is easily felt, so there is no doubt about when the
torquing process is complete.
The following precautions should be observed
when using torque wrenches:
1. Do not use the torque wrench as a hammer.
2. When using the micrometer setting type,
do not move the setting handle below the lowest
torque setting. Place the micrometer at its lowest
setting before returning it to storage.
3. Do not use the torque wrench to apply
greater amounts of torque than its rated capacity.
4. Do not use the torque wrench to break
loose bolts that have been previously tightened.
5. Never store a torque wrench in a toolbox
or in an area where it may be damaged.
Torquing can be described as the twisting
stress that is applied to the fasteners to secure
components together. These fasteners can be nuts,
bolts, studs, clamps, and so forth. Torque values
for these fasteners are expressed in inch-pounds
or foot-pounds. Unless otherwise stated, all
torque values should be obtained with the
manufacturer’s recommended thread lubricant
applied to the threads.
Torque values are listed in the appropriate
section of the applicable instruction manual. In
case there is no torque specified, standard torque
values can be found in the Structural Hardware
Manual, NA 01-1A-8, or in your particular aircraft general information MIM. Regardless of
whether torque values are specified or not, all nuts
in a particular installation must be tightened alike
amount. This permits each bolt in a group to carry
its share of the load. It is a good practice to use
a torque wrench in all applications.
Using the proper torque allows the structure
to develop its design strength and greatly reduces
the possibility of failure due to fatigue. One word
of caution—never rely on memory for torque
information, but look up the correct torque value
each time it is needed. A nut or bolt that is not
torqued to the proper value may cause damage
to the component or equipment.
The proper procedure is to tighten at a
uniformly increasing rate until the desired torque
is obtained. In some cases, where gaskets or other
parts cause a slow permanent set, the torque must
be held at the desired value until the material is
sealed. When applying torque to a series of bolts
on a flange or in an area, select a median value.
If some bolts in a series are torqued to a minimum
value and others to a maximum, force is concentrated on the tighter bolts and is not distributed
evenly. Such unequal distribution of force may
cause shearing or snapping of the bolts.
Torque wrench size must be considered when
torquing. The torque wrenches are listed according to size and should be used within this
recommended range. Use of larger wrenches that
have too great a tolerance results in inaccuracies.
When an offset extension wrench is used with a
torque wrench, the effective length of the torque
wrench is changed. The torque wrench is so
calibrated that when the extension is used, the
indicated torque (the torque that appears on the
dial or gage of the torque wrench) may be
different from the actual torque that is applied
to the nut or bolt. The wrench must be preset to
compensate for the increase when an offset
extension wrench is used.
Occasionally, it is necessary to use a special
extension or adapter wrench together with a
standard torque wrench. To arrive at the resultant
required torque limits, use the following formula:
S = reading of setting of torque wrench;
T = recommended torque on part;
L = length of torque wrench (distance between center of drive and center of hand
grip); and
E = length of extension of adapter (distance
between center of drive and center of
broached opening measured in the same
place as L).
EXAMPLE: Recommended torque is 100
inch-pounds. Using a 12-inch torque wrench and
a 6-inch adapter, determine reading on torque
An example of the measurement of this
formula is shown in figure 2-12. When the extension is pointed back toward the handle of the
torque wrench, subtract the effective length of the
extension from the effective length of the torque
wrench. If the extension is pointed at a right angle
to the torque wrench, then the actual value does
not change.
An engineering study revealed a widespread
lack of understanding as to what happens when
an adapter is used with a torque wrench, To see
if you understand the effects of adapters and
extension handles, read the situation below and
answer the questions.
Figure 2-12.-Torque wrench with extensions.
Situation: Given a dial-indicating torque
wrench, an adapter, and an extension handle,
perform the following steps:
taken. If this is not done, the torque obtained will
be in error.
Torque a nut to a predetermined indicated
value, gripping the torque wrench at the end of
the extension handle. Return to this torque value
twice to ensure the nut is fully seated.
MICROMETERS.— It is important that
a person repairing and building up engines
thoroughly understand the use, and care of the
micrometer. Micrometers are used to set
clearances and to measure damage or repair limits.
Figure 2-13 shows an outside micrometer caliper
with the various parts clearly indicated.
Micrometers are used to measure distances to the
nearest one thousandth of an inch, The measurement is usually expressed or written as a decimal.
You must know the method of writing and
reading decimals.
Remove the extension handle and apply the
same indicated torque again.
Question: Will the nut rotate before the
previously indicated torque is reached?
Answer: Yes! Think it through—if you
decrease the handle length (L), you increase the
torque (T) if the indicated torque (S) remains
Types of Micrometers. — There are three types
of micrometers that are most commonly used
throughout the Navy. They are the outside
micrometer caliper (including the screw thread
micrometer), the inside micrometer, and the depth
Question: What would happen if a micrometer
torque wrench was used?
Answer: Nothing would change. The type of
torque wrench has no effect at all. It is simply a
function of the length of the lever arm between
the torque wrench square drive and where the
hands are placed on the handle.
Using solid-handle torque wrenches with
extension handles can cause significant over- or
undertorquing. This problem exists based on the
handle length chosen for the computation and
where hand force is actually applied on the
handle. When applying the formula, force must
be applied to the handle of the torque wrench at
the point from which the measurements were
Figure 2-13.-Nomenclature of an outside micrometer
micrometer. See figure 2-14. The outside
micrometer is used for measuring outside dimensions, such as the diameter of a piece of round
stock. Use the screw thread micrometer to
determine the pitch diameter of screws. The
inside micrometer is used for measuring inside
dimensions. Use inside micrometers to determine
the inside diameter of a tube or hole, the bore of
a cylinder, or the width of a recess. Use the depth
micrometer for measuring the depth of holes or
Care of Micrometers.— Keep micrometers
clean and lightly oiled. Make sure they are placed
in a case or box when they are not in use. Anvil
faces must be protected from damage and must
not be cleaned with emery cloth or other
The importance of aircraft hardware is often
overlooked because of the small size of most
items. The safe and efficient operation of any aircraft depends upon the correct selection and use
of aircraft hardware. You must make sure that
items of aircraft hardware remain tightly secured
in the aircraft.
Inflight mishaps continue to happen at an
alarming rate. Many of these mishaps are due to
improper hardware selection and installation. For
example, mishaps involving aircraft fires can often
be attributed to the chafing of fluid lines and wire
bundles, caused by improperly clamped parts.
In modern aircraft construction, thousands of
rivets are used. Many parts require frequent
dismantling or replacement. It is more practical
for you to use some form of threaded fastener.
Some joints require greater strength and rigidity
than can be provided by riveting. We use various
types of bolts and nuts to solve this problem.
Bolts and screws are similar in that both have
a head at one end and a screw thread at the other.
However, there are several differences between
them. The threaded end of a bolt is always
relatively blunt. A screw may be either blunt or
pointed. The threaded end of a bolt must be
screwed into a nut. The threaded end of the screw
may fit into a nut or directly into the material
being secured. A bolt has a fairly short threaded
section and a comparatively long grip length (the
unthreaded part). A screw may have a longer
threaded section and no clearly defined grip
length. A bolt assembly is generally tightened by
turning a nut. The bolt head mayor may not be
Figure 2-14.-Common types of micrometers.
designed to be turned. A screw is always designed
to be turned by its head. Another minor difference
between a screw and a bolt is that a screw is
usually made of lower strength materials.
Threads on aircraft bolts and screws are of the
American National Aircraft Standard type, This
standard contains two series of threads—national
coarse (NC) and national fine (NF). Most aircraft
threads are of the NF series.
Bolts and screws may have right- or left-hand
threads. A right-hand thread advances into
engagement when turned clockwise. A left-hand
thread advances into engagement when turned
the nut at least two full threads. If the bolt is too
short, it will not extend out of the bolt hole far
enough for the nut to be securely fastened. If it
is too long, it may extend so far that it interferes
with the movement of nearby parts.
In addition, if a bolt is too long or too short,
its grip will usually be the wrong length. As
shown in figure 2-16, the grip length should be
approximately the same as the thickness of the
material to be fastened. If the grip is too short,
the threads of the bolt will extend into the bolt
hole. The bolt may act like a reamer when the
material is vibrating. If the grip is too long, the
nut will run out of threads before it can be
tightened. In this event, a bolt with a shorter grip
should be used. If the bolt grip extends only a
short distance through the hole, a washer maybe
A second bolt dimension that must be considered is diameter. As shown in figure 2-15, the
diameter of the bolt is the thickness of its shaft.
The results of using a wrong diameter bolt
should be obvious. If the bolt is too big, it
cannot enter the bolt hole. If the diameter is too
small, the bolt has too much play in the bolt hole.
The third and fourth bolt dimensions that
should be considered when you choose a bolt
Many types of bolts are used in modern aircraft, and each type is used to fasten something
in place. Before we discuss some of these types,
it might be helpful to list and explain some
commonly used bolt terms. You should know the
names of bolt parts and be aware of the bolt
dimensions that must be considered in selecting
a bolt.
The three principal parts of a bolt are the
head, grip, and threads, as shown in figure 2-15.
Two of these parts might be well known to you,
but perhaps grip is an unfamiliar term. The grip
is the unthreaded part of the bolt shaft. It extends
from the threads to the bottom of the bolt head.
The head is the larger diameter of the bolt and
may be one of many shapes or designs.
To choose the correct replacement for an
unserviceable bolt, you must consider the length
of the bolt. As shown in figure 2-15, the bolt
length is the distance from the tip of the threaded
end to the head of the bolt. Correct length selection is indicated when the bolt extends through
Figure 2-15 .-Bolt terms and dimensions.
Figure 2-16.-Correct and incorrect grip lengths.
replacement are head thickness and width. If the
head is too thin or too narrow, it might not be
strong enough to bear the load imposed on it. If
the head is too thick or too wide, it might extend
so far that it interferes with the movement of
adjacent parts.
AN bolts come in three head styles—hex head,
clevis, and eyebolt. NAS bolts are available in
countersunk, internal wrenching, and hex head
styles. MS bolts come in internal wrenching and
hex head styles. Head markings indicate the
material of which standard bolts are made. Head
markings may indicate if the bolt is classified as
a close-tolerance bolt. See figure 2-17. Additional
information, such as bolt diameter, bolt length,
and grip length, may be obtained from the bolt
part number. Refer to the Structural Hardware
Manual, NAVAIR 01-1A-8, for complete identification of threaded fasteners.
Figure 2-18.-Non-self-locking nuts.
Aircraft nuts may be divided into two general
groups: non-self-locking and self-locking nuts.
Non-self-locking nuts are those that must be
safetied by external locking devices, such as cotter
pins, safety wire, or locknuts. The locking feature
is an integral part of self-locking nuts.
common of the non-self-locking nuts are the castle
nut, the plain hex nut, the castellated shear nut,
and the wing nut. Figure 2-18 shows these nonself-locking nuts.
Castle nuts are used with drilled-shank AN
hex-head bolts, clevis bolts, or studs. They are
designed to accept a cotter pin or lockwire for
Castellated shear nuts are used on such parts
as drilled clevis bolts and threaded taper pins.
They are normally subjected to shearing stress
Figure 2-17.-Bolt head markings.
only. They must not be used in installations where
tension stresses are encountered.
Plain hex nuts have limited use on aircraft
structures. They require an auxiliary locking
device such as a check nut or a lock washer.
Use wing nuts where the desired tightness can
be obtained by the fingers and where the assembly
is frequently removed. Wing nuts are commonly
used on battery connections.
SELF-LOCKING NUTS.— Self-locking nuts
provide tight connections that will not loosen
under vibrations. Self-locking nuts approved for
use on aircraft must meet critical specifications.
These specifications pertain to strength, corrosion
resistance, and heat-resistant temperatures. Use
new self-locking nuts each time components are
installed in critical areas throughout the entire aircraft. Self-locking nuts are found on all flight,
engine, and fuel control linkage and attachments.
There are two general types of self-locking nuts.
They are the all-metal nuts and the metal nuts with
a nonmetallic insert to provide the locking action.
The elastic stop nut is constructed with a
nonmetallic (nylon) insert, which is designed to
lock the nut in place. The insert is unthreaded and
has a smaller diameter than the inside diameter
of the nut. Its use is limited to engine cold
sections, since high heat could melt the
nonmetallic insert.
Self-locking nuts are generally suitable for
reuse in noncritical applications provided the
threads have not been damaged. If the locking
material has not been damaged or permanently
distorted, it can be reused.
NOTE: If any doubt exists about the
condition of a nut, use a new one!
Installation of Nuts and Bolts
It is of extreme importance to use like bolts
in replacement. In every case, refer to the
applicable maintenance instruction manual and
illustrated parts breakdown. Be certain that each
bolt is of the correct material, size, and grip
length. Examine the markings on the head to
determine whether a bolt is steel or aluminum
alloy. This knowledge is especially important if
you are to use the bolt in the engine hot section.
Use washers under the heads of both bolts and
nuts unless their omission is specified. A washer
guards against mechanical damage to the material
being bolted. A washer also prevents corrosion
of the structural members. An aluminum alloy
washer used under the head and nut of a steel bolt
securing aluminum alloy or magnesium alloy
members will corrode the washer rather than the
members. Steel washers should be used when
joining steel members with steel bolts.
Whenever possible, the bolt should be placed
with the head on top or in the forward position.
This positioning helps prevent the bolt from
slipping out if the nut is accidentally lost.
Figure 2-19.—Various types of washers.
Sufficient friction is provided by the spring
action of the washer to prevent loosening of the
nut because of vibration. Lock washers must not
be used as part of a fastener for primary or
secondary structures.
The star lock or shakeproof washer is a round
washer made of hardened and tempered carbon
steel, stainless steel, or Monel. This washer can
have either internal or external teeth. Each tooth
is twisted, one edge up and one edge down. The
top edge bites into the nut or bolt, and the bottom
edge bites into the working surface. It depends
on spring action for its locking feature. This
washer can be used only once because the teeth
become compressed after being used.
Washers used in aircraft structures may be
grouped into three general classes: plain, lock
washers, and special washers. Figure 2-19 shows
some of the most commonly used types.
Plain Washers
Tab lock washers are round washers designed
with tabs or lips that are bent across the sides of
a hex nut or bolt to lock the nut in place. There
are various methods of securing the tab lock
washer to prevent it from turning. An external
tab bent downward 90 degrees into a small hole
in the face of the unit, an external tab that fits
a keyed bolt, or two or more tab lock washers
connected by a bar are some of the present
methods used. Tab lock washers can withstand
higher heat than other methods of safetying. The
washer can be used safely under high vibration
conditions. Use tab lock washers only once
because the tab tends to crystallize when bent a
second time.
Plain washers are widely used under AN hex
nuts to provide a smooth bearing surface. They
act as a shim in obtaining the correct relationship
between the threads of a bolt and the nut. They
also aid in adjusting the position of castellated
nuts with respect to drilled cotter pin holes in
bolts. Plain washers are also used under lock
washers to prevent damage to surfaces of soft
Lock Washers
Lock washers are used whenever the selflocking or castellated-type nut is not used.
Special Washers
Special washers such as ball-seat and socket
washers and taper pin washers are designed for
special applications. For example, the ball-seat
and socket washer is used where the bolt may be
installed at an angle to the surface. The washer
is also used where perfect alignment with the
surface is required, such as engine mount bolts.
Clamps used on aircraft engines prevent lines
from chafing on parts or against other lines. They
can also connect two lines or pieces of material,
Figure 2-20 shows examples of the Adell clamp
to maintain line clearance and prevent chafing.
The Adell clamp shown in figure 2-20 comes
in two different types. One is made of a rubber
or Teflon® cushion for low-range temperatures.
The second type has an asbestos cushion for high
When installing clamps, be sure to use the
proper size and material. Although clamps may
be reused, make sure reused clamps are in good
condition. Inspect support clamps for deterioration of the cushion material, to prevent the metal
Figure 2-21.-Examples of (A) incorrect and (B) correct
installation of hinged clamp.
Figure 2-20.-Securing lines using support clamps.
part of the clamps from cutting or chafing the
supported line. Carefully inspect clamps for
proper installation. Figures 2-21 and 2-22 shows
two examples of correct and incorrect installations, Figure 2-23 shows how improperly installed
or wrong size clamps hide the damage they cause.
Safetying is a process of securing all aircraft
bolts, nuts, capscrews, studs, and other fasteners.
Safetying prevents the fasteners from working
loose due to vibration. Loose bolts, nuts, or
screws can ruin engines or cause parts of the aircraft to drop off. To carry out an inspection on
an aircraft, you must be familiar with the various
methods of safetying. Careless safetying is a sure
road to disaster. Always use the proper method
for safetying. You should always inspect for proper safetying throughout the area in which you
are working.
There are various methods of safetying aircraft
parts. The most widely used methods are safety
Figure 2-23.-Hidden damage from improperly installed
wire, cotter pins, lock washers, snap rings, and
special nuts. Some of these nuts and washers have
been described previously in this chapter. You
should know how to use safety wire and cotter
Safety Wiring
Safety wiring is the most positive and satisfactory method of safetying. It is a method of wiring
Figure 2-22.-Examples of (A) incorrect and (B) correct installation of Adell clamp.
together two or more units. Any tendency of one
unit to loosen is counteracted by the tightening
of the wire.
Nuts, bolts, and screws are safety wired by the
single-wire double-twist method. This method is
the most common method of safety wiring. A
single-wire may be used on small screws in close
spaces, closed electrical systems, and in places
difficult to reach.
Figure 2-24 shows various methods commonly used in safety wiring nuts, bolts, and
screws. Examples 1, 2, and 5 of figure 2-24 show
proper methods of safety wiring bolts, screws,
square-head plugs, and similar parts when wired
in pairs. Examples 6 and 7 show a single-threaded
component wired to a housing or lug. Example 3
shows several components wired in series.
Example 4 shows the proper method of wiring
castellated nuts and studs. Note that there is no
loop around the nut. Example 8 shows several
components in a closely spaced, closed
geometrical pattern, using the single-wire method.
Figure 2-25 shows safety wire techniques for
T-bolt clamps.
When drilled-head bolts, screws, or other parts
are grouped together, they are more conveniently
Figure 2-24.-Safety wiring methods.
Figure 2-25.-V-band coupling safety wiring techniques.
safety wired to each other in a series rather than
individually. The number of nuts, bolts, or screws
that may be safety wired together depends on the
application. For instance, when you are safety
wiring widely spaced bolts by the double-twist
method, a group of three should be the maximum
number in a series.
When you are safety wiring closely spaced
bolts, the number that can be safety wired by a
24-inch length of wire is the maximum in a series.
The wire is arranged in such a manner that if the
bolt or screw begins to loosen, the force applied
to the wire is in the tightening direction.
Figure 2-26.-Cotter pin installations.
When you use the safety wire method of
safetying, follow these general rules:
Cotter Pins
Use cotter pins to secure bolts, screws, nuts,
and pins. Some cotter pins are made of lowcarbon steel, while others consist of stainless steel
and thus are more resistant to corrosion. Use
stainless steel cotter pins in locations where nonmagnetic material is required. Regardless of shape
or material, use all cotter pins for the same general
purpose— safetying.
1. Torque all parts to the recommended
values, and align holes before you attempt to
proceed with the safetying operation. Never overtorque or loosen a torqued nut to align safety wire
2. The safety wire must be new upon each
3. When you secure castellated nuts with
safety wire, tighten the nut to the low side of the
selected torque range unless otherwise specified.
If necessary, continue tightening until a slot aligns
with the hole.
4. All safety wires must be tight after
installation but not under such tension that
normal handling or vibration will break the wire.
5. Apply the wire so that all pull exerted by
the wire tends to tighten the nut.
6. Twists should be tight and even, and the
wire between the nuts should be as taut as
possible without being overtwisted.
7. A pigtail of 1/4 to 1/2 inch (three to six
twists) should be made at the end of the wiring.
This pigtail must be bent back or under to prevent
it from becoming a snag.
NOTE: Whenever uneven-prong cotter
pins are used, the length measurement is
to the end of the shorter prong.
Cotter pin installation is shown in figure 2-26.
Use castellated nuts with bolts that have been
drilled for cotter pins. Use stainless steel cotter
pins. The cotter pin should fit neatly into the hole,
with very little side-play. The following general
rules apply to cotter-pin safetying:
Do not bend the prong over the bolt end
beyond the bolt diameter. (Cut it off if necessary.)
Do not bend the prong down against the
surface of the washer. (Again, cut it off if
Safety wire comes in different sizes and
material. Use the size and material appropriate
for the job. When using the single-wire method,
you should use the largest size wire that the hole
will accommodate. Different types of safety wire
include Monel for normal use, Inconel for high
temperatures (800- 1500 degrees), and alclad for
nonmagnetic applications.
Do not extend the prongs outward from
the sides of the nut if you use the optional
wraparound method.
Bend all prongs over a reasonable radius.
Sharp angled bends invite breakage. Tap the
prongs lightly with a mallet to bend them.
After completing this chapter, you will be able to:
Recognize the means to identify different types
of support equipment.
Identify the function of the Support Equipment Training and Licensing Program.
Identify the purpose, operation, and safety
precautions in using both powered and nonpowered support equipment.
As naval aircraft have become more complex,
the equipment used to support them has become
more complex. The Aviation Machinist’s Mate
uses many different types of support equipment
(SE) to maintain aircraft in top condition. Some
support equipment such as tow tractors and power
units are common to many different aircraft.
There is also a long list of SE that applies only
to a specific type or model of aircraft. Using SE
correctly is a challenging, sometimes dangerous,
but never routine operation. The support equipment manuals or the maintenance instruction
manuals (MIMs) cover the proper operating procedures and safety precautions for the use of SE.
Read the manuals, learn to use the equipment,
and become qualified on it before you are required
to use it.
each year. Each year the Navy spends millions of
dollars to repair damaged SE and aircraft
caused by the improper use of SE. Navy personnel
are injured, maimed, or killed by improper
use of SE because of failure to follow prescribed
safety precautions. We must do something to
eliminate these tragedies and costs.
This chapter discusses SE identification and
the use of different types of SE. You will
learn about the hazards, safety precautions,
and proper procedures to follow when using
both powered and nonpowered SE. Finally,
you will learn about the SE Training and
Licensing Program, as discussed in OPNAVINST 4790.2.
Over the years safety procedures and precautions for operating SE have developed mainly
from direct experience. Unfortunately, much of
that experience was gained as a result of
accidents. Each of us must be aware that accidents
don’t “just happen.” People cause accidents. We
are all capable of having an accident, for any
number of reasons. Carelessness, complacency,
haste, ignorance, shortcuts, fatigue, and stress are
some of reasons given for SE accidents. It’s
amazing, and a little sad, that the same type
of SE accidents happen over and over again
In previous years, identifying SE has been
somewhat difficult. You learned the designations
and applications of the equipment by association.
You knew that an MD-3 was a tow tractor; so was
a TA-18 and a JG-75. There were several more
tow tractors, but there was nothing in their
designations that showed that they had anything
in common. Support equipment is now undergoing
Figure 3-1.-Equipment indicator code.
a change in designations to group them by application. Newly constructed and modified support
equipment is now identified by MIL-STD-875A.
This designation system for aeronautical and
support equipment will be identical throughout
the military services. Present SE with old designations remain the same until they undergo an
alteration or modification; then they are redesignated.
alignment, or calibration of aircraft systems or
components. Avionic SE include general-purpose
electronic test equipment (GPETE) and automatic
test equipment (ATE). Examples of this type of
SE include multimeter, pressure testers, and fuel
quantity indicator test sets.
Nonavionic SE (common and peculiar) includes all equipment that is nonelectric in nature
and may be powered or nonpowered. Examples
of powered equipment are mobile electric power
plants (NC-10C), aircraft tow tractors
(A/S32A-31), and mobile air-conditioners
(A/M32C-17). Examples of nonpowered SE are
engine stands (4000A) and maintenance
workstands (B-4).
Some time ago the tow tractor, known to the
fleet as the MD-3, underwent the service life
extension program (SLEP). It emerged fleet-ready
and redesignated as the A/S32A-31. The new
system is different from those you may be
accustomed to, but it is a long-needed improvement. After you study the equipment indicators
chart from MIL-STD-875A, you’ll see it isn’t so
difficult after all. See figure 3-1. Figure 3-2 shows
a breakdown of the A/S32A-31.
The most common types of powered SE are
the mobile electric power plants (MEPPs), the
mobile motor-generator sets (MMGs), the mobile
air-conditioners, the gas turbine compressors
(GTCs), and the portable hydraulic power supplies/hydraulic test stands.
Support equipment is all equipment required
on the ground to make an aeronautical system,
system command and control system, support
system, subsystem, or end item of equipment
operational in its intended environment. SE is
primarily that equipment covered by the Aircraft
Maintenance and Material Readiness List
(AMMRL) Program.
Mobile electric power plants supply regulated
electrical power for aircraft servicing, starting,
maintenance, and testing. There are various types
of motor generator assemblies. Some supply dc
power only, while others furnish both dc and ac
The MEPPs used today are designed for
operation on shore stations and aboard aircraft
carriers. On aircraft carriers, these units are
usually mobile with minimum vehicular dimensions
Support equipment is categorized as common
(general-purpose) and peculiar (special-purpose).
SE is normally identified as either powered or
nonpowered. SE types maybe further divided into
the categories of avionic SE and nonavionic SE.
Avionic SE (common and peculiar) includes
all equipment of an electronics nature used for,
but not limited to, the testing, troubleshooting,
Figure 3-2.-Equipment type designation.
115/200-volt, three-phase ac, and 28-volt dc
power. The electrical power controls are located
on two panels at the right side of the operator’s
compartment. The propulsion control and engine
instruments are on the control panel forward of
the steering wheel.
The power plant uses two electronic governors—the engine governor assembly and the drivecontrol module governor assembly. The engine
governor assembly monitors the output frequency
of the ac generator and controls the engine speed.
Speed control is achieved by the governorcontrolled torque motor, which adjusts the engine
internal fuel control. The drive-control module
governor assembly controls the torque motor by
the position of the accelerator pedal when the
START/DRIVE position. The accelerator pedal
uses a variable resistor, not mechanical linkage.
One of the primary hazards of this MEPP is
the unusual driving characteristic. The rear wheel
steering puts the maneuvering part of the vehicle
behind you. It takes lots of practice to become
familiar with rear-wheel steering, and you should
be totally familiar with it before maneuvering
close to aircraft on the flight line.
and weight. They are designed for the utmost
maneuverability and mobility. On shore stations,
these units may be mobile or they maybe mounted
on trailers and require towing. There are many
types of MEPPs in use. The type used depends
upon the type of aircraft to be serviced.
MEPPs, especially the self-propelled type, are
high on the list of SE involved in ground accidents
with aircraft. In addition to the hazards of driving
or towing MEPPs with cables still plugged into
the aircraft, there is the possibility of damage to
the aircraft’s electrical or electronic systems due
to improper electrical operation. High voltage is
certainly a hazard in the use of all MEPPs.
Although protective insulation and covers provide
protection, malfunctions or improper operation
can create electrical shock hazards.
NC-2A Power Plant
The NC-2A is a self-propelled MEPP. See
figure 3-3. It is designed primarily for use aboard
aircraft carriers. This unit is powered by a threecylinder, two-cycle, series 53 Detroit diesel engine
governed at 2,550 rpm. The NC-2A is front-wheel
driven and steered by the rear wheels. The steering
is similar to that of a forklift, and is highly
maneuverable in congested areas. The unit has a
turning radius of approximately 11 feet. The front
wheels are driven by a variable-speed, reversible,
28-volt dc electric motor. It is capable of
propelling the unit up to 14 mph on level terrain.
The diesel engine drives the generators through
a speed-increasing transmission to produce a
NC-8A Power Plant
The NC-8A is a mobile, self-propelled unit
used for servicing and starting rotary- and
fixed-wing aircraft. See figure 3-4. This MEPP
is used primarily aboard shore stations. It is
powered by a four-cylinder, two-cycle, diesel
engine controlled by an electrohydraulic governor.
This unit has one dual-purpose generator
capable of supplying both ac and dc power
simultaneously. It consists of a dc generator and
a synchronous alternator enclosed in one housing.
Figure 3-4.-NC-8A mobile electric power plant (MEPP).
Figure 3.3.-NC-2A mobile electric power plant (MEPP).
incorporated to prevent propulsion unless the
power cables are properly stowed has decreased
this hazard, this accident still occurs because of
defective cable receptacles. Noxious gases, heat,
and exhaust sparks from the diesel engine are
hazardous. Noise from diesel engine SE is a
serious hazard. Most operate above the 90 dBA
level, so hearing protection is a must! Other
hazards associated with the NC-8A include
personnel with poor driving habits or inadvertent
NC-8A movement while close to the aircraft.
The dc generator provides 28 volts, 500
amperes continuously, or 28 volts, 750 amperes
intermittently. The ac generator provides 120/208
volts, three phases, 400 Hz, 60 kVA, and 68 amps
at a .8 power factor when only the ac power is
being used,
All engine controls and instruments are located
directly in front of the operator. Controls and
instruments for the generators are located to the
operator’s right.
Vehicle propulsion is provided by a 28-volt dc,
reversible, variable-speed motor. The motor is
connected to the rear wheels via an automotivetype differential. The speed is controlled by a
variable resistor in the accelerator pedal. The
direction of travel is controlled by a switch
mounted on the engine instrument/control panel.
A primary hazard associated with the NC-8A
is the tendency for the SE operator to drive off
with the power cables still connected to the aircraft. Although a SE change (SEC) has been
NC-10C Power Plant
The NC-10C is designed for shore-based
facilities. See figure 3-5. The unit will supply
regulated electric power up to 90 kVA at .08
power factor, 120/208-volt, three-phase, 400-Hz
ac for servicing, maintenance, and starting of
helicopter and jet aircraft. The unit is powered
by a Detroit diesel six-cylinder, two-cycle engine.
Figure 3-5.-NC-10C mobile electric power plant (MEPP).
Operation of the unit requires a three-phase,
60-hertz, 220- or 440-volt ac external power
source. It is used both aboard ship and ashore.
The unit is not self-propelled and must be
towed or manually moved. The four-wheel trailer
is equipped with tie-down rings, pneumatic tires,
and a tow bar in front for towing or manual steering. A mechanically actuated hand brake, located
on the front of the unit, is connected to
drum/shoe type of brakes on the rear wheels.
Maximum towing speed is restricted to 5 mph.
The MMG-1A weighs 4,120 pounds. It is 7 feet
10 inches long, 4 feet 2 inches wide, and 3 feet
6 inches high.
Stowage compartments are located at the rear
of the unit for 30-foot ac and dc output cables.
The 30-foot input power cable is stowed in a
compartment at the left front side of the unit.
Control panels and a utility power connector
panel are located under fold-open panels on the
right front side of the unit. A portable floodlight,
which can be attached to a mounting bracket atop
the unit, is stored in a door compartment at the
front of the unit.
A portion of the generated ac power is rectified
to supply 28-volt dc at 750 amperes continuously.
The unit is self-contained and requires no
external electrical or mechanical sources of power.
It may be towed at speeds up to 20 mph. The
efficiency of the NC-10C is not affected on
inclines up to 15 degrees maximum in any direction from horizontal. Climatic conditions of
operation are from – 18°F to 120°F ( – 22°C to
50°C) and under relative humidity up to 100
percent. It will operate efficiently at altitudes from
sea level to 8,000 feet.
The general hazards of high voltage, hot
cables, noise, noxious gases, and exhaust heat are
all applicable to the NC-10C. There is also no
lockout circuit to prevent moving the unit with
the cables still plugged into the aircraft.
Mobile motor-generator sets perform the same
function as the mobile electric power plants, but
they are not self-contained and require an external
source of electrical power for operation. The
MMG-1A and MMG-2 are primarily used in
hangars on shore stations, or on the hangar decks
of aircraft carriers where the running of an
internal combustion engine would be objectionable and where external power (220-440/60 Hz)
is readily available.
MMG-2 Motor Generator Set
The MMG-2 is physically quite small and
compact. See figure 3-7. It is a trailer-mounted,
electric motor-driven generator set used to
provide 120/208-volt, 400-Hz ac power, and
28-volt dc power for use in ground maintenance,
calibration, and support for all fighter/interceptor
types of aircraft equipment.
MMG-1A Motor Generator Set
The MMG-1A, shown in figure 3-6, is a small,
compact, trailer-mounted, electric motor-driven
generator set used to provide 115/200-volt, threephase, 400-hertz ac, and 28-volt dc power for
ground maintenance, calibration, and support for
various types of aircraft systems and equipment.
Mobile air-conditioners are designed to
cool, ventilate, dehumidify, and filter the air used
in aircraft cabins and compartments. Although
mobile air-conditioners are usually associated
Figure 3-7.-MMG-2.
Figure 3-6.-MMG-1A.
with them since you will be working around
Some of the hazards for air-conditioning units
are the same as other diesel or electrical support
equipment. These hazards include noise, high
voltage, high-pressure fluids, and exhaust fumes.
In addition, air-conditioners have large, whirling
fans and blowers and refrigerant 22 in both the
liquid and gaseous state.
Refrigerant 22 (R-22) is nonflammable, nontoxic, nonexplosive, and odorless. However, it can
still be dangerous. It can cause serious “burns”
in its liquid state. R-22 vapors displace oxygen in
the air, and if enough is inhaled, it can cause
A/M32C-17 Mobile Air-Conditioner
The A/M32C-17 mobile air-conditioner is a
self-contained, trailer-mounted unit. See figure
All the power-consuming components of the
A/M32C-17 are driven by an industrial diesel
Figure 3-8.-A/M32C-17 mobile air-conditioning unit.
NR-5C Mobile Air-Conditioner
engine. The operating controls are mounted on
the control panel located in a compartment above
the left rear wheel of the trailer. See figure 3-9.
Figure 3-10 describes the operation of each
control on the control panel.
The NR-5C air-conditioner is a mobile, trailermounted unit. See figure 3-11. The NR-5C is
electric powered. The compressor is powered by
Figure 3-9.-A/M32C-17 operating control panel.
1. Outlet air temp.
Indicates outlet air temperature
As desired
2. Clutch indicator light
Light ON indicates clutch engaged
3. Outlet air pressure
Indicates outlet air pressure
1 - 4.5 psig
4. Compressor suction
Indicates compressor suction
48 - 65 psig
5. Compressor discharge
Indicates compressor head pressure
100 - 365 psig
6. Panel light switch
Illuminates control panel
7. Fuel level gauge
Indicates fuel level in large tank
8. Fuel warning light
Indicates large fuel tank is empty;
small fuel tank is low
9. Engine coolant temperature
Indicates engine coolant temperature
Approx 180°F (82°C)
10. Starter switch
Rotates engine to start
Momentary ON
11. Throttle
Controls engine speed (rpm)
(Refer to tachometer)
12. Tachometer and opersting-time meter
Indicates engine speed (rpm) and
elapsed operating time
500 rpm (Idle)
2000 rpm (Normal)
13. Fuel shutoff
Cuts off fuel supply to engine
14. Oil pressure gauge
Indicates engine oil pressure
30 - 60 psig
15. Voltmeter
Indicates operation of alternator and
battery condition
16. Selector switch
Selects operating modes
As desired
17. Emergency stop switch
(Not on units with
HKQ serial numbers)
Stops unit in emergency situations
without allowing for engine cooling or
refrigeration pump down
Move up to stop unit
Figure 3-10.-A/M32C-17 control functions and ranges.
a 440-volt ac, three-phase, 60-cycle,
30-horsepower electric motor, which is an integral
part of the compressor.
The NR-5C is mounted on four wheels. The
two rear wheels are nonsteerable, shock absorbing
on heavy-duty cushion tread tires. Two swivel
shock-absorbing wheels incorporate parking
brakes that are applied or released by a single
manual control lever located at the front of the
unit. Access doors and panels are provided for
Figure 3-11.-NR-5C mobile air-conditioning unit.
full accessibility. Four lifting rings are mounted
on the upper corners of the unit, and four tiedown rings are mounted on the bottom corners
of the unit.
that are controlled by two parking brake handles
mounted on the front of the vehicle.
Gas turbine compressors supply compressed
air for systems requiring large quantities (volume)
of air at low pressure. The gas turbines you will
be using are used to supply air for starting jet aircraft engines. Some of the units also include a
power generating system.
Anytime you are around GTCs you must be
aware of the dangers associated with them. A
noise hazard exists anytime you are around GTCs.
The exhaust gases of the GTC are also very
hazardous. Hot, high-velocity gases can burn you
or any part of the aircraft they hit. If the aircraft
is loaded with external fuel tanks or weapons, the
exhaust can be extremely dangerous. Another
danger comes from the high-pressure air in the
air start hoses. Disconnects or ruptures happen
most often during the initial surge of air through
the hose. A flailing hose end can severely injure
personnel and damage the aircraft.
NR-10 Mobile Air-Conditioner
The NR-10 mobile air-conditioner is also a
trailer-mounted, self-contained, air-conditioning
unit. See figure 3-12. A six-cylinder, turbocharged, diesel engine supplies all the power for
the operation of the air-conditioner.
The engine is liquid cooled by means of a
radiator. Airflow through the radiator is provided
by the condenser fan. The trailer assembly
consists of the tow bar and four independent
suspension wheels. The tow bar assembly is
designed so that when the front wheels attain their
maximum angular position, a cam on the tow bar
assembly is released, allowing the tow bar to
continue following the motion of the towing
vehicle. All four wheels are provided with a
hydraulic braking system on the tow bar for
towing, and all four wheels have parking brakes
GTC-85 Compressor
The GTC-85 gas turbine engine is basically a
two-stage centrifugal compressor directly coupled
to a radial inward-flow gas turbine. Compressed
air is obtained as bleed air from the second stage
of the compressor at a 3.6:1 pressure ratio. This
pneumatic power (bleed air) is used for the operation of large pneumatic equipment, such as jet aircraft turbine starters.
NCPP-105 Compressor
The NCPP-105 is a complete, self-contained
unit consisting of a flyaway assembly enclosed in
a skid-mounted, weather-resistant enclosure.
See figure 3-13. The top view of figure 3-13 shows
the NCPP-105 as a skid-mounted unit. This unit
Figure 3-12.-NR-10 mobile air-conditioning unit.
Figure 3-13.-Model NCPP-105 gas turbine compressor power unit.
assembly, upon arrival at its temporary location,
can be operated by attaching it to a fuel supply.
can be installed on a trailer, as shown in the lower
view of this figure. This permits ease of movement from aircraft to aircraft and from place to
The control panel is part of the flyaway, and
is located on one end of the NCPP-105 unit, as
shown in figure 3-13. The control panel contains
the complete operating instructions for the
operation of the unit.
The NCPP-105 supplies compressed air at two
pressure ratios (5:1 and 3.6:1) for aircraft engine
starting, and ac and dc electrical power for operation of aircraft ac and dc electrical components.
It is equipped with an ac output cable, a dc output cable, and a bleed-air duct assembly.
The unit enclosure consists of a forward and
aft enclosure (hinged together), a cable stowage
compartment, a muffler assembly, a fuel tank
structural assembly, and a base assembly.
Portable hydraulic power supplies/test stands
provide a means of simulating the aircraft’s
hydraulic pump. By connecting a hydraulic power
supply/test stand to the aircraft’s hydraulic
system, the various actuating systems can be
operated without turning up the aircraft engine.
One of the systems that might be checked out this
way is the aerial refueling store. The power supply
is connected to the aircraft system at ground test
couplings (quick disconnects) provided on the
The flyaway assembly is normally operated
while in the NCPP-105 enclosure, with the dc
power supply mounted in the forward enclosure.
However, when it is required to transport the
flyaway assembly by aircraft to a temporary
location, the dc power supply is removed and
relocated on the flyaway assembly structure. The
fuel line and ac and dc electrical output cables are
disconnected, the forward and aft enclosures are
lifted off the structure assembly, and the flyaway
assembly is then removed from the base assembly.
The flyaway assembly, with its remote cable, ac
and dc electrical output cables, and bleed-air duct
The A/M27T-5 portable hydraulic power
supply is a single system hydraulic pumping unit.
See figure 3-14. It is rated at 20 gpm at 3,000 psi
and 10 gpm at 5,000 psi.
Figure 3-14.-A/M27T-5 portable hydraulic power supply.
The power supply is self-contained. It is
designed to check the performance and characteristics of aircraft hydraulic systems. The
A/M27T-5 is capable of performing the following functions:
pump, and lines. Eventually the stand will fail,
either jamming or collapsing.
B-2 Workstand
A type of workstand in common use is the B-2,
shown in figure 3-15. The B-2 consists basically
of a fixed height, 10-foot lower structure; a
variable height upper structure; and a manual
pump-actuated hydraulic system for raising and
lowering the upper structure. The upper structure
includes a work platform with guardrails and steps
with handrails. The platform and steps, because
of parallelogram linkage, stay horizontal
throughout their upward or downward travel. The
lower structure includes fixed steps and handrails,
a tow bar, and four free-swivel caster wheels for
mobility. Each caster is equipped with a safety
locking device containing a spring-loaded pin,
which snaps into notches on the caster pivot axle
to lock the caster swivel. The lower structure also
includes four immobilizing jacks with baseplates.
The jack plates press against the ground and act
as brakes, but not supports, for the structure. You
may find some B-2 stands with the foot-lever
brakes (like the B-4A and B-5A) instead of the
The height range for the B-2 work platform
is from 13 feet to 20 feet. Overall height, including
the 3 1/2-foot guardrails, is 16 1/2 feet lowered
1. Delivering hydraulic fluid at controlled
pressures. This enables operation of the aircraft’s
hydraulic system without the need to start the aircraft’s engines.
2. Testing the flow rate of aircraft hydraulic
3. Testing the aircraft hydraulic system and
components for leakage or malfunction.
4. Flushing and refilling aircraft hydraulic
systems with MIL-H-83282 hydraulic fluid filtered
to 3-micron absolute.
So far we have discussed only powered SE.
This portion of the text will discuss nonpowered
support equipment. Nonpowered SE is all the
equipment that has no engine or motor installed
to supply power for equipment operation.
Maintenance stands, platforms, or workstands
(the names are commonly interchangeable) give
us a means to reach parts of the aircraft we can’t
safely reach or work on from the ground. There
are a large variety of types and models. Some of
the stands are common SE used on almost any
type of aircraft. Others are very large stands used
only at shore activities or on one specific type of
Most adjustable aircraft maintenance platforms are hydraulically operated. A platform and
ladder assembly are mounted on a caster-equipped
base. This enables maintenance personnel to safely
work at heights from 3 feet to a maximum of 20
feet, depending on the stand selected. Since the
design, use, safety precautions, and procedures
are generally very similar, we will cover only a
few of the more common stands.
Most maintenance workstands become defective through abuse and lack of care. Most small
stands are designed to hold 500 pounds safely.
Overloading the stand can cause some part of the
platform structure to bend. That generally causes
the lift structure, or steps to bind. That, in turn,
puts abnormal pressure on the hydraulic cylinder,
Figure 3-15.-B-2 maintenance platform.
and 23 1/2 feet raised. The base structure is 10
feet wide and 14 feet long; however, the upper
work platform extends the length of the whole
workstand to 21 feet when it’s in the lowered
position. The work platform space is 4 feet by 4
feet square. The complete workstand weighs 1,900
The hydraulic system on the B-2 includes a
hand pump, hydraulic lines, a reservoir, and a
hydraulic lift cylinder with a safety lock. The
pump is located on the left-hand angle iron of the
platform. Hydraulic lines lead from the pump to
the lift cylinder reservoir, which attaches to the
scissor section and platform structure. The workstand is raised and lowered by using the pump and
the release valve, the same as a jack.
When the B-2 work platform is raised, the
inner barrel of the hydraulic cylinder is exposed.
This inner barrel has spaced grooves around it to
hold a safety barrel lock. Most of the models have
a barrel lock consisting of a ring with four spring
grips, which rides out on the piston. When the
ring is rotated, cams force the grips out free
of the barrel. When the ring is rotated farther,
the cams allow the grips to press against
the barrel and snap into one of the grooves.
The lock then prevents the cylinder piston from
collapsing in the event of hydraulic failure. You
may run across some models that have a U-shaped
bolt attached to the piston by a chain. This
U-lock is inserted into a barrel groove to lock the
piston up.
B-4A and B-5A Platforms
The two most common maintenance platforms
are the B-4A and the B-5A, as shown in figures
3-16 and 3-17. Both workstands are movable,
hydraulically operated, adjustable platforms with
ladders. They are mounted on free-swivel caster
wheel assemblies. Each wheel has a foot-lever
actuated mechanical brake and a swivel lock
The steel-grated platforms are equipped with
safety rails on three sides, and there are handrails
on the ladder. Both stands are equipped with locking pins that, when inserted through the top of
the platform frame, lock the scissors. This
prevents the platform from collapsing in the event
of hydraulic failure.
Both stands are raised by using a hydraulic
pump, which is located on the platform to the left
Figure 3-17.-B-5A maintenance platform.
Figure 3-16.-B-4A maintenance platform.
trailer, model 3000B. The removal and positioning (or installation/removal) trailer, as the name
shows, is used to remove and install engines and
move them for short distances. The transportation trailer is used for transporting engines over
longer distances and to transfer engines from
other pieces of the matched rail ground-handling
system. Actual work on the engine is normally
performed after it is transferred to the engine
workstand. The workstand is usually in a fixed
location in the hangar or shop.
of the ladder. The stands are lowered by using
the hydraulic release valve on the pump.
The major difference between the B-4A and
the B-5A stands is their size and height range. The
B-4A extends for a working height of between
3 and 7 feet. The B-5A extends for a working
height of between 7 and 12 feet. Both stands have
a capacity of 500 pounds. The B-4A is 8 feet long,
3 feet wide, and weighs 460 pounds. The B-5A
is 8 feet 4 inches long, 8 feet wide at the base, and
weighs 860 pounds.
3000B Trailer
Other Maintenance Platforms
Figure 3-18 shows the 3000B trailer. The unit
is a four-wheel trailer incorporating a detachable,
telescopic tow bar at the front and a tow coupling
at the rear. The twin parallel rails are equipped
with male and female quick-disconnect couplings
and spring-loaded roller adapter stops on both
ends of each rail. The rails can be mated to the
model 4000A or B engine removal stand or the
model 3110 engine workstand.
There are many more types of workstands
available to you from both Navy and commercial
sources—from foot-high work stools to step
stools, stepladders, and phase platforms. These
are generally designed for the specific jobs they
are used for and incorporate the strength, ruggedness, and features required for safety. If used
properly and with care they are safe.
What isn’t safe is anything that wasn’t
designed as a ladder or workstand; such as folding
steel chairs, swivel or even solid chairs, boxes,
card tables, cans, barrels, drums, tractor hoods,
or the top of any other SE. There are a hundred
other things that people try to use every day
instead of proper workstands. These substitutes
are usually available and convenient, although
they are NOT safe. They are dangerous and cause
a tremendous number of falls and disabling
The main purpose of the 3000B trailer is to
move or transport engines for short or long
distances, such as hangar to hangar or from
squadron to the aircraft intermediate maintenance
department (AIMD). The trailer is one part of the
universal matched rail ground-handling system.
The trailer weighs 600 pounds and has a load
carrying capacity of 8,000 pounds. It is equipped
with pneumatic tires inflated to 30 psi. The rails
are 12 feet 8 inches long. The overall trailer is 2
feet 10 inches high and 6 feet wide.
Since the days of the early axial flow turbojet
engines, the Navy has moved toward universal
engine installation, removal, and transportation
trailers and workstands. These basic trailers and
engine workstands are a matched rail groundhandling system that can be modified to handle
different types of engines, installations, and aircraft by the use of various peculiar support equipment (PSE) adapters and, in some cases, hoisting
The equipment in common use today are the
engine removal and positioning trailer, models
4000A and 4000B, and the engine transportation
Figure 3-18.-3000B engine transportation trailer.
4000A and 4000B Trailers
Connecting lines and fittings
These two models are very similar. Figure 3-19
shows the model 4000A. It is a four-wheel,
mobile, hydraulically controlled, self-supporting
unit. The trailer consists of a main frame
supported by four wheels, a lift linkage system,
an upper frame holding two cradle assemblies,
and a tube and rail assembly. A detachable,
telescopic tow bar provides a means of manually
steering or towing the trailer. Some trailers may
be equipped with a tow coupling on the rear.
The hydraulic system consists of the following:
Foot-lever actuated drum/shoe types of parking brakes are located on the two rear wheels.
Large foot assemblies, which can be manually
lowered, are provided to give the stand maximum
stability and support when required. The tie rods
that hold the rear wheels fore and aft, and those
that control tow bar steering of the front
wheels, are configured so that they can easily
be disconnected. This permits all four wheels
to be manually positioned for maximum maneuverability in close quarters.
Four hydraulic frame lift rams that raise
and lower the upper frame assembly (rails)
All four wheels are attached to the main frame
by hydraulically controlled wheel support arms,
operated by wheel lift rams. A ratchet and pawl
system is provided on the rams to safely lock the
rams, mechanically and automatically, as they
extend. Pawl handles on each wheel lift cylinder
must be actuated and held to permit the rams to
Four (two on some models) wheel lift rams
that raise and lower the main frame
Two hand pumps with release valves that
operate either the lift rams or the wheel rams
The wheel lift rams permit raising or lowering
the main (lower) trailer frame. The main frame
can be lowered right to the deck, provided the four
manual foot assemblies are all the way up. The
main frame full up position gives maximum
ground clearance and is used when towing or
A two-position selector valve labeled LIFT
A hydraulic fluid reservoir (two on some
Figure 3-19.-4000A engine removal/installation trailer.
moving the trailer, particularly when loaded. The
forward hydraulic pump and release valve raise
and lower the front end of the main frame. The
aft pump and release valve raise and lower the
rear end of the main frame. Operated together,
the pumps or release valves raise and lower the
whole main frame at once.
or light load. Some stands have two hydraulic
Some stands have only two, instead of four,
wheel lift rams. The pumps on some stands are
located on top of the left side main frame, whereas
some pumps are inside the main frame member.
The 4000A and 4000B trailers should never be
used to transport engines, even for short distances.
The lift linkage consists of four upper and four
lower links, centrally hinged in a jackknife
position. The linkage system is raised and lowered
by four frame lift rams. These lift rams are also
equipped with a ratchet and pawl system to provide a safe mechanical lock in case of a hydraulic
system leak or failure. Pawl knobs located on all
four upper links must be actuated and held to
permit the rams to retract.
3110 Workstand
The model 3110 workstand is a 49-gauge
matching rail-type unit designed to mate with railtype trailers for the roll transfer of the engine.
Model 3110, usually located in the hanger
or power plants work center, allows for the
horizontal maintenance and storage of aircraft
engines. These stands can be used on any hard
surface, and are easily erected and maintained.
See figure 3-20.
The upper frame is attached to the lift linkage
system and holds a cradle assembly at each end.
Inside each cradle are two rollers upon which the
semicircular support tubes, holding the two
parallel rails, can roll (rock from side to side). A
rotation adjustment knob, located on the left side
of the forward support tube, permits ±10 degrees
of roll adjustment of the rails. Two traverse
adjustment knobs, located on the left side of each
cradle assembly, permit ±3 inches of horizontal
lateral (side) movement of the rails. Yaw adjustments up to ±2.25 degrees left or right of the
center line of the rails can be made using just one
of the traverse adjustment knobs, or both in
different directions.
The Aviation Machinist’s Mate has a requirement to use special support equipment to
accomplish tasks such as engine removal and
corrosion control purposes. The aero bomb hoist
and the jet engine corrosion control carts are
examples of this special-purpose gear.
A mechanical winch assembly with a hand
crank is provided on the unit for moving an engine
onto the trailer when the rails are tilted. The ends
of the two upper rails are equipped with male and
female quick-disconnect couplings to permit
mating and load transfer from one type of stand
to another. Roller adapter assemblies are provided
on the rails to hold the engine. The roller adapters
can be locked in any position on the rail to center
the load. The very ends of the rails are equipped
with spring-loaded stop pins to prevent an engine,
on unlocked roller adapters, from accidentally
rolling off the rails.
Aero Bomb Hoists
Aero bomb hoists are used in conjunction with
the air logistic trailer for aircraft engine removal
The model 4000A stand has been in service for
many years. During that time many modifications
and changes have resulted in various configurations of stands in the fleet. The U.S. Air Force
also uses this stand. Some stands have pneumatic
tires, and some have solid rubber tires. Some
hydraulic pumps have selector collars for heavy
Figure 3-20.-Model 3110 workstand assembly.
and installation in some aircraft. See figure 3-21.
Prior to using the bomb hoist, a PREOPERATIONAL
inspection, which includes checking the cable
for frayed or broken strands, must be conducted. Always be sure that the hoist load
test date is current and that the cable is routed
The primary components of the unit are a
large solution tank, two air cylinders, a work
platform with guardrail, four spray applicator
wand assemblies, and the trailer.
A drawbar at the front of the trailer
provides towing and steering capabilities.
It also incorporates a spring loaded “deadman” brake. If the drawbar is released
from the horizontal towing position, it
returns to the vertical position with considerable force. If a person is unaware of
this feature when disengaging the tow bar
from a tractor, there is the possibility of
personnel injury.
Never leave an engine unattended while it
is being supported by hoists. Never work
or get under an engine while it is being
supported by hoists. When lowering or
raising an engine, do it slowly. Constantly
check the engine clearance with the aircraft
nacelle and controls to prevent damage or
The 33-gallon solution tank is separated into
two separate compartments. The forward section
is a 7-gallon preservative tank. The rear section
is a 26-gallon freshwater tank. Each tank has its
own filler neck and cap. There is a 4-inch opening
for water and a 2-inch opening for the preservative
tank. Each tank has a pipe plug at the bottom
for draining. The freshwater and preservative
fluid systems each have a shutoff valve, a quick
Jet Engine Corrosion
Control Cart
The corrosion control cart provides freshwater
rinsing or the application of preservation compound to the compressor section of an engine
through a low-pressure spray. See figure 3-22.
Figure 3-21.-Aero bomb hoist.
work platform on top of the unit provides a work
platform to help operators reach the intakes of
helo engines.
The purpose of the SE Training and Licensing
Program is to make sure you receive effective
training in the safe and efficient operation of SE
on specific types of aircraft. The improper use of
SE has resulted in excessive ground accidents and
repair and replacement costs amounting to
millions of dollars each year. It also results in
reduced operational readiness. The major reasons
for improper use of SE are lack of effective training and lack of effective supervision.
Proper licensing of SE operators takes the
coordinated effort of both the IMA SE division
and the user activity. OPNAVINST 4790.2 (latest
edition) lists the procedures and responsibilities
required for the training and licensing of support
equipment operators.
The SE Training and Licensing Program has
two distinct parts. Part one, taught by the
supporting IMA, covers the proper operation or
use of the SE. Part two, taught by the user
activity, consists of on-the-job training (OJT),
practical exams, and written tests to operate the
SE on a specific type/model/series of aircraft.
Once this training is accomplished and documented, the division officer initiates an SE
operator’s license and forwards it for approval.
Figure 3-22.-Corrosion control cart.
acting lever valve, and an applicator wand on/off
Two 500 cubic-inch, 3,000 psi, air cylinders
mount on the left side of the unit to supply air
pressure to pump freshwater or preservative from
the storage tanks. Four applicator wands are
stored on the right side of the unit. Each wand
is about 6 feet long and attaches to the two 10-foot
supply hoses off the storage tanks. A steel grate
After completing this chapter, you will be able to:
Identify the types of fuels used in aviation and
recognize the operational properties required
in aviation fuels.
Identify the types and operation of engine fuel
system parts.
Identify procedures for troubleshooting
malfunctions in airframe and engine fuel
Identify the parts and recognize a typical airframe fuel system. Also recognize the safety
precautions required in airframe fuel system
Identify the major safety precautions, and
recognize the basic procedures for fuel cell
removal and installation.
The jet engine fuel system usually includes an
emergency system to supply fuel to the engine in
case of main system failure. In some cases, the
emergency system is a duplicate of the main
system. However, in others the emergency system
is not fully automatic and must be controlled
by the pilot. With the nonautomatic type of
emergency system, the pilot must accelerate and
decelerate slowly, or there will be danger of rich
blowout, lean blowout, stall, or overheating of
the combustion and turbine areas.
Both the airframe fuel system and engine fuel
system, as well as the type, designation, and
requirements of the aircraft fuels, are discussed
in this chapter.
The purpose of the fuel system is to deliver
a uniform flow of clean fuel under constant
pressure to the engine under all operating
conditions. To accomplish this task the fuel
system must be properly maintained. The AD’s
responsibility to maintain and troubleshoot the
fuel system include the following:
1. Knowing the different types of fuels and
their characteristics.
2. Knowing the different types of aircraft and
engine fuel systems and their parts.
In general, aircraft fuel systems are divided
into two parts:
1. The airframe fuel system consisting of fuel
tanks, float-operated transfer valves, selector and
shutoff valves, and fuel tank boost pumps.
2. The engine fuel system includes some
combination of different parts. These parts are
filters, fuel control units, engine-driven fuel
pumps, flow dividers, pressurizing valves, drain
valves, afterburner fuel controls, and fuel nozzles
or injectors.
For many years it was popularly believed that
gas turbine engines could burn any type of fuel,
from crude oil to aviation grade 115/145 gasoline.
Of course, this is not true. Because of the wide
range of operating conditions and high rate of fuel
consumption, jet engines require specific fuels to
operate efficiently and maintain a reasonable
engine service life. Various grades of jet fuels were
developed to meet specific operating or handling
characteristics. A study of the basic characteristics
of turbine fuels will help you understand the
importance of delivering the proper fuel to the
aircraft. Such a study is also valuable in
understanding the need for safety and caution in
handling these fuels. This section includes the
basic characteristics of engine fuels.
Volatility measures the ability of a liquid to
convert to a vaporous state. Fuel must vaporize
and the vapor be mixed in a given percentage of
air for it to burn or explode. Only fuel-air
mixtures within the flammable range will burn
(fig. 4-1). Volatility of a fuel effects starting,
range, and safety. A highly volatile fuel starts
easier (especially at low temperatures or under
adverse conditions) and has less range (due to fuel
evaporation in flight). The fuel has a higher
tendency to vapor lock and is more susceptible
to a fire during a crash. The volatility of a
petroleum fuel is usually measured in terms of
vapor pressure and distillation.
The vapor pressure shows the tendency to
vaporize at specific temperatures. Vapor pressure
is measured in a Reid Vapor Pressure Test Bomb.
In the test, one volume of fuel and four volumes
of air are contained in a sealed bomb fitted with
a pressure gauge. The container and fuel are
heated to 100°F, shaken, and then you read the
pressure on the gauge. The pressure shown on the
gauge is known as the Reid Vapor Pressure (RVP)
and is expressed in pounds per square inch (psi).
The higher the pressure the more volatile the fuel.
The distillation measurement for volatility
measures the amount of fuel boiled off at specific
temperatures. Since turbine fuels are a mixture
of hydrocarbons (gasoline and kerosene), they
have a wide range of boiling points. This
test records the boiling ranges. The military
Aircraft engine fuels are petroleum products
manufactured from crude oil by oil refineries.
They are classified as inflammable liquids.
Any material easily ignited that burns rapidly is
inflammable. (NOTE: The terms flammable and
inflammable mean the same.) Under proper
conditions, fuel can explode with force similar to
dynamite. Death can result if the vapors of fuel
are inhaled in sufficient quantities. Serious skin
irritation can result from contact with the fuel in
the liquid state. In liquid form, aircraft fuels are
lighter than water, and in vapor form they are
heavier than air. Consequently, water in the fuel
usually settles to the bottom of the container. And
vapors of these fuels, when released in the air,
tend to remain close to the ground, thus increasing
the danger to personnel and property. From safety
and health standpoints, aircraft engine fuels must
be handled with caution.
In the selection of a fuel, several factors must
be considered. Since one fuel cannot have all the
requirements to the greatest degree, the fuel
selected is a compromise of various factors.
Specific properties of fuels are determined
through testing. These tests determine the
volatility, density, heating value, combustion,
safety, and handling characteristics of the fuels.
There are hundreds of test that determine the
physical, chemical, and performance properties
of fuel. We limit this discussion to the most
common and important ones as follows:
1. Volatility (vapor pressure and distillation
2. Flash point and fire point
3. Heat energy content
4. Viscosity
5. Handling characteristics
6. Combustion products
7. Effects of additives and impurities
8. Freeze point
Figure 4-1.-Vaporization of aviation fuels at atmospheric
Handling Characteristics
specification for fuels will give these temperatures
and the percentages of the fuel allowed to boil
off to meet the desired standards.
For a fuel to have satisfactory handling
characteristics, it must be noncorrosive and should
not clog fuel filters, even at very low temperatures.
The fuel should not produce vapor lock in the fuel
tanks or in the various fuel pumps or slugging out
of the fuel tank vents. (Slugging is the process by
which liquid fuel is carried along with vaporized
fuel when the vapor escapes to the atmosphere.)
As far as possible, the fuel should have enough
of the properties of a lubricant to avoid significant
wear of the fuel-metering pumps.
Flash Point and Fire Point
The flash point is the temperature at which
the fuel vaporizes enough to ignite with an outside heat source. The flash point of a fuel is an
index of its potential safety for handling and
storage. Ships require at least a 140°F flash point
for storage for safety reasons. The fire point is
the temperature where the vapors continue to burn
without an outside heat source.
Combustion Products
Aircraft fuels must have a minimum tendency
to form solids or carbon on combustion. A loss
in the efficiency of the engine results when these
deposits build up in the engine.
Heat Energy Content
For aircraft engine use, it is important that the
fuel contain as much heat energy (thermal value)
as possible, both per unit weight and per unit
volume. The thermal value is the amount of heat
produced as a result of complete combustion and
expressed in calorie or British thermal units (Btu).
Additives, Impurities, and Their Effects
Only materials that will be effective when
added in a maximum concentration of 5 percent
are considered as liquid additives. Beyond this
concentration, the material may be considered as
a fuel.
NOTE: A calorie is the amount of heat
needed to raise the temperature of 1 gram
of water 1 degree Celsius. A Btu is the
amount of heat required to raise the
temperature of 1 pound of water 1 degree
Fahrenheit. One Btu equals 252 calories.
Gum inhibitors used in military gas turbine
fuels are the same as those used for military
aviation gasolines. In aviation gasoline, gum is
almost always completely soluble and becomes
apparent only when the gasoline is evaporated.
Both soluble and insoluble gum, especially the
insoluble form, can be expected to have serious
effects on the fuel system of the turbine engines.
The fuel-metering pumps, fuel pumps, and fuel
filters are likely to be seriously affected by
insoluble gum. The soluble type can be expected
to cause difficulty in the fuel system, at points
where microscopic leakage occurs and exposes
thin films of fuel to air, and thus to evaporation.
The microscopic fuel leaks will usually appear at
fuel valves.
Thermal value per unit of weight increases as
gravity increases. Energy content and density
influence fuel selection when range or payload are
the limiting factors. This is important to understand when the aircraft will be weight-limited
rather than volume-limited.
Viscosity is the internal resistance of a liquid
that tends to prevent it from flowing. Turbojet
engine fuels should be able to flow through the
fuel system and strainers under the lowest
operating temperatures to which the engine will
be subjected. Fuel viscosity and density also have
considerable effect on nozzle performance,
especially when varied over a wide range. The
most important fuel property influencing nozzle
performance is viscosity. It affects drop size, flow
range, and spray angle. Changes in fuel density
affect fuel flow.
Certain aircraft require a minimum concentration of fuel system icing inhibitors (FSII).
This is put in the fuel to prevent icing in the airframe fuel system, engine filter, or engine fuel
control. FSII materials are considered to be
dangerous before their additions to fuel;
therefore, shipboard injection is not approved.
Freeze Point
store, thermally stable, and has a high heat
content per gallon. JP-5 is a kerosene-type fuel
with a vapor pressure close to 0 psi. Its high flash
point makes it safe for shipboard handling. In
fact, it is the only jet aircraft fuel used aboard
ships. It has a lower tendency to vaporize than
the more volatile grades. The vapor-air mixture
in tanks or containers above its liquid surfaces will
generally be too lean to be ignited until the surface
of the liquid reaches a temperature of about
The freezing point of a fuel is the temperature
at which solid particles begin to form in the fuel.
These particles are waxy crystals normally held
in suspension in the fuel. These particles can
readily block the filters in an aircraft fuel system.
The fuel almost always becomes cloudy before the
solid particles form. This cloud is caused by
dissolved water coming out of the solution and
JP-8. JP-8 (NATO Code F-34) is similar to
JP-5 in most characteristics, except flash point
and freeze point. JP-8 is now available only in
Europe. JP-8 represents significant advantages
over JP-4 in fuel handling and operational safety.
Although, like JP-4, its flash point is lower than
shipboard safety standards. The disadvantages of
cost, availability y, and low temperature starting
problems prevent it from replacing JP-4.
The U.S. Military grades of jet fuel are
designated by the letters JP followed by a
number. The grade number merely shows the
approximate sequence the fuel specifications
were accepted by the military. NATO codes
show compatible fuel standards. When changing
to a different fuel, it is usually not necessary
to drain out the old fuel. Some aircraft
prohibit fuel mixing or require different settings
on some fuel components (fuel controls) when
switching fuel grades.
Commercial Fuels. Common commercial fuels
used include types A, A-1, and B. Commercial
fuels are authorized for use in military aircraft
when JP fuel is not available. The characteristics
of commercial fuel are similar to military fuels.
A-1 is designated NATO code F-34, or equal to
JP-8. Jet A is equal to JP-5, and Jet B is equal
to JP-4.
JP-1. JP-1 was the original kerosene-type fuel
used in turbine engines. Its characteristics were
low vapor pressure and high energy content per
volume of fuel.
JP-3. JP-3 was a mixture of 65 percent
gasoline and 35 percent kerosene. Because of its
high percentage of gasoline, its characteristics
were similar to gasoline. This included low flash
point (–40°F), easy cold weather starting, and
poor lubricating qualities. There is also a high fuel
loss due to evaporation and a tendency to vapor
The complex fuel systems of modern aircraft do not function properly if the fuel is
contaminated with dirt, rust, water, or other
foreign matter. Very small quantities of water may
form into ice at altitude affecting small fuel
control orifices. Contaminated fuel has caused aircraft accidents with a tragic loss of life and
valuable aircraft. This means clean fuel is a LIFEOR-DEATH matter with aviation personnel.
JP-4. JP-4 (NATO Code F-40) is an alternate
fuel to JP-5 for USN jet aircraft used at shore
stations only. It is never used on ships. Its low
vapor pressure reduces fuel tank losses and vapor
lock tendencies. Its fuel density is 6.5 pounds per
gallon, and its flash point is below 0°F. When
switching to JP-4 from JP-5, engine operating
characteristics may change. Changes include easier
starting, slower acceleration, lower operating
temperature, and shorter range.
Besides being deadly, contaminants are also
sneaky. A certain type of emulsion resulting from
water and rust particles can adhere to the sides
of aircraft’s fuel cells and go undetected, even
with fuel sampling. It will continue to build up
until parts of it wash off, blocking fuel filters,
lines, or fuel control passages. Contamination
causes unnecessary man-hours in troubleshooting
and fixing fuel problems and possible engine
JP-5. JP-5 (NATO Code F-44) is now the
Navy’s primary jet fuel. It is relatively safe to
In addition to causing extra maintenance and
engine failure, fuel contamination causes serious
delays in flight operations. Contaminated fuel
must be tracked back to the source of contamination and the problem corrected. Until the cause
of contamination is found and corrected, the
contaminated system cannot be used. The fuel
system may be a mobile refueler, air station
hydrant refueling system, or the entire fuel system
of an aircraft carrier. Contaminated fuel could
affect one aircraft or the operation of an entire
air wing.
Part per million is the reference for water
contamination. If you take a 32-ounce sample
bottle and fill it 3 1/4 inches from the bottom,
you have about 500 cubic centimeters (cc).
Break the 500 cc down into one million little
pieces. You now have 1ppm. Of course, you
must use accurate surveillance equipment to
perform measurements that small. Normally,
the organizational maintenance level does not
require this precise testing and inspection.
Instead, the organizational level visually inspects
fuel samples for contamination.
Measuring Contamination
Types and Limits of Contamination
How do you determine how much contamination is too much? First, you have to understand
the units of measurements used to identify
contamination. The two units for measuring
contamination are microns for solids and parts
per million (ppm) for water.
Acceptable fuel is clean and bright with no
visually detected free water. The terms clean and
bright have no relation to the natural color of the
fuel. Jet fuels are not dyed and vary from clear,
water white to straw-yellow in color. Clean means
the absence of any cloud, emulsion, visible
sediment, or free water. Bright means the fuel has
a shiny, sparkling appearance. A cloud, haze,
specks of particulate matter, or entrained water
indicates contaminated fuel that cannot be used.
Steps must be taken to find the source of
There are about 25,400 microns in 1 inch.
Figure 4-2 gives you a microscopic view of
a human hair compared with small particle
Figure 4-2.—Enlargement of small particles and comparison to human hair.
Fuel suspected of microbiological contamination
must not be defueled into a clean system.
contamination and correct it. Figure 4-3 shows
an acceptable sample and common types of
contamination usually detected visually.
Sampling Procedures
WATER.— Water in fuels is either fresh or
saline and present as dissolved or free water.
Dissolved water is water in the fuel that is NOT
visible. Free water is a cloud, emulsion, droplets,
or gross amounts in the bottom of the container.
Any form of free water could result in icing,
corrosion, or malfunctioning of fuel system parts.
Saline water will cause corrosion faster than fresh
Fuel samples are taken from the fuel cell lowpoint drains as specified in the applicable
Maintenance Requirement Card (MRC) deck.
NOTE: Obtain fuel samples prior to
refueling. Only trained personnel shall take
fuel samples; personnel taking samples
must have clean hands. Improper containers and poorly drawn or mishandled
samples result in meaningless or misleading
SEDIMENT.— Sediment appears as dust,
powder, fibrous material, grain, flakes, or stain.
Specks or granules of sediment indicate particles
in the visible size (fig. 4-2) of about 40 microns
or larger. In a clean sample of fuel, sediment
should not be visible except upon the most
meticulous inspection. Sediment or solid contamination is either course or fine.
Course sediment is 10 microns or larger in size.
Course particles can clog orifices and wedge in
sliding valve clearances and shoulders, causing
malfunction and excessive wear. They can also
clog nozzle screens and other fine filter screens
throughout the fuel system. Fine sediments are
less than 10 microns and are not visible as distinct
or separate particles. They appear as a dark
shellac-like surface on sliding valves.
1. Ensure exterior of low-point drain is
cleaned prior to sampling.
2. Drain off 1 pint from low-point drain,
using a 1-quart, clear glass or polyethylene
3. Inspect sample for loose drops of water
puddled under fuel.
NOTE: If dark stringy or fibrous material
that tends to float in the fuel is noted in
any sample, forward the sample(s) to the
nearest Navy Petroleum Laboratory for
microbiological growth determination.
4. If water is detected, discard sample and
repeat steps 1 and 2 until no water is found.
5. Swirl the sample by briskly rotating the
6. If water is present under the swirling
vortex, draw another sample and reinspect.
7. Inspect fuel sample for any discoloration,
cloudiness, and loose sediment under the swirling
8. If small amounts of particulate material are
noted, discard the sample, draw another sample,
and reinspect.
9. If relatively large quantities of water or
foreign matter are noted, or small amounts persist
from one or more cell drains, perform the
a. Keep the fuel sample.
b. Immediately notify maintenance control, who will ground the aircraft and notify the
quality assurance division to perform an investigation to determine the source of contamination.
c. If the source of contamination is not
isolated to the aircraft, notify the cognizant fuel
handling activity. The source of contamination
must be identified. See table 4-1 for types of
contamination and limits.
Microbiological growth consists of living
organisms that grow at a fuel/water interface.
These organisms include protozoa, fungus, and
bacteria. Fungus is the major constituent causing
many of the problems associated with biological
contamination of jet fuels. Fungus is a vegetable
life that holds rust and water. It is also a
stabilizing agent for fuel-water emulsion.
Microbiological growth can develop wherever free
water exists in the fuel tanks. Traces of metallic
elements are also necessary, but water is the key
ingredient. Remove free water and growth ceases.
Microbiological growth is a brown, black, or gray
color and has a stringy, fibrous-like appearance.
It clings to glass and metal surfaces, causing
problems such as severe corrosion or erratic
operation of fuel system components. Microbiological growth causes erroneous readings in
fuel quantity systems, sluggish fuel control
operation, and clogging of filters. It is more
prevalent in tropical and semitropical locations
because of higher temperatures and humidity.
Airframe fuel system maintenance is the
responsibility of more than one work center. For
instance, ADs remove and install bladder and selfsealing fuel cells. Personnel of the AM rating
perform the repairs on integral tanks. Personnel
from the AO rating usually help in the installation and removal of external tanks (drop tanks).
To maintain the aircraft fuel system pertaining
to the AD rating, you must be familiar with the
aircraft fuel system as well as the engine fuel
To meet the particular needs of the various
types of aircraft, fuel tanks vary in size, shape,
construction, and location. Sometimes a fuel tank
is an integral part of a wing. Most often fuel tanks
are separate units, configured to the aircraft
design and mission.
The material selected for the construction of
a particular fuel tank depends upon the type of
aircraft and its mission. Fuel tanks and the fuel
system in general are made of materials that will
not react chemically with any fuels. Fuel tanks
that are an integral part of the wing are of the
same material as the wing. The tank’s seams are
sealed with fuelproof sealing compound. Other
fuel tanks may be synthetic rubber, or self-sealing
cells or bladder-type cells that fit into cavities in
the wing or fuselage of the aircraft.
Fuel tanks must have facilities for the inspection and repair of the tank. This requirement is
met by installing access panels in the fuselage and
wings. Fuel tanks must be equipped with sump
and drains to collect sediment and water. The
construction of the tank must be such that any
hazardous quantity of water in the tank will drain
to the sump, so the water can be drained from
the fuel tank. The AD should be familiar with the
different types of fuel tank/cell construction as
described in the following paragraphs.
Self-Sealing Fuel Cells
A self-sealing cell is a fuel container that
automatically seals small holes or damage caused
during combat operations. A self-sealing cell is
not bulletproof, merely puncture sealing. As
illustrated in figure 4-4, the bullet penetrates the
outside wall of the cell, and the sticky, elastic
sealing material surrounds the bullet. As the bullet
passes through the cell wall into the cell, the
sealant springs together quickly and closes the
hole. Now some of the fuel in the tank comes in
contact with the sealant and makes it swell,
completing the seal. In this application, the
natural stickiness of rubber and the basic qualities
of rubber and petroleum seal the hole. This sealing
action reduces the fire hazard brought about by
leaking fuel. It keeps the aircraft’s fuel intact so
the aircraft may continue operating and return to
its base.
Figure 4-4.-Bullet sealing action.
from flowing through to the exterior of the fuel
cell (fig. 4-4.)
The mechanical reaction results because
rubber, both natural and synthetic, will “give”
under the shock of impact. This will limit damage
to a small hole in the fuel cell. The fuel cell
materials will allow the projectile to enter or leave
the cell, and then the materials will return to their
original position. This mechanical reaction is
almost instantaneous.
The chemical reaction takes place as soon as
fuel vapors penetrate through the inner liner
material and reach the sealant. The sealant, upon
contact with fuel vapors, will extend or swell to
several times its normal size. This effectively closes
the rupture and prevents the fuel from escaping.
The sealant is made from natural gum rubber.
The retainer material is the next material used
in fuel cell construction. The purpose of the
retainer is to provide strength and support. It also
increases the efficiency of the mechanical action
by returning the fuel cell to its original shape when
punctured. It is made of cotton or nylon cord
fabric impregnated with Buna N rubber.
The most commonly used types of self-sealing
fuel cells are the standard construction type and
the type that uses a bladder along with the selfsealing cell. Of the two, the standard construction
cell is used the most. It is a semiflexible cell, made
up of numerous plies of material.
The combination bladder and self-sealing cell
is made up of two parts. One part is a bladdertype cell, and the other part is identical to the
standard construction cell. It is designed to selfseal holes or damage in the bottom and the lower
portions of the side areas. The bladder part of
the cell (nonself-sealing) is usually restricted to the
upper portion. This type of cell is also semiflexible.
SELF-SEALING CELL (STANDARD CONSTRUCTION).— There are four primary layers
of materials used in the construction of a selfsealing cell. These layers are the inner liner, nylon
fuel barrier, sealant, and retainer. All self-sealing
fuel cells now in service contain these four primary
layers of materials. If additional plies are used in
the construction of the cell, they will be related
to one of the primary plies.
The inner liner material is the material used
inside the cell. It is constructed of Buna N
synthetic rubber. Its purpose is to contain the fuel
and prevent it from coming in contact with the
sealant. This will prevent premature swelling or
deterioration of the sealant.
Buna rubber is an artificial substitute for crude
or natural rubber. It is produced from butadiene
and sodium and is made in two types, Buna S and
Buns N. The Buna S is the most common type
of synthetic rubber. It is unsuitable for use as
inner liner material in fuel cells. It causes the
petroleum fuels used in aircraft to swell and
eventually dissolve. The Buna N is not affected
by petroleum fuels, making it ideal for this
application. However, the Buna N is slightly
porous, making it necessary to use a nylon barrier
to prevent the fuel from contacting the sealant.
The nylon fuel barrier is an unbroken film of
nylon. The purpose of the nylon fuel barrier is
to prevent the fuel from diffusing farther into the
cell. The nylon is brushed, swabbed, or sprayed
in three or four hot coats to the outer surface of
the inner liner during construction.
The sealant material is the next material used
in fuel cell construction. It remains dormant in
the fuel cell until the cell is ruptured or penetrated
by a projectile. It is the function of the sealant
to seal the ruptured area. This will keep the fuel
CONSTRUCTION).— One variation from the
standard construction, self-sealing fuel cell
previously discussed is shown in figure 4-5. It has
four primary layers—an inner liner, a nylon fuel
barrier, two sealant plies, and three retainer plies.
The cords in the first retainer ply run
lengthwise of the cell. The cords in the second
retainer run at a 45-degree angle to the first. The
cords in the third retainer run at a 90-degree
angle to the second. The outside is coated with
Figure 4-5.-Self-sealing fuel cell (standard construction).
Buna-Vinylite lacquer to protect the cell from
spilled fuel and weathering.
Baffles and internal bulkheads are used inside
the cell to help retain the shape of the cell and
prevent sloshing of the fuel. They are constructed
of square woven fabric impregnated with Buna N
Flapper valves are fitted to some baffles to
control the direction of fuel flow between
compartments or interconnecting cells. They are
constructed of Micarta, Bakelite, or aluminum.
These plies, baffles, internal bulkheads, and
flapper valves with the necessary fittings and
combinations make up a typical self-sealing fuel
Figure 4-6.-Bladder cell construction.
The nylon barrier consists of three to four
coats of nylon applied hot by brush, swab, or
spray. The purpose of the nylon barrier is to keep
fuel from diffusing through the cell wall.
The retainer consists of Buna N coated squarewoven fabric (cotton or nylon) or cord fabric. The
purpose of the retainer ply or plies is to lend
strength to the fuel cell and provide protection
for the nylon fuel barrier.
Bladder-Type Fuel Cells
A nonself-sealing fuel cell is commonly called
a bladder-type cell. It is a fuel container that does
not self-seal holes or punctures. The advantage
of using a bladder fuel cell results from the saving
in weight. Some of the other advantages are the
simplicity of repair techniques and the reduced
procurement costs over self-sealing fuel cells.
Bladder-type cells are usually made of very
thin material to give minimum possible weight.
They require 100-percent support from a smooth
cavity. The cell is made slightly larger than the
cavity of the aircraft for better weight and
distribution throughout the aircraft’s fuel cavity
The thinner wall construction increases the
fuel capacity over the self-sealing cells, thus
increasing the range of the aircraft. Many of our
aircraft that were formerly equipped with selfsealing cells have been changed to bladder-type
There are two types of bladder fuel cells—
rubber type and nylon type.
NYLON-TYPE BLADDER CELL (PLIOCEL).— Nylon bladder cells differ in construction
and material from the Buna N rubber cells. This
type of cell may be identified by the trade name
“Pliocel” stenciled on the outside of the cell. The
Pliocel construction consists of two layers of
nylon woven fabric laminated with three layers
of transparent nylon film.
The repair of this type of cell must be
accomplished by entirely different methods and
with different materials. The adhesive and Buna N
rubber used to repair the rubber-type bladder cell
cannot be used on the nylon-type cell.
rubber-type bladder cells are made in the same
manner as self-sealing cells. They have a liner,
nylon barrier, and a ret airier ply. The sealant
layers are omitted. All three plies are placed on
the building form as one material in this order—
liner, barrier, and retainer. Figure 4-6 illustrates
this type construction.
The inner liner may consist of Buna N rubber,
Buna N coated square-woven fabric (cotton or
nylon), or Buna N coated cord fabric. The
purpose of the inner liner is to contain the fuel
and provide protection for the nylon barrier.
External fuel systems increase range or mission
by providing additional fuel for increased range
or tanking. The external fuel system consists of
the fuel (drop) tanks, a transfer system and
jettison system.
Drop Tanks
The 150- and 300-U.S. gallon Aero 1C and the
300-U.S. gallon Aero 1 D external fuel tanks
are droppable, streamlined, metal containers
(fig. 4-7). These are carried under the wing to
Figure 4-7.-External fuel tank (with cutway view).
tank is equipped with a pressure-fueling float
switch and an air pressure and vent shutoff valve.
The pressure-fueling float switch is a floatoperated device that accomplishes the shutoff of
fuel flow when the external fuel tank is filled to
capacity. This is an electrical connector provided
on the top of the tank for connecting the float
supplement the internal fuel supply for extended
Threaded bosses and threaded inserts are
provided on the top of the tank to accommodate
the installation of adapter fittings. These are used
to connect the fuel tank to the aircraft fuel system
and the fuel tank air pressurization system. The
switch electrical wiring to the aircraft wiring. The
air pressure and vent shutoff valve vents the tank
to the atmosphere during the pressure fueling
procedure. However, the valve is also used with
the external fuel tank air pressurization system.
This valve uses engine bleed air as a means of
pressurizing the tank and forcing fuel into the
wing tank, or tanker store. A gravity filler port
is provided to accomplish gravity fueling when
pressure fueling equipment is not available.
Figure 4-8.-Ejector pumps. (A) Dual seat; (B) single seat; (C) wing transfer.
Auxiliary fuel pumps or booster pumps are
required in every pressure feed system. They are
needed to supply fuel pressure for starting the
engine and to supply fuel to the primer system.
They are also used as an emergency pump in case
of failure of the engine-driven unit.
The submerged boost pump is essentially an
integral unit composed of a centrifugal pump and
an electric motor. A screen is provided to protect
the pump from foreign matter. A submerged
boost pump is shown in figure 4-9.
External Fuel Transfer
External fuel tank pressurization and transfer
is accomplished with regulated engine bleed air.
An external tank pressure regulator maintains 15
to 18 psi air pressure to each of the external tanks.
Once the tank is pressurized, fuel then transfers
through the refuel/transfer shutoff valve into
the refueling manifold. External fuel is then
transferred to any of the fuel tanks that will
accept the fuel. The refuel/transfer valve will close
automatically when the external tank is empty.
The external tank air pressure regulator closes
when there is weight on the wheels or when the
in-flight refueling probe is extended. This will
prevent the tanks from being pressurized while the
aircraft is on the ground, during an arrested
landing, or during in-flight refueling.
Strainers.— Strainers are installed in the tank
outlets and frequently in the tank filler necks.
These strainers are of fairly coarse mesh and
prevent only the larger particles from entering the
fuel system. Other strainers are provided in the
fuel inlets and in the fuel lines themselves. The
latter are fine-mesh strainers.
External Fuel Tank Jettison
FUEL QUANTITY INDICATORS.— Quantity-indicating units will vary. A fuel counter or
indicator, mounted on the instrument panel, is
electrically connected to a flowmeter installed in
the fuel line to the engine. The fuel counter is
similar in appearance to an automobile
speedometer. When the aircraft is serviced with
fuel, the counter is automatically set to the total
number of pounds of fuel in all tanks. As fuel
passes through the measuring element of the
flowmeter, it sends electrical impulses to the fuel
counter. These impulses actuate the fuel counter
mechanism in such a way that the number of
The external fuel tanks can be selectively
jettisoned or all jettisoned at one time, such as
during an emergency situation. The external
tank to pylon fuel/air coupling valves will
automatically close the fuel transfer and air
pressurization tubes once the tanks are jettisoned.
Fuel Tank Components
Common fuel tank parts include pumps,
strainers, fuel quantity indicators, valves to
control fuel level or routing, and vents and drains.
These parts provide capabilities for fueling,
defueling, and fuel system management.
PUMPS.— The airframe fuel system uses
transfer pumps and boost pumps to deliver a
continuous supply of fuel to the engine(s) under
all operating conditions.
Transfer Pumps.— Fuel transfer pumps are
installed in the fuel system to pump fuel from the
various tanks of the aircraft to the main or sump
tank. There are several different types of transfer
pumps; common ones are electrically driven or
an ejector-type motive-flow pump. See figure 4-8.
Since the type of pump may differ from one aircraft model to another, the applicable maintenance instruction manual should be consulted
for proper identification and maintenance.
Boost Pumps.— All Navy fixed-wing aircraft
use pressure feed fuel systems. The basic source
for this pressure is the engine-driven pump.
Figure 4-9.-Submerged boost pump.
pounds passing to the engine is subtracted
from the original reading. Thus, the fuel counter
continually shows the total quantity of fuel (in
pounds) remaining in the aircraft. However, there
are certain conditions that cause the fuel counter
indication to be inaccurate. Any fuel remaining
in the droppable tanks when they are jettisoned
is indicated on the fuel counter as fuel still
available for use. Any fuel that leaks from a tank
or a fuel line upstream of the flowmeter is not
counted. Any fuel supplied to the engine by the
emergency pump is not counted.
Some continuous-flow fuel systems have a fuel
quantity gauge for each tank or group of interconnected tanks. If the system has a main tank
with auxiliary tanks feeding into it, a fuel quantity
gauge is normally for the main tank. In this type
of system, the pilot relies on the indication of the
fuel counter (flowmeter). All fuel in the auxiliary
tanks is transferred to the main tank and fed to
the engine. When all fuel except that in the main
tank has been consumed, the fuel quantity gauge
provides a more reliable indication of the fuel still
available. The accuracy of its indication is not
affected by the conditions listed in the preceding
paragraph; that is, leakage and emergency system
The fuel quantity gauge normally used in aircraft is an electronic (capacitor) type for measuring fuel in aircraft in pounds. Normally, the
capacitor-type fuel gauge is used without a
flowmeter, although most engines have provisions
for installing one if it is required,
A low-level switch is incorporated in the fuellevel transmitter. This switch turns on an indicator
light in the cockpit when the fuel in the tank drops
to a specific low level. This signal informs the pilot
that the fuel supply is almost exhausted.
Figure 4-10.-Motor-operated shutoff valve (gate valve).
Fuel Level Control Valves.— Fuel level
control valves control fuel levels in a tank during
ground fueling or fuel transfer to the main tank.
There is one fuel level control valve for each tank,
auxiliary tank, or group of interconnected tanks.
When used for fuel transfer, the valves are located
at different levels in the main tank. Fuel is then
transferred from the auxiliary tanks in the order
designed by the manufacturer. During normal
operation of the fuel system, the boost pumps for
all the tanks are turned on before the engine is
started. Each auxiliary tank boost pump continues
to operate until the tank is emptied; then the fuel
pressure warning light comes on and the boost
pump is turned off by the pilot. Thus, fuel is
delivered under boost pump pressure to each fuel
level control valve. The fuel then remains in the
tank or group of tanks to which it is connected.
In the sectional views of the valve in
figure 4-11, note how the float rises and lowers
with the fuel level. When the fuel level in the main
tank is high, the float is raised. This closes the
pilot valve and lifts the ball check from its seat.
Fuel, under boost pump pressure, then passes
through the main valve stem into the valve body.
Note how the fuel pressure exerted against the
bottom surface of the synthetic rubber diaphragm
holds the main valve closed. This prevents fuel
from entering the main tank from the transfer
When the fuel level in the main tank drops,
the float moves downward. Note in figure 4-11
how this action allows the ball check to seat in
used to regulate and control the flow of fuel in
the airframe and engine fuel systems. Some of
these valves are discussed in the following
Shutoff Valves.— Shutoff valves are twoposition (open and closed) valves. The manually
operated type is installed to shut off the fuel while
a unit in the system is being removed or replaced.
Electrically operated shutoff valves control flow
during fuel transfer and when fuel is being bypassed because of a defective or damaged unit.
Figure 4-10 shows a motor-operated shutoff valve,
commonly referred to as a gate valve.
Figure 4-11.-Fuel level control valve.
the main valve stem. It then shuts off the fuel
pressure in the bottom side of the diaphragm. The
pilot valve opens and permits fuel to drain from
the main valve body. As the pressure on the under
surface of the diaphragm is relieved, the main
valve opens to admit fuel from the auxiliary tank.
is a one-direction valve, It is important that the
valve be installed so that the arrow points in the
desired direction of flow.
In some turbojet engine fuel systems, there is
a check valve between the fuel control and the fuel
dump valve. The check valve remains closed until
a certain pressure is reached in the fuel line. A
bypass to the top of the dump valve transmits this
bypass pressure. Then, upon engine starting, the
controlled fuel pressure builds up. The dump
valve is then actuated to close the drain port and
open the flow into the fuel manifolds. The check
valve remains fully open during engine operation.
Check Valves.— Check valves are installed in
the fuel system wherever fuel flow in one direction is required. Fuel pressure in the direction of
flow-indicated by an arrow on the valve-forces
the valve open against spring pressure. Spring
force and reversal of fuel flow close the valve. This
Selector Valves.— In the continuous-flow
system, selector valves are not used for tank-toengine selection during normal operation.
However, in many installations there are selector
valves to enable the pilot or mechanic to control
the fuel flow for special purposes. This includes
fuel integrity checks, shutting off fuel to the
engine, bypassing fuel components to allow
manual operations in emergency conditions, and
cross-feeding fuel to different tanks or engines to
prevent an unbalanced fuel load.
lines between the various tanks and between the
tanks and the engine-driven pump are of the
conventional type. They consist of metal tubing
or flexible hose. There are drain cocks at low
points in the lines so that any water that collects
at these locations may be drained. A quickdisconnect fitting is often installed in the main fuel
line to the engine. This fitting permits quick
disconnection of the main fuel line when an engine
change is performed.
The line connecting the various fuel system
units installed on the engine are made of either
metal tubing or flexible hose. Since these lines
and fittings must withstand the high pressures
encountered on the discharge side of the enginedriven fuel pump, special types are used.
Hose Assemblies.— The
lightweight engine hose assembly is designed for
continuous operating temperatures of –40°
to +300°F. The inner tube is seamless and is of
a specially formulated synthetic compound. The
reinforcement and cover are of stainless steel wire
braid and consist of a partial inner braid and a
full-coverage outer braid. This hose can be
identified by the bright wire braid outer cover with
red markings. These markings are repeated
6 inches apart.
This hose is designed for aircraft power plant
and airframe fuel and oil lines. It is widely used
in jet engines. It is flexible, lightweight, and has
the ability to withstand high operating
temperatures where maximum fire resistance is a
prime consideration. This hose may be used in
submerged applications.
The fitting on this type of hose uses a lip-seal
principle, instead of compression, to effect a fluid
seal. This lip-seal is formed during assembly by
a sharp knifelike spur, which cuts an annular flap
in the hose inner tube. Fitting retention against
blowoff is affected by the cutting action of the
spur. This separates the wire braid, which is then
gripped between the nipple and the socket. These
fittings must be marked with a painted stripe to
detect hose pushout after assembly or proof test.
Rigid Tubing.— The majority of rigid tubing
used in naval aircraft is manufactured from
aluminum. However, exposed lines and lines
subject to abrasion or intense heat are made of
stainless steel. Therefore, you will be concerned
more with stainless steel lines. Whenever an
engine fuel line requires replacement, the normal
procedure is to obtain from supply a preformed
line with fittings attached. If a line must be
manufactured locally and installed on an engine
or component, the original line must be duplicated
as exact as possible.
Figures 4-12 and 4-13 show a few of the
correct and incorrect methods of installing metal
tubing and flexible hose.
Figure 4-12.-Correct and incorrect methods of installing tubing.
Figure 4-13.-Correct and incorrect installation of flexible
Figure 4-14.-Fuel line fittings. (A) bulkhead fitting;
(B) universal fitting; (C) universal fitting with O-ring seal
and seal ring.
Bulkhead Fitting. — Bulkhead fittings must be
properly installed. To ensure proper installation
of the fitting shown in figure 4-14, view A, the
mechanic must check to see that the bulkhead has
the required thickness for which the fitting was
the seal ring should be coated sparingly with
petrolatum or hydraulic fluid. Work the seal ring,
with the smooth (hair) side toward the O-ring seal,
into the counterbores of the nut. You then turn
the nut down until the O-ring seal is pushed firmly
against the lower threaded section of the fitting.
Install the fitting into the boss. You then keep the
nut turning with the fitting until the O-ring
contacts the boss. This point can be determined
by a sudden increase in torque. With the fitting
in this position, put a wrench on the nut to
prevent its turning; then turn the fitting in 1 1/2
turns. Then position the fitting by turning in not
more than one additional turn. Hold the fitting
and turn the nut down tight against the boss.
Slight extrusion of the ring is not considered
detrimental. See figure 4-14, view C.
Fitting With O-Ring Seal.— The nut should be
assembled on the fitting end until the washer face
of the nut lines up with the upper corner of the
seal groove. The O-ring seal should be lubricated
sparingly with petrolatum and placed on the
fitting groove so it contacts the nut. Then screw
the fitting (and nut simultaneously) into the boss
until the seal contacts the boss chamfer and the
nut contacts the boss. Before tightening the
locknut, position the fitting direction by turning
it three-fourths turn or turning it out one-fourth
turn. Assemble the fluid line to the fitting end.
Holding the fitting stationary in the selected
position, tighten the locknut. See figure 4-14,
view B.
FUEL DRAINS.— So the moisture content
can be checked and moisture drained from
the fuel system, the drain valve(s) is/are installed in the low point (or points) in the system
(or units).
Fitting With O-Ring Seal and Seal Ring.—
The threads of the fitting, the O-ring seal, and
Figure 4-15, views A through G, shows seven
different types of fuel drain valves used on
The valve shown in view A is usually located
in the boost pump or in the low-point drain. This
fitting needs to be pushed up and held to have
it in the OPEN position. To close the valve, you
should release the plunger.
The valve shown in view B is usually found
in the main fuel filter drain. To open this type
of drain, you should rotate the bar counterclockwise to lock it in the OPEN position. To close the
drain, rotate the bar in the clockwise direction.
The valve shown in view C is usually located
in the inboard or outboard compartment lowpoint drain. To open and lock it in the open
position, insert a screwdriver in the slot and turn
it clockwise (about 90°). To close this valve, turn
the screwdriver counterclockwise.
The fuel drain valve shown in view D can be
opened by inserting a screwdriver in the slot,
pushing in, and holding it, which will allow fuel
to flow. It can be closed by releasing the
The fuel drain shown in view E is for the aft
boost pump drain. It can be opened and locked
in the OPEN position by rotating it in the
counterclockwise direction. Rotating it in the
clockwise direction will close the valve.
The valve shown in view F is usually found
in the low-point drain and in the main vent line
Figure 4-15.-Fuel drain valves.
of the low-point drain. The valve is automatically
actuated open at 1.0 psi minimum and closes at
3.0 psi maximum.
The valve shown in view G is usually found
in the low-point drain, forward sump cell, and
is opened by pushing and holding. It is closed by
releasing the plunger.
Fuel from the airframe fuel system is supplied
to the engine-driven fuel pump through the engine
fuel supply hose. The engine fuel supply hose is
the last link between the airframe fuel system and
the engine fuel system. Fuel from the enginedriven fuel pump is directed to the fuel control.
Then it is regulated and distributed to the
combustion chambers. Components of the engine
fuel system are discussed in the following
paragraphs, along with operation.
The JFC 25-3 hydromechanical fuel control,
shown in figure 4-16, is a lightweight, highcapacity, fuel-flow-metering unit. It is designed
to permit selection of a desired engine jet thrust
level. It also provides automatic compensation
through the full range of thrust for the ambient
operating conditions encountered during flight.
Engine thrust during ground operation and under
various flight conditions is controlled by a single
power lever. It also regulates fuel for engine
starting and shutdown. The variables sensed by
the fuel control are power lever angle, burner
high-pressure compressor speed
and compressor inlet temperature
using these variables, the fuel control accurately
governs the engine’s steady-state. It is selected
through a speed-governing system of the proportional or droop type. The fuel control also uses
these same variables to control fuel flow for
acceleration and deceleration.
The fuel control consists of a fuel-metering
system and a computing system. The metering
system regulates fuel supplied to the engine by the
engine-driven fuel pump to provide the engine
thrust demanded by the pilot. Fuel regulation is
also controlled by engine operating limitations,
as sensed and scheduled by the fuel control
computing system. The computing system senses
and combines various operational parameters to
govern the output of the metering system of the
fuel control under all engine operating conditions.
High-pressure fuel is supplied to the control
inlet from the engine-driven pump. At the inlet
Figure 4-16.-Fuel control (JFC 25-3).
of the control, the fuel is filtered by a coarse
(80-mesh) screen and a fine (40-micron) screen.
The coarse screen protects the metering system
from large particles of fuel contaminants. If this
screen becomes clogged, a filter relief valve will
open, permitting continued operation with
unstrained fuel. The fine screen protects the
computing system against solid contaminants.
This screen is self-cleaning. It traps particles by
removing the high-velocity of the fuel flowing past
the screen into the metering section.
Next, the fuel encounters the pressure-regulating valve, which is designed to maintain a
constant pressure differential across the throttle
valve. All high-pressure fuel in excess of that
required to maintain this pressure differential is
bypassed to the pump interstage by the pressureregulating valve. This valve is servo controlled.
The actual pressure drops across the throttle valve
orifice and is compared, by the sensor, with
a selected pressure drop, and any error is
hydraulically amplified. The amplified error
positions the pressure-regulating-valve spring,
altering the force balance of this valve so that
sufficient high-pressure fuel is bypassed to
maintain the selected pressure drop. The pressureregulating-valve sensor also incorporates a
bimetallic disc to compensate for any variation
in the specific gravity of the fuel, which results
from fuel temperature change.
The high-pressure fuel, as regulated by the
pressure-regulating valve, then passes through the
throttle valve. This valve consists of a contoured
plunger that is positioned by the computing system of the control within a sharp-edged orifice. By
virtue of the constant pressure drop maintained
across the throttle valve, fuel flow is a function
of the plunger position. An adjustable stop limits
the motion of this plunger in the decrease fuel
direction to permit minimum fuel flow.
The final part to act upon the metered flow
prior to its exit from the control is the minimum
pressure and shutoff valve. This valve is designed
to shut off the flow of metered fuel to the engine
when the power lever is in the OFF position. This
causes the power-lever-operated sequencing valve
to transmit a high-pressure signal to the spring
side of the shutoff valve. This forces the latter
against the seat, thus shutting off the flow of fuel
to the engine. When the power lever is moved out
of the OFF position, the high-pressure signal is
replaced by pump interstage pressure. Then
metered fuel pressure is increased sufficiently to
overcome the spring force, the valve opens, and
fuel flows to the engine. Thereafter, the valve will
provide a minimum operating pressure within the
fuel control. This ensures that adequate pressure
is always available for operation of the servos and
valves at minimum flow conditions.
The power-lever-operated sequencing valve
also incorporates a windmill bypass feature, which
functions when the shutoff valve is closed. This
feature bleeds throttle valve discharge flow to the
fuel pump interstage to increase the throttle valve
pressure drop and opens the pressure-regulating
valve. Damage to the fuel pump from excessive
pressure is thus prevented during engine windmilling. The sequencing valve functions in both
the normal and manual operating systems.
The following designators are used in the
description of the computing system of the
JFC 25-3 fuel control. These designators should
be referred to during study of the fuel control.
High-pressure compressor rotor
speed (RPM)
Compressor inlet temperature
Burner can pressure
Ratio of metered fuel flow
to burner can pressure
The computing system positions the throttle
valve to control steady-state engine speed,
acceleration, and deceleration. This is accom(the ratio of
plished by using the ratio
metered fuel flow to engine burner pressure) as
a control parameter. Throttle valve positioning
of this parameter is achieved through a multiplying system whereby the
signal is used for
acceleration or deceleration. The steady-state
speed control is multiplied by a signal proporto provide the required fuel flow.
tional to
is sensed in the following manner: A motor
bellows is internally exposed to
and the
resulting force is increased by the force of an
evacuated bellows of equal size. It is directly
connected to the motor bellows. The net force,
absolute burner pressure, is transmitted through
a lever system to a set of rollers having a
These rollers
position proportional to
ride between the bellows-actuated lever and a
multiplying lever. The force proportional to
is thus transmitted through the rollers to the
multiplying lever. Any change in the roller
or the
signals upsets the
equilibrium of this lever. This changes the
position of a flapper-type servo valve, which is
supplied with regulated high-pressure fuel through
a fried bleed orifice. The resulting change in servo
pressure between the two orifices is controlled by
the position of a piston attached to the throttle
valve plunger. The motion of this piston
compresses or relaxes a spring that will return the
multiplying lever to its equilibrium position. An
adjustable minimum-ratio stop on the
signal controls engine deceleration. This arrangement provides a linear relationship between
which results in blowoutdecreasing
free decelerations.
An adjustable maximum-ratio stop on the
signal controls engine acceleration. This
stop is positioned by an acceleration-limiting cam.
It is rotated by a speed-sensing servo system and
translated by a compressor inlet temperature
sensing servo system. The cam is so contoured as
versus engine
to define a schedule of
speed for each value of
that will permit engine
accelerations. This avoids engine overtemperature
and surge limits without compromising engine
acceleration time.
A burner pressure limiter incorporated in the
fuel control senses burner pressure with respect
to ambient pressure. When this differential
exceeds a preset maximum, the pressure signal to
the burner pressure motor bellows. This reduces
bleeding through the limiter valve to ambient
pressure. This causes a limitation on fuel flow,
which prevents burner pressure from exceeding
a maximum, safe value.
A flyweight-type, engine-driven, speed-sensing
governor controls movement of the speed servo
piston through a pilot valve. When
changes, the flyweight force varies and the pilot
valve is positioned to meter either low- or
high-pressure fuel to the speed servo piston. The
motion of the piston repositions the pilot valve
until the speed-sensing system returns to
equilibrium. The piston incorporates a rack that
meshes with a gear segment on the threedimensional acceleration cam to provide the speed
signal for acceleration limiting. This piston
position is also used to indicate actual engine
speed, and it is connected by a droop lever to a
droop cam.
The temperature-sensing bellows and servo
assembly are connected through a lever and yoke
assembly to the acceleration-limiting cam. The
position of this servo piston is indicative of
and is used
compressor inlet temperature
to translate the acceleration cam. It integrates the
temperature and speed signals. The position of
the speed-set cam is also translated by the servo
piston by means of a cross-link to the acceleration
cam. Engine steady-state condition is a function
of high compressor
speed, compressor inlet
burner pressure
power lever position.
In the event that the primary control system
malfunctions, the manual system may be engaged
by operating a switch in the cockpit. It then
energizes the manual transfer solenoid to close the
flapper valve. The flapper valve will remain in the
closed position because of residual magnetism,
regardless of whether or not the solenoid is
continuously energized. Servo action positions the
shuttle valve to direct pump discharge pressure
to the spring side of the manual and normal
systems transfer valve. This pressure, combined
with spring pressure, positions the valve to close
off the primary operating system and direct highpressure fuel to the manual system.
Fuel-pressurizing Valve
The fuel-pressurizing valve is usually required
on jet engines. It incorporates a duplex-type
fuel nozzles to divide the flow into primary and
secondary (main) manifolds. At the low fuel flows
required for starting and altitude idling, all the
fuel passes through the primary line. As the fuel
flow increases, the valve begins to open the main
line. At maximum flow, the main line is passing
about 90 percent of the fuel.
Fuel-pressurizing valves will usually, through
incorporation of spring-loaded inlet check valves,
trap fuel forward of the manifold, giving a
positive cutoff. This cutoff prevents fuel from
leaking into the manifold and through the fuel
nozzles. This eliminates afterfires and carbonization of the fuel nozzles. Carbonization occurs
when low combustion chamber temperatures
cause incomplete burning of the fuel.
An example of this arrangement is the fuelpressurizing and dump valve. This valve performs
two major functions, as indicated by its name.
During engine operation, it divides metered fuel
flow into two properly pressurized portions,
primary and secondary. During engine shutdown,
it provides a dump system that connects the fuel
manifolds to an overboard drain. The features of
the fuel-pressurizing and dump valve are shown
in figure 4-17.
Fuel valves in the engine fuel system aid in
starting, stopping, and as safety factors. Valves
may differ slightly from engine to engine, and they
may be called by different nomenclature, although
they perform identical functions. Some of these
valves and their functions are discussed in the
following paragraphs.
Figure 4-17.-Fuel-pressurizing and dump valve.
The fuel-pressurizing and dump valve is
connected to the fuel manifold. It is composed
of an inlet check valve, a 200-mesh fuel inlet
screen, a pressurizing or flow-dividing valve, and
a manifold dump or drain valve.
Flow Divider
A flow divider performs essentially the same
function as a pressurizing valve. It is used, as the
name implies, to divide flow to the duplex fuel
nozzles. It is not unusual for units performing the
same functions to be called different names on
different engines or by different manufacturers.
Drain Valves
Drain valves drain residual fuel from the
various parts of jet engines where accumulated
fuel is most likely to present operating problems
The chance of a fire hazard exits in a combustion
chamber if fuel accumulation occurs during
shutdown. Residual lead and gum deposits from
evaporated fuel cause problems in fuel manifolds
and fuel nozzles.
In some instances, the function of draining
fuel manifolds is accomplished by an individual
unit known as a drip or dump valve. This type
of valve may operate by pressure differential, or
it may be solenoid operated. See figure 4-18.
The combustion chamber drain valve drains
raw fuel that accumulates in the combustion
chamber. It drains after each shutdown when the
engine fire has gone out, and it drains fuel that
collects during a false start. The can type
combustion chambers drain fuel, by gravity, down
through the flame tubes or interconnector tubes
until it gathers in the lower chambers. It is fitted
with drain lines to the drain valve. In the basket
annular type combustion chamber, the fuel drains
through the airholes in the liner and collects in
a trap in the bottom of the chamber housing. A
Figure 4-18.-Solenoid-operated drip (dump) valve.
typical combustion chamber drain valve is shown
in figure 4-17.
When the fuel collects in the drain lines, the
drain valve allows the fuel to drain when pressure
in the combustion chamber manifold is reduced
to near atmospheric pressure. As shown in figure
4-17, the drain valve is spring-loaded in an open
position. It is closed as pressure within the
manifold and lines to the burners increases above
that of the spring tension trying to keep the valve
open. It is imperative this valve be in good working condition to drain accumulated fuel after each
shutdown. Otherwise, a HOT START during the
next starting attempt or an AFTERFIRE after
shutdown may occur.
time. It is very important that the fuel be evenly
distributed by the spray to prevent the formation
of any hot spots in the combustion chambers. It
is of particular value for this reason that the spray
be well centered in the flame area of the liners.
Fuel nozzle types vary between engines; mostly
fuel is sprayed into the combustion area under
pressure through small orifices in the nozzles. The
nozzles generally used are of the vaporizing orifice
type and include the simplex and the duplex
configurations. The duplex nozzle usually requires
a dual manifold and a pressurizing valve or flow
divider. This is to divide primary and secondary
(main) fuel flow, while the simplex nozzle requires
only a single manifold for proper fuel delivery.
Fuel Spray Nozzles and Fuel Manifolds
fuel nozzle was the first type of nozzle used in
turbojet engines, but it was replaced in most
installations with the duplex nozzle, which gives
better atomization at starting and idling speeds.
The simplex nozzle is still being used to a limited
degree. A simplex nozzle is shown in figure 4-19.
Each of the nozzles of the simplex type consists
of a nozzle tip, an insert, and a strainer made of
a fine-mesh screen and a support.
In jet engines, the fuel spray nozzles function
is to inject fuel into the combustion area in a
highly atomized, precisely patterned spray. It then
burns evenly and in the shortest possible space and
DUPLEX FUEL NOZZLE.— The duplex fuel
nozzle is the type nozzle most widely used in
present-day engines. Its use requires a flow
divider, which gives a desirable pattern of spray
for combustion over a wide range of operating
pressures. A nozzle of this type is shown in
figure 4-20.
Figure 4-19.-Simplex fuel nozzle.
Figure 4-20.-Duplex fuel nozzle.
The duplex nozzle may be represented in many
configurations, depending upon the type of
combustion chamber installation. Therefore, the
nozzle parts will vary between duplex nozzles of
various engines. Figure 4-21 shows a duplex
nozzle for use in a can-annular combustion
The primary fuel entry line (manifold) of the
duplex nozzle is smaller than the secondary entry
line. This feature permits fuel within the primary
line to reach a comparatively high degree of
pressure and atomization during starting and
altitude idling conditions. The secondary fuel
entry line also starts supplying fuel when engine
rpm raises fuel pressure to a predetermined
level—usually after engine rpm is stabilized after
a start.
At sufficient pump outlet pressure, the
pressurizing valve or flow divider allows fuel
to enter the main or secondary line. The
spray orifice will increase its spray angle because of the increased fuel flow and pressure.
Figure 4-20 shows the spray angle of a typical
duplex nozzle.
Valves covered in this section are designed to
control the flow of fuel in all aircraft fuel
systems. Construction and flow control are similar
for most valves used in modern aircraft. The
construction of the selector is basically a ported
body housing a rotor. Controlling these valves can
be manual or electrical. Graphite sealing discs are
so arranged on the rotor that the ports are sealed
or opened in sequence by rotation of the rotor.
The ends of the rotor bores in the body are closed
by top and bottom caps with O-rings. The rotor
stem extends through the top cap with an O-ring
seal to prevent leakage. This stem is rotated by
an electrical actuator assembly or by either a
handle or a yoke for manual actuation. Where
manual actuation is used, the top cap incorporates
a spring-loaded ball and a stop pin to index the
various rotor positions. Figure 4-22 shows a
selector valve.
The motor-operated gate valve provides a
means of controlling the flow of fuel to various
Figure 4-21.-Duplex nozzle spray pattern.
Figure 4-22.-Selector valve.
The single manifold of the simplex nozzle
does not have the above-mentioned feature
and must supply fuel under all operating
conditions. So duplex nozzles provide better
low-speed performance then simplex-type nozzles.
parts of the fuel system. It is designed as an openand-closed valve and is motor-operated. The gate
or sliding portion of the valve slides between
O-rings or other suitable sealing devices in the
body of the valve. On some models, an indicator
is attached to the gate to show the position of the
valve while installed in the system. Some of these
valves have a cable and drum between the motor
and valve mechanism to provide for manual override. This mechanism may be used if the electrical
motor is defective. Figure 4-23 shows a motoroperated gate valve with a manual-override
mechanism. The installation and rigging of motoroperated gate valves are similar to those of the
fuel selector valves. However, the motor-operated
gate valves that have no manual override require
no adjustment on installation.
The three most common types of filters in use
are the microfilter, the wafer screen filter, and the
plain screen mesh filter. The individual use of each
of these filters is dictated by the filtering
treatment required at a particular location.
Figure 4-24.-Aircraft fuel filter (microfilter).
cellulose material, frequently used in the construction of filter cartridges, removes foreign matter
measuring 10 to 25 microns. The minute openings
make this type of filter susceptible to clogging;
therefore, a bypass valve is a necessary safety
The microfilter, shown in figure 4-24, has the
greatest filtering action of any present-day filter,
and it is rated in microns. (A micron is a
thousandth part of 1 millimeter.) The porous
Since the microfilter does such a thorough job
of removing foreign matter, it is especially
valuable between the fuel tank and engine. The
cellulose material also absorbs water, preventing
it from passing through the pumps. If water
does seep through the filter-and this happens
occasionally when filter elements become
saturated with water—the water can and does
quickly cause damage to the working elements of
the fuel pump and control units. These elements
depend solely on the service fuel for their
lubrication. To reduce water damage to pumps
and control units, periodic servicing and replacement of filter elements are imperative.
The most widely used filters are the 200-mesh
and the 35-micron filters. They are used in fuel
pumps, fuel controls, and between the fuel pump
and fuel control where removal of microscopic
particles is needed. These filters, usually made of
a fine-mesh steel wire, are a series of layers of
wire. This type of filter replaces the wafer screen
described in-the next paragraph.
Figure 4-23.-Motor-operated gate valve and override
Wafer Screen Filter
The wafer screen filter, shown in figure 4-25,
has a replacement element made of layers of
screen discs of bronze, brass, and steel. This type
of filter can remove minute particles. It also has
the strength to withstand high pressure.
Plain Screen Mesh Filter
The plain screen mesh filter is the most
common type. It has long been used in internalcombustion engines of all types for fuel and oil
strainers. In present-day turbojet engines, it is
used in units where filtering action is not so
critical, such as in fuel lines before the highpressure pump filters. The mesh size of this type
of filter varies greatly according to the purpose
for which it is used.
operation of the aircraft engine. The engine-driven
fuel pumps must be capable of delivering the
maximum needed flow at high pressure to obtain
satisfactory nozzle spray and accurate fuel
Fuel pumps for engines are generally positive
displacement gear, piston, or rotary vane types.
The term positive displacement means that the
pump will supply a fixed quantity of fuel to the
These pump types may be divided into two
groups—constant displacement and variable
displacement. Their use depends on the fuel
control system used to regulate the flow of fuel
to the fuel controls.
Gear-Type Pumps
Gear-type pumps have straight-line flow
characteristics. However, fuel requirements vary
with flight or ambient air conditions. Hence, a
pump of adequate capacity at all engine operating
conditions will have excess capacity over most of
Engine-driven fuel pumps deliver a continuous
supply of fuel at the proper pressure during
Figure 4-25.-Wafer screen filter.
the range of operation. This characteristic requires
the use of a pressure relief valve for disposing of
excess fuel. A constant-displacement gear-type
pump is illustrated in figure 4-26.
Variable-Displacement Pump
The variable-displacement pump system differs from the constant-displacement pump system.
Pump displacement is changed to meet varying
fuel flow requirements; that is, the amount of fuel
that is discharged from the pump can be made
to vary at any one speed. With a pump 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,
connected in parallel. Either pump can carry the
load if the other fails during normal parallel
operations. At times, one pump is not enough to
meet power requirements. Pump duplication
increases safety in operation, especially in takeoffs
and landings.
The positive-displacement, variable-stroke
type of pump incorporates a rotor, a piston, a
maximum speed governor, and a relief valve
Figure 4-26.-Cutaway view of a dual-element gear-type pump.
Figure 4-27.-Variable-stroke fuel pump.
mechanism. A variable-stroke pump is shown in
figure 4-27.
The most important consideration in working
with any fuel system maintenance task is the
safety of personnel. Aircraft fuels are extremely
hazardous because of the explosive and toxic
dangers that are always present. The health
hazards associated with aviation fuels (breathing
of vapors, spillage on skin or in the eyes, or
swallowing) must be avoided. It is not possible
to describe all the potential problems or dangers
that may arise in the performance of any type of
fuel system maintenance. As an AD, it is your
responsibility to be thoroughly aware of all the
safety practices and procedures that must be
strictly followed.
Fuel vapors are very harmful when they are
inhaled. It takes only a very small percentage of
these vapors to cause very serious effects on
personnel. Fuel vapors are heavier than air and
will collect in the lower areas of the fuel tank/cell.
Unless these vapors are removed by the use of
forced-air ventilation, they can present a hazard
for an indefinite period. Personnel should avoid
the inhalation of these vapors, and always be alert
An engine-driven rotary-vane type of pump
and a diagram showing the operation of the unit
are shown in figure 4-28.
The engine-driven fuel pump is turned by a
gear train in the accessory section of the engine.
Constant pressure is maintained by a springloaded pressure relief valve. Figure 4-28 shows the
pressure relief valve in operation, bypassing
excess fuel back to the inlet side of the pump.
Fuel is bypassed before the engine is started,
when the engine-driven fuel pump is not turning.
An auxiliary fuel booster pump delivers fuel under
pressure. Fuel pumped by the booster pump will
pass through the stationary engine-driven pump;
it is necessary to incorporate a bypass valve in the
engine-driven pump. Both the fuel pressure relief
valve and the bypass valve may be contained in
the same mechanism.
Refer to Fluid Power, NAVEDTRA 12964,
for a detailed description of the principles of
operation of the various types of pumps.
Figure 4-28.-Engine-driven, rotary-vane type of pump.
to recognize the first signs of the toxic effect of
breathing these vapors. The symptoms of inhalation include nausea, dizziness, and headaches. If
a person should experience these symptoms during
fuel system maintenance, immediately stop and
move the individual to a source of fresh air. If
the individual appears to be completely overcome
by the vapors, get prompt medical attention.
When working with any type of aviation fuel,
personnel should always avoid prolonged contact
with the fuel. If a person’s clothing becomes
saturated, they should remove them as soon as
possible and wash off the affected areas with soap
and water. It is essential to know the location of
approved eyewash stations and how they are used.
In modern aircraft, the fuel systems are
designed to operate satisfactorily under all
conditions, such as acceleration and deceleration,
temperature, pressure and flight attitudes.
However, no matter how good the design, the fuel
system will not function as designed if it is not
maintained properly. A significant number of fuel
leaks can be attributed to incorrect maintenance
procedures used in installing fuel tanks/cells,
components, lines, and fittings. By referring to
the applicable aircraft maintenance manual and
learning the general procedures discussed in this
section, you will have little difficulty in locating
the source of an aircraft fuel leak.
Leak source analysis is the process of using
the aircraft maintenance manual, fuel system
schematic diagrams, installation diagrams, and
troubleshooting charts. The most common
method of analysis is the methodical process of
elimination to isolate the source of a fuel leak.
In addition, you should first screen the aircraft
discrepancy book (ADB) to possibly save many
man-hours looking for a leak. The review of a
prior fuel system discrepancy may reveal that
spilled fuel was not properly cleaned or components were improperly installed. Never assume
that the first leak you find is the only leak in the
system. Completely check and test the entire fuel
system as directed by the applicable maintenance
Severe leaks in the tank/cell drain system are
caused by a rupture, loose interconnecting fittings,
or cut or distorted O-rings. These leaks can usually
be detected immediately after refueling the
tank/cell. Dripping leaks are usually found at fuel
system plumbing connections. Leaks are caused
by undertorquing or overtorquing lines, hoses, or
fittings. Never assume that the leaks can usually
be detected by operating the fuel transfer
pump/boost pump to pressurize the fuel system.
Intermittent leaks are most often caused by loose
cell fittings or connections. Fuel quantity probes
that are mounted on the high side of the tank/cell
usually leak when the aircraft is in a climb or
descent. In some cases, servicing the fuel tank/cell
to capacity may aid in locating these types of
Fuel Dye to Locate Leaks
The use of colored dye to detect hidden fuel
leaks is a practical means you can use in fuel
system leak source analysis. The dyed fuel will
leave a stain that can be traced back to the
source of the fuel leak. (The use of dyed fuel is
particularly useful in checking for leakage. This
is especially true near the engine’s hot section
where high temperatures prevent the fuel from
leaving a wet spot.) When using a dye to aid in
the troubleshooting of fuel leaks, a logbook entry
in the miscellaneous history section of the aircraft
logbook should be made. The fuel color, resulting
from the use of dye, can be disregarded in fuel
sample analysis. Additionally, a similar entry
should be made for aircraft serviced with dyed
fuel. You should always select a dye color that
will provide the highest visibility in the area where
the leaking fuel is suspected. The use of 2 ounces
of dye for each 100 gallons of fuel in the cell or
tank is required. The appropriate information for
ordering the dye can be found in Appendix A of
NA 01-1A-35. The addition of unmixed dye to
empty fuel systems should always be avoided
because it can cause deterioration of the cell
lining. The dye should always be added to the fuel,
rather than fuel added to the dye. For information and correct procedures for the use of dyes
in fuel system leak detection, refer to the proper
maintenance manual.
A liquid red or yellow dye that can be added
directly to the aircraft fuel tank is available.
Before the dye is used to determine the source of
fuel leaks, any visible fuel leaks in the tanks and
plumbing must be eliminated. If you must use dye
to locate a hidden leak, test only one questionable
tank at a time. Add the liquid dye to the suspected
tank when it is one-third full, using a full 2-ounce
can of dye for each 100 gallons of tank capacity.
NOTE: Never use more than one 2-ounce
can of dye to each 100 gallons of fuel.
A very small leak may require an hour or more
for color to appear. If no coloration appears after
a reasonable waiting period, fill the tank to the
two-third level. Add another 2-ounces of dye for
each 100 gallons of fuel added. Wait as before.
Again, if no coloration appears after a reasonable
waiting period, repeat the process at full tank
capacity. The dye will leave a stain, which can be
traced to the source of the leak even after the tank
unit has been emptied.
Prior to an inspection, entry of personnel, or
repair of any fuel tank/cell, specific functions
must be accomplished. These functions are
discussed in the following paragraphs. A definition of each function is provided to allow you to
become familiar with it.
1. Defueling. Defueling is the process of
removing fuel from the aircraft tank/cell.
2. Depuddling. Depuddling is the process of
removing residual fuel from cells/tanks after
defueling and low-point draining. Depuddling is a
necessary step prior to air purging when a nontoxic and noncombustible atmospheric state is
required in a fuel cell or tank.
3. Purging. Purging is the process for
removing fuel vapors capable of producing a
combustible or toxic atmosphere.
NOTE: Do not return the colored fuel to
bulk tanks or trucks, as there is sufficient
dye in a 2-ounce can to color 10,000 gallons
of fuel.
The colored fuel is suitable for use in aircraft
engines since the dye does not have a harmful
effect on the usefulness of the fuel. If you do not
empty the aircraft fuel tank to repair the leak, the
dyed fuel can be burned in the engine. Fuel from
tanks tested with dye will remain colored until the
tanks have been filled and emptied several times.
Stains on the aircraft structure or clothing can be
removed with aircraft fuel or an approved drycleaning solvent.
Before you perform any defueling, depuddling,
or purging on an aircraft, you park it in an area
specifically authorized for such operations. You
must be familiar with the safety precautions and
procedures listed in the maintenance instruction
manuals and NAVAIR wing and squadron
Preparation for Fuel Cell Maintenance
Before any maintenance is performed on a fuel
tank/cell, a check of the applicable aircraft
maintenance manual is required. If the aircraft
maintenance manual is not specific enough to
cover the type of maintenance that is required,
refer to the Aircraft Fuel Cell and Internal/
External Tank manual, NA 01-1A-35, for
additional information. If you find conflicting
information between the specific fuel system
portion of the aircraft maintenance manual and
the NA 01-1A-35, the procedures in the NA
01-1A-35 manual take precedence.
General defueling precautions of aircraft
include the following:
1. Position the aircraft at least 100 feet from
any building or smoking area or in the designated
defueling area.
2. Fire extinguishers must be inspected for
serviceability and manned at all times.
3. The defueler must be parked as far from
the aircraft as possible. It should be parked
heading away from the aircraft in case it becomes
necessary to move the defueler in an emergency.
To protect personnel from the health hazards
associated with aviation fuels, protective clothing
and equipment are required and should be the first
priority before starting any fuel cell maintenance.
Specific items such as respirators, coveralls,
proper shoes, and safety goggles are usually
available for use by personnel. All of these are
required to work with aviation fuels cell or tanks.
Appendix B of NA 01-1A-35 contains specific
information on all of the required safety equipment.
All the required grounding and bonding
cables must be attached before the aircraft
or defueler tanks are opened. Bonding and
grounding wires must be attached to clean,
unpainted, conductive surfaces to be
4. You should always ground the aircraft to
an approved grounding point. The aircraft must
be bonded to the defueler. The grounding cable
for the nozzle must be grounded to a metal part
remote from the tank/cell. This minimizes static
electricity y between the nozzle and the aircraft.
Then you attach the bonding cable from the
nozzle to the aircraft.
5. Personnel requirements are one person to
man each fire extinguisher, one to operate the
defueler, and one person to operate the aircraft
defueling panel. You also need one person to
operate the fuel system control panel inside the
aircraft, if applicable.
tank/cell tested by a gas-free engineer to ensure
the tank/cell is safe for personnel to initiate
depuddling. If after this time a “safe” condition
is not reached, reinstall the air blower for at least
an additional 15 minutes and have the test
repeated. Continue the venting and testing, if
necessary, until the tank/cell can be certified safe
for personnel. The air inside the tank/cell has to
be certified and documented as safe. The outside
safety observer and the individual who is going
to enter the tank/cell should obtain all the
necessary protective clothing and equipment and
proceed with the depuddling.
NOTE: The two individuals should always
be connected by a safety line in case of an
6. Once defueling is complete, drain remaining fuel from low-point drains into an approved
safety container.
The next step in depuddling is to remove all
the necessary access panels and covers required.
Then, immediately after entering the tank/cell,
the individual must cap or seal all openings leading
from other possible sources of fuel or fuel vapors.
Depuddling can be accomplished by using an
approved explosionproof vacuum cleaner. You
can also use a cellulose sponge or cheesecloth to
remove the residual fuel from the tank/cell.
Do not defuel aircraft in the vicinity of an
electrical storm. No maintenance of any
type will be allowed on the aircraft during
When you perform maintenance on a fuel
tank/cell, the next step is purging. There are now
four approved methods you may use to purge the
aircraft fuel tank/cell. They are the air blow, air
exhaust, oil purge methods, and JP-5 method.
The air blow purging method uses an air
blower and ducting to force fresh outside air into
the tank/cell. The air exhaust purge method uses
an air blower and ducting to draw fresh outside
air through the tank/cell. The oil purge method
uses lubricating oil, MIL-L-6081, grade 1010, to
dilute the fuel vapors in the defueled tank/cell.
The oil purge method is the most desirable of the
three methods. It is necessary to perform extensive
repairs to the aircraft other than maintenance
solely related to the fuel system. The oil purge
method will normally keep the tank/cell safe for
personnel for approximately 10 to 15 days. The
JP-5 method uses JP-5 fuel to dilute and help
remove all residues from low flash point fuels
including JP-4 or AVGAS.
Depuddling of the aircraft fuel tank/cell is a
hazardous operation because it requires the entry
or partial entry of personnel into an aircraft
tank/cell. They remove any residual fuel that was
not removed from the tank/cell during defueling.
In an effort to minimize the hazards associated
in depuddling, all maintenance personnel are
required to work in pairs. One person should
remain outside the tank/cell to act as a safety
observer while the other actually enters the
tank/cell to do the depuddling. The following
general safety precautions apply to depuddling:
The aircraft battery connector and aircraft power
receptacle should always be tagged with an
appropriate warning placard. This is to indicate
power is NOT to be applied to the aircraft under
any circumstances. Before you perform any
depuddling, refer to the aircraft maintenance
manual and NA 01-1A-35 for the proper support
equipment that must be used.
When you purge a tank/cell, attach an
approved air blower to the tank/cell and ensure
that all personnel remain clear of the removed
access panel. After allowing approximately 30
minutes for the blower to remove the toxic vapors,
you should stop the air blower and have the
NOTE: In all methods of purging, it is
mandatory that the tank/cell be certified.
This is done by a gas-free engineer and
documented as being safe for personnel or
safe for hotwork.
Carefully remove the cell, observing the
following precautions:
General fuel cell removal and installation
procedures are discussed in the following
paragraphs. These procedures are applicable to
the removal and installation of all fuel cells.
However, the latest technical publications must
be used for actual removal and installation of fuel
cells on any naval aircraft.
1. Do not pull the cell by its fittings.
2. Carefully guide the protruding fittings past
all obstructions.
3. If the cell binds while removing it, do not
force it. Stop to determine the cause of the trouble
and remedy it before continuing. Sprinkle the cell
with talc or other suitable powder if it becomes
necessary to squeeze the cell around or between
structural members.
4. Do not pry on rubber fittings or on the cell
with sharp instruments; use large wooden paddles.
After the aircraft is defueled, depuddled,
and purged, the following steps should be
accomplished for the removal of the cell:
When removal of the cell is necessary because
of major repairs or other reasons, it should be
inspected. You then reinstall it, provided it is fit
for further service in the aircraft. Fuel cells should
be removed when signs of leakage appear. These
signs are rubber particles in the strainer, loose
seams, loose or cracked fittings, or if swollen
sealants are found. If the cell is considered to be
repairable beyond organizational level, it should
be crated and sent to the nearest fuel cell repair
When a fuel cell remains empty for more than
72 hours, a thin coat of oil, Specification
MIL-L-6081, Grade 1010, is applied to the inner
liner. This should be accomplished whether the
cell is installed or removed from the aircraft for
storage. The oil will act as a temporary plasticizer,
and it will prevent the inner liner from drying out
and cracking.
Fuel cells that are to be returned to storage
until repairs can be accomplished at a later date
should have a coating of oil, Specification
MIL-L-6082, Grade 1065. It is applied to the
interior of the cell. The heavier type of oil will
act as a preservative over a sustained period. Oil
should not be applied to the interior of self-sealing
cells that have exposed sealant. It is applied when
the exposed area has been covered with an oilresistant tape. Although complete coverage of the
cell interior is necessary, preservative oil should
not be allowed to puddle in the bottom of the cell.
1. Remove required access covers.
2. Remove all interior parts, lines, clamps,
fittings and plates from cell.
NOTE: Clean dust covers must be installed
on all open tubes, ports, and disconnected
electrical plugs and receptacles.
3. Cap or plug all lines, fittings, and parts
removed from the cell to prevent contamination.
4. Place removed items in a separate container
for each cell, and identify with cell number and
aircraft bureau number.
5. If possible, locate and mark with yellow
crayon (SS-C-635) any damaged areas showing
evidence of leakage.
6. Disconnect cell fittings and interconnects.
Fuel cells are easily damaged. Use caution
when cutting nylon lacing cords.
7. Untie and remove lacing cords. If the cord
is cut during removal, retain old cord to determine
length of replacement cord.
8. Remove the screws or hangers that secure
the cell to the cavity. Install lifting device if
The cell must be handled very carefully to
prevent abrasions, cuts, and punctures. Tape
should be applied to sharp edges of all cavity
openings to eliminate chafing of the cell upon
If necessary, the cell may be collapsed and
strapped in a folded position. Bends should not
occur at any of the fittings.
Handling Procedures
Always carry or haul fuel cells carefully. The
purpose of carrying or handling fuel cells is to
protect the outside (retainer ply) wall. It serves
to support the shape of the cell and protect the
self-sealing (sealant) layer underneath it from fuel
spillage. The cord construction and lacquer
coating must be cautiously safeguarded.
To avoid any undue damage to the cell during
handling, follow the following instructions:
1. Always transport the cell by a well-padded
truck or dolly, or by hand carrying.
2. Never use any of the cell fittings for handholds while carrying the cell.
3. Never allow the cell to be dragged or
rolled on the deck.
4. Before placing the cell on the deck, you
should spread an appropriate barrier material on
the area where the cell will be placed.
5. Never place the cell on a bench, pallet, or
table where parts of the cell are allowed to
6. If the cell was removed during cold
weather, you should warm the cell to at least 60°F
(16°C) before collapsing or folding.
7. Never use unnecessary force or pressure
to compress a collapsed cell into a small package.
The undue pressure will produce sharp folds that
cause damage to the cell.
8. Never allow the cell to be folded across
or beside any of the cell fittings.
9. Never leave a self-sealing cell in a collapsed condition for a period longer than 1 hour.
Bladder-type fuel cells may be left collapsed for
a longer period of time, providing the cell is not
walked on, severely creased, or abused.
10. Always install protective caps on the cell
hanger receptacles while the cell is removed from
the aircraft.
soaked with this solution should be used to apply
the preservative. Fuel cells that are returned to
storage until repairs can be accomplished should
have a coating of oil, MIL-L-6082, Grade 1065,
applied to the interior of the cell. The heavier type
of oil will act as a preservative during the sustained
period. Although complete coverage of the cell
interior is necessary, the preservative oil should
not be allowed to puddle in the bottom of the cell.
When uncrating a fuel cell, you must always
follow the opening instructions on the crate
or shipping container. These instructions are
provided for your use to prevent possible damage
to the cell and to preserve the crate/shipping
container for future use. Before removing the cell
from the container, you should be sure that a
clean, smooth surface, larger than the cell
itself, has been cleared and protected with an
appropriate barrier material before unfolding the
cell. Fuel cells that have been stored for a long
time can shrink or become distorted. Cells in this
condition will be difficult for you to install, and
they often cause misalignment of the cell fittings
with the aircraft fittings. To restore a shrunken
or distorted fuel cell to its original condition, you
should soak the cell in water. The length of time
required for soaking will normally depend on the
condition of the cell. Normally, 72 hours is
enough, as long as the water temperature remains
at least 70°F (21°C). Soaking time can be reduced
by placing the cell in an air-circulating oven at
a maximum temperature of 120°F (49°C) for
about 4 hours. It must also be maintained at high
NOTE: Bladder-type fuel cells and nylon
Pliocels are much more delicate than selfsealing cells and require extremely careful
handling. However, the handling precautions are the same as for self-sealing cells.
When a fuel cell remains empty for more than
72 hours, a thin coat of oil, MIL-L-6081, Grade
1010, should be applied to the inner liner. This
process should be accomplished when the cell is
installed, or when it is removed from the aircraft
for storage. The oil will act as a temporary
plasticizer, and it will prevent the inner liner from
drying out and cracking. To get a uniform coat
on the entire inner lining, you should use an oilsoaked cheesecloth to apply the preservative oil.
Inaccessibility is a problem with some fuel cells
and the only way to properly protect the cell is
to apply the preservative by spraying. Pliocel-type
bladder cells do not require internal preservation
except when they are folded and stored for a
period in excess of 2 weeks. If it becomes
necessary to preserve this particular type of cell,
the inner. lining must be coated with equal parts
of glycerin, MIL-G-491, and water. A cheesecloth
The steps outlined below are generally followed when installing a fuel cell in an aircraft.
Check the cell to make certain that it is the
proper one for the cavity. Tape all cell openings.
Inspect the fuel cell cavity for cleanliness and loose
bolts, nuts, etc.; and make certain there are no
sharp metal or protruding edges that may damage
the cell during or after the installation of the fuel
cell. Tape or otherwise protect the edges of the
fuel cell if necessary.
specifies required torque values and refers to the
applicable bolt torque sequence for securing the
Apply talc or other suitable powder to the
outer surface of the cell and the cell cavity to make
it easier to move the cell into position.
If necessary, collapse or fold the cell as
required, and secure it with webbed straps. The
cell should be warmed to room temperature.
When applying straps, place them and the buckles
so they are easily accessible after the cell is
Guide the fuel cell into the cavity, making sure
it is installed in the right direction. Wooden
paddles with rounded edges may be used to guide
the cell into the cavity; never use tools with sharp
edges or points. If any binding occurs, determine
the trouble and remedy it before damage is caused
to the fuel cell. Be very careful that protruding
fittings are not damaged.
Remove straps if cell was collapsed; then check
the interior of the cell to make certain that no tools
or foreign materials were left inside.
Install all fittings and components. New seals
and gaskets must be used.
When a new or repaired fuel cell is installed
in an aircraft, it should be tested for possible leaks
before it is filled with fuel. The air pressure test
is the best method of determining if any leaks
exist. This test consists of applying air pressure
to a sealed cell and checking for the existence of
leaks with a mercury manometer. Further details
on this type of testing can be found in the specific
aircraft MIM.
The maintenance of the aircraft and engine
fuel system is primarily the responsibility of the
AD rating. Besides fuel cell repairs, some of these
maintenance tasks that are the responsibility of
the AD are discussed in the following paragraphs.
These tasks include inspecting, cleaning, replacing
fuel parts, and the rigging and adjusting of various
fuel system controls.
NOTE: The use of any sealing compounds
on rubber fuel cell fittings is prohibited.
Sealing compounds may be used only on
connections when the adjoining surfaces
are metal.
Torquing Requirements
Fuel System Component Inspection
One of the main causes of fuel leaks is
improper torquing of bolts used to secure fuel cell
access covers, access plates, and cell fittings. Overtorquing or improper torquing sequence causes
excessive rubber cold flow, warps fitting plates,
and, in some cases, breaks the metal insert in fuel
cell fittings. It is important that torquing be
performed properly.
Before the bolts are installed, threads should
be inspected for burrs or other defects that could
damage cell fitting inserts or give incorrect torque
readings. Threaded cell fittings should be inspected to ensure that they are not filled with rustpreventive compound or dirt. Presence of such
foreign material will result in incorrect torque
All bolts should be fully installed fingertight
before they are torqued. Bolts should be of proper
length. A bolt that is short will not safely engage
the mated part; one that is too long will bottom
out, giving incorrect torque values and causing
Each work package (WP) that requires
removal and installation of a fuel cell access cover
Periodically, the entire fuel system must be
inspected for wear, damage, and leaks. Fuel
system parts must be inspected for security of
mounting, leaks, and loose connections. Maintenance should be limited to such items as the
tightening of connections to eliminate leaks and
the replacement of defective units. Repairs
involving the disassembly of units are made at
overhaul activities.
The entire system should be checked for wear,
damage, and leaks. All units must be securely
mounted and all connections tight and properly
safetied. Boost pumps should be used to buildup
pressure to check for leaks.
The boost pump should be checked periodically for proper operation and correct pressure
output. The pump assemblies must be checked for
leaks, the condition of the fuel, and the condition
of the electrical connections. The drain lines must
be free of traps, bends, and restrictions.
Main-Line Strainer
The main-line strainer must be drained at each
preflight inspection to eliminate any water and/or
sediment. The screen must be removed and
cleaned at the intervals specified in the applicable
technical publications. The sediment removed
from the housing should be examined thoroughly.
Particles of rubber are often early warnings of
deterioration of hose or self-sealing tanks. The
strainer must be checked for leaks and damaged
Fuel Lines and Fittings
The lines must be inspected to see that they
are properly supported and that the nuts and
clamps are securely tightened. A hose-clamp
torque wrench should be used to tighten hose
clamps to the proper torque. If this wrench is not
available, the clamp should be tightened fingertight plus the number of turns specified for the
hose and clamp. Clamps should be tightened only
when the engine is cold. If the clamps do not give
a seal at the specified torque, the clamp, the hose,
or both should be replaced. After a new hose has
been installed, the hose clamps should be checked
daily and tightened if necessary. When this daily
check shows that cold flow (the flowing of the
rubber from the clamping area) has ceased, the
clamps should be inspected at the less frequent
periods specified for hose clamps throughout the
system. The hose should be replaced if the plies
have separated, if there is excessive cold flow, if
there are signs of chafing, or if the hose is hard
and inflexible. Permanent impressions from the
clamps in the tube or cover stock indicate excessive
cold flow. Hose that has collapsed at the bends
as a result of misaligned fittings or lines should
be replaced. Some hose tends to flare at the ends
beyond the clamps. This is not an unsatisfactory
condition and does not indicate leakage. At each
engine change, all hose connections forward of
the firewall should be inspected and defective hose
Selector Valves
Selector valves should be rotated and checked
for free operation, excessive backlash, and
accurate pointer indicators. If the backlash is
excessive, the entire operating mechanism should
be checked for worn joints, loose pins, and broken
drive lugs. Defective parts in the operating
mechanism must be replaced. The cable control
systems should be inspected for worn or frayed
cables, damaged pulleys, and worn pulley
High-Pressure Fuel Lines
Since the fuel lines installed on the discharge
side of the engine-driven fuel pump are subjected
to high pressures. You should take special care
when inspecting for leaks and damage. The lines
must be properly connected; otherwise, units such
as the governor and barometric control will not
function correctly and may be seriously damaged.
Combustion Chamber Drain Valve
The engine will have to be turned up to check
the operation of the drain valve. If fuel does not
run from the overboard drain after shutdown, the
cause should be determined.
High-pressure Filter
The high-pressure filter must be inspected for
security of mounting, leaks, and proper safetying.
The filter element must be removed, cleaned, and
inspected at regular intervals. Regardless of their
condition, filter elements must be discarded at the
period(s) specified in authorized maintenance
instructions. Whenever the element is removed,
the housing should be cleaned and the seals should
be replaced.
Fuel Nozzles
Periodically, the fuel nozzles and screens must
be removed and inspected. The screens should be
cleaned and defective nozzles replaced. The
inner surface of the exhaust cone must be
inspected for heavy streaks—discoloration of the
metal due to overheating. These inspections are
rough checks for a combustion chamber in which
the fuel nozzle is not functioning properly.
Ram Air Turbine, Hose,
and Reel Inspection
One of the inspections that is of importance
to you as an AD is the ram air turbine, hose, and
reel inspection. It is performed every 10 hours of
reel operation on the model D-704 air refueling
store. The inspection is accomplished as follows:
d. Reception coupling for binding of
swivel joint.
e. Direct a rearward stream of air (shop
air at 90 psi) across the generator blades and check
the lighting circuit for proper operation.
f. Shroud and shoe assembly for damage.
g. T5 switch arm for damage.
Be sure that the hose jettison cartridge is
removed before performing any maintenance on the store.
1. Remove the filler cap and side access doors
and inspect the interior for corrosion, contaminants, and security of components. Reinstall
cap and doors.
2. Inspect the ram air turbine for the
a. Axial blade freedom. No looseness is
b. Angular blade freedom. A maximum of
10 degrees is permitted.
Ensure that the hose reel snubber valve is
installed prior to performing step 4.
4. Operate the reel for three cycles of
extension and retraction in accordance with the
MIM and check for the following:
a. Presence of hose tension (response) at
fully extended position.
b. Proper operation of tail cone lights and
control panel.
c. Lockpin engagement in reel for stowed
5. Perform the fuel flow test as follows:
a. Fuel the store to 50 gallons minimum.
b. Extend the hose at least 25 feet in
accordance with the MIM.
c. Remove the reception coupling and
insert the hose end in the filler cap opening.
Be sure that your hands remain away from
the turbine blades’ plane of rotation when
performing step c.
c. With the power OFF, rotate the turbine
counterclockwise (viewed from front) until blades
are fully feathered. Hold the spinner stationary
and turn the power ON. The blades should snap
to the unfeathered position.
d. With the power ON, grasp any two
opposite blades and twist toward the feathered
and unfeathered position. Check for binding of
internal mechanism. Turn the power OFF.
e. With the power OFF, attach a 1-inch
C-clamp on the end of the blade 12 inches from
the center of the spinner. Rotate the turbine
counterclockwise (viewed from front) until the
blades are fully feathered. Then rotate the turbine
for at least two more revolutions. To measure
brake slippage, place the hook end of a fishscale
(0-15 pounds) at the C-clamp opening and
continue to rotate the turbine. The force should
be 3 to 5 1/2 pounds.
3. Extend the hose fully in accordance with
the applicable maintenance instruction manual
(MIM) and inspect as follows:
a. Hose and fittings for slippage.
b. Hose for cuts, abrasion, and soft spots.
c. Drogue assembly for torn canopy, bent
struts, and frayed cables.
Be sure that the hose is held firmly in the
tank filler opening when performing step d.
d. Place the control panel TRANS-OFF
switch in TRANS and check fuel flow. Minimum
flow as observed on the store control panel should
be 40 gallons in 15 seconds (160 GPM) if JP-5
fuel is used and 37.5 gallons in 15 seconds (150
GPM) if MIL-L-6081C (1010) oil is used.
Fuel Part Removal/Installation
The most important item for you to remember
when removing or installing any engine part is to
follow the procedures set forth in the applicable
MIMs. The following basics should prove helpful.
1. Make sure all special and required tools are
2. Have a new O-ring kit available for component installation.
While rigging the fuel selector, power control,
and shutoff valve linkages, you should follow the
step-by-step procedures for the particular aircraft
model being rigged. The cables should be rigged
with the proper tension with the rigging pins
installed. The pins should be removed without any
binding; if they are hard to remove, the cables
are not rigged properly and should be rechecked.
The power lever should have the proper cushion
at the IDLE and FULL POWER positions. The
pointers or indicators on the fuel control should
be within limits. You must take all of these things
into consideration while rigging or adjusting the
parts of the fuel system. Also, the fuel selectors
must be rigged so that they have the proper travel
and will not restrict the fuel flow to the engines.
Rigging the fuel control of a turbojet engine
is an exacting job. The power lever assembly and
its related linkage provide manual control of the
engine thrust. The power lever assembly is located
in the cockpit, and its related mechanical linkage
connects it with the fuel control unit of the engine.
Positioning the power lever at any selected setting
mechanically actuates the linkage to the fuel
control unit, resulting in the desired engine thrust.
Modern jet aircraft use various power lever
control systems. One of the common types is the
cable and rod system. In this system, you will find
bell cranks, push-pull rods, drums, fairleads,
3. Have all of the appropriate caps and plugs
available to prevent the contamination of open
fuel lines during component replacement.
4. Have parts bags and tags available to keep
all removed nuts, bolts, and washers in one
location, and be sure to tag all fuel lines before
removal. This makes reinstallation much easier
and less confusing.
5. If the removed component is to be turned
into supply, make sure it is purged and filled with
an approved preservative before turn-in.
This section covers some of the basic
inspections and procedures to be used in the
rigging and adjusting of fuel controls, fuel
selectors, and fuel shutoff valves. Inspect all bell
cranks and rod bearings for looseness, cracks, and
corrosion. Particular attention should be given to
the rod and bell cranks where the bearing is
staked. This area is subject to stress cracking and
corrosion. The adjustable rod ends should be
inspected for damaged threads and the number
of threads remaining after final adjustment. The
drums should be inspected for wear, and the cable
guards should be checked for proper positioning.
If the cables have been loosened, the tension
should be set.
Figure 4-29.-Engine power control system rigging.
would be most concerned with is at the fuel
control and throttle quadrant. See figure 4-29.
Before starting the adjustment of the power
controls at the engine, you should make sure
that the power lever is free from binding. Use a
tensionometer to ensure correct cable tension. If
the power controls do not have full throw or are
binding, the entire system should be checked and
the discrepancies repaired before adjusting the
power control system. Low cable tension may
cause sluggishness or insufficient travel of the
control. High cable tension may result in damaged
pulleys, bell cranks and cables, or vibrations in
the controls.
flexible cables, and sheaves. All of these parts
make up the control system and must be adjusted
or rigged from time to time. On single-engine aircraft, the rigging of the power lever controls is
not very difficult. The basic requirement is to have
the desired travel on the power lever and correct
travel at the fuel control. However, on multiengined jet aircraft, the power levers must be
together or married at all power settings.
The rigging of the proper control cables and
push-pull rods is usually accomplished at the
factory, and no rigging is required except when
a part has been changed. The control system you
After completing this chapter, you will be able to:
Identify the types, characteristics, and requirements of lubricants in aircraft engines; and
recognize sources and procedures for the prevention of oil contamination.
Recognize the two main types and major parts
of lubrication systems.
Identify maintenance procedures of oil
systems and system parts, including oil pressure adjustments; oil filter removal, inspection, and installation; and metal particle
Recognize the goals and requirements of the
Navy Oil Analysis Program (NOAP) to include proper oil sampling techniques, required
NOAP forms and recommended action, and
required logbook entries for NOAP.
The increased complexity of aircraft engines
has added to the requirements for proper lubrication. Jet engines require lubrication to prevent
friction from reducing the engines’ efficiency. Oil
is the lifeblood of the aircraft engine. If the oil
supply to the bearings stops, the lubricating films
break down and cause scoring, seizing, and
burning between moving parts. Fortunately, the
engine oil pump and oil system are dependable.
Like the heart and circulatory system of the
human body, they quietly perform their function
so well we forget their importance.
operating conditions. Maintenance instruction
manuals (MIMs) or maintenance requirements
cards (MRCs) list the type of lubricant required
for specific aircraft maintenance tasks. With an
understanding of the different types of lubricates,
their characteristics, and purposes, you will know
why we must use the proper lubricant. Using the
wrong type of lubricant, mixing different types,
or improper lubrication can cause extra
maintenance man-hours, part failures, and
Lubricants are classified according to their
source—animal, vegetable, petroleum, mineral,
or synthetic. Animal oils are not suitable
lubricants for internal-combustion engines. They
form fatty acids, which cause corrosion when
exposed to high temperatures. Vegetable oils have
good lubricating qualities, but break down (they
change in chemical structure) after long periods
of operation in internal-combustion engines.
Mineral-base lubricants are usually divided into
three groups—solids, semisolids, and liquids.
Petroleum-based oils (for example, MIL-L-6081
The primary purpose of any lubricant is to
reduce friction caused by metal-to-metal contact.
Lubricating oils provide a film that permits
surfaces to glide over one another with less
friction. Therefore, lubrication is essential to
prevent wear in mechanical devices where surfaces
rub together.
The Navy uses many different types of
lubricants. The selection of the proper lubricant
depends on the design of the equipment and the
2. Cooling. Lubricating oil must cool moving
parts by carrying heat away from gears and
bearings. This is an important function considering the many parts located next to burners or
turbine wheels, where temperatures are over
Liquid lubricants cool by pumping or spraying
oil on or around bearings or gears. The oil
absorbs the heat and later dissipates it through
oil coolers.
3. Cleaning. Another major function of a
lubricating oil is cleaning. Oil carries dirt, small
carbon and metal particles, and gum and varnish
to filters. This has become increasingly important
with the higher compression ratios, engine speeds,
operating temperatures, and closer tolerances
between parts in newer engines.
grade) were used in early jet engines. This oil was
distributed in two grades—1010 for normal
use and 1005 for extremely low temperatures.
MIL-L-6081 grade 1010 is still used as a preservation oil in fuel systems.
The types of lubricants used in the engines of
today are different from the lubricants used years
ago. As the power output of jet engines increased,
aircraft were able to fly higher. The result of jet
engines operating at these higher, colder altitudes
and higher engine temperatures created greater
demands on lubricating oils. This, in turn,
required the development of synthetic lubricants
that could withstand these higher bearing
MIL-L-7808 was the first synthetic oil
developed to meet these demands. Today, most
jet engines use another synthetic-based oil,
MIL-L-23699. These two oils are completely
compatible and may be mixed when necessary.
However, certain 23699 characteristics are
downgraded in proportion to the quantity of 7808
oil, if mixed. Synthetic oils are based on acids
and other chemicals; therefore, they are not
compatible with the mineral- or petroleum-based
All lubricating oils used by the Navy have a
classification number, which shows the grade and
intended use of the oil. Aircraft engine lubricating
oils are given a four-digit grade number, such as
1065. The Navy and the Air Force use the Saybolt
scale for designating the viscosity of oil. The
designation consists of four digits. The first
digit designates the use of the oil; the 1 indicates
aviation engine lubrication. The last three digits
give the viscosity using the Saybolt scale.
NOTE: You should consult the applicable
technical instructions for the grade number
or MILSPEC of oil recommended for use
in an engine. Reciprocating engines use
MIL-L-228S1, W-120, or E-120 oil, which
is not compatible with the turbojet engine.
NOTE: You are probably more familiar
with the Society of Automotive Engineers
(SAE) numbers for grading viscosity. If
you want a comparison between the two
systems, take the 3 numbers for the Saybolt
system, divide by 2, and round to the
nearest multiple of 10. For example, 1065
has an SAE rating of 30.
Lubricating oils must perform three basic
functions in a jet engine: (1) lubrication, (2) cooling, and (3) cleaning.
1. Lubrication. Oils should have the following
characteristics to lubricate properly:
a. It must be of low enough viscosity to
flow readily between closely fitted, rapidly moving
parts. It must also have a high enough viscosity
to prevent metal-to-metal wear.
b. It must not break down under high
temperatures and pressures.
c. It must have a low enough pour point
to flow readily when starting under extremely low
d. It must have a high enough flash and
fire point so it doesn’t burn or vaporize under high
e. It should not form and deposit excessive
amounts of carbon, varnish, or gum deposits.
Synthetic oils use military specification
numbers for references. See table 5-1.
Contaminated fuel in the lubricating system
of an engine can be disastrous to engine operation. Lubricating oils can be contaminated
through operational conditions (dusty or sandy
places, high operating temperatures), faulty
maintenance practices, and part failures.
An example of operational contamination is
carbon. Carbon forms when oil evaporates,
Table 5-1.-Classification of Lubricating Oils
especially where heat is concentrated; for example,
in - the bearing compartments near hot turbine
sections. This carbon eventually breaks off and
circulates through the engine lubricating system.
The pieces of carbon are usually not hard enough
or large enough to cause failure of the pumps.
However, they may be large enough to clog small
filters or nozzles. The presence of sand, dirt, and
metallic particles in the lube system is another
source of operational contamination.
Faulty maintenance practices that contaminate
lubricating systems include using the wrong type,
or mixing oils, and improper servicing. The lube
system parts of an engine are made of materials
based upon the type of oil that is to be used.
Synthetic oils attack the common rubber materials
used in the O-ring, seals, and gaskets of lubrication systems that use mineral-based oil. This
attack causes the material to soften, swell, and
lose its ability to seal properly. This condition
causes the oil to leak from the system.
The contamination of oil by rust maybe due
to water in the oil system. There is also contamination from storage containers or servicing equipment. Over time, rust in the lube system will
eventually discolor the bearings. Ordinary rust will
leave a red discoloration on the bearing elements,
and black iron oxides will leave a black indication. These rust particles are not large enough to
cause pump failure.
The contamination of oil by engine fuel can
result from a ruptured fuel-oil cooler. Since the
fuel system operates at a higher pressure than the
lube system, the flow will be into the oil supply.
The presence of fuel in the oil will cause oil
dilution. It also changes the oil properties so the
oil cannot cool and lubricate the bearings
Another serious type of contamination is the
oil itself. Synthetic oil will cloud or form other
contaminants if stored too long. This is why there
is a shelf life for all synthetic oils. Never use
overaged oil. Follow the applicable instruction for
shelf life of synthetic oil (it is usually 6 months)
to prevent problems.
Another type of lubricant ADs should be
familiar with is grease. Grease is used on bearings,
outside the engine lubricating system, control
arms and linkages, and actuators, The most
important requirements of greases are as follows:
1. Stability. It must be free from bleeding
(separation of oils), oxidation, and changes in
consistency during periods of storage and use.
2. Noncorrosiveness. The lubricant must not
chemically attack the various metals and other
material it comes in contact with.
If it is too cold, the oil will resist movement
between the parts and flow too slowly for proper
lubrication. If the oil pressure is too low, it will
not supply enough oil to the bearing for proper
cooling. If the pressure is too high, it may cause
high-speed antifriction bearings to skid and not
roll properly.
It would be impossible to cover all the
different parts of every type of engine oil system
in use today. Therefore, this text presents a
representative sample of various parts common
to different types of oil systems.
3. Water resistance. In some cases, a grease
that is insoluble in water is required. In other
cases, the grease must be resistant only to weathering or during washing.
4. Satisfactory performance in use. The grease
must perform satisfactorily in the equipment and
under the conditions it was intended.
Properties of greases vary with the type of
soap used in manufacturing. Military specifications specify the operating conditions or applications. Table 5-2 contains information on some of
the most frequently used greases.
Engines use a wet-sump, dry-sump, or a
combination of both as lubricating systems. Wetsump engines store the lubricating oil in the engine
or gearbox. Dry-sump engines use an external
tank mounted on the engine or somewhere in the
aircraft structure near the engine. Under the drysump lubrication system is another type called the
hot tank system. You should know the similarity
and operation of these systems.
Oil systems used in jet engines are relatively
simple in design and operation, but their
function is important. The principal purposes of
the oil system are the same as those covered under
lubricating oils—to provide an adequate supply
of clean oil to bearings and gears at the right
pressure and temperature, to remove heat from
the engine, and to remove contaminants from the
system and deposit them in the filters.
The ability of the oil to lubricate correctly
depends upon its temperature and pressure. If the
oil is too hot, it will not have enough viscosity.
Wet-sump System
Engines needing a limited supply of oil and
cooling can use a wet-sump type (fig. 5-1). The
Figure 5-1.-Wet-sump lubrication system.
Table 5-2.-Common Military Lubricants and Their Uses
reservoir for the wet-sump system is either the
accessory gear case or a sump mounted to the
bottom of the accessory gear case. This system
is similar to your car’s engine. In the wet-sump
oil system, the gearbox provides space for foaming and heat expansion because the oil level only
partly fills the casing. Deaerators, in the gearbox,
remove oil from the air and vent the air outside.
The main disadvantages of a wet-sump system
are as follows:
1. The oil supply is limited by the sump
2. It is hard to cool the oil. Oil temperatures
are higher because the oil is continuously subjected to the engine temperature.
3. The system is not adaptable to unusual
flight altitudes, since the oil supply would flood
the engine.
Dry-sump System
The dry-sump system is the most common. In
the dry-sump lubrication system, a tank located
in the airframe or mounted on the engine holds
the oil. This type of system carries a larger oil
supply, and an oil cooler is usually included to
control temperature. The lubrication design of the
engine may use either an air-oil or a fueloil cooler. The axial-flow engines keep their
comparatively small diameter through a streamlined design of the oil tank and oil cooler. A drysump oil system is shown in figure 5-2.
The hot tank system stores oil in the tank
before it is cooled. The oil tank is an integral part
of the accessory drive case. The other parts
include the main oil pump, pump pressureregulating valve, and main oil filter (fig. 5-3).
Figure 5-3.—Hot tank oil system.
As we just discussed, there are two primary
types of oil systems. Some of these parts are
unique to one type of system, while other parts
are used in both systems. The following paragraphs cover oil system parts regardless of type
unless otherwise noted. The main parts of a
typical oil system include an oil tank, oil pumps,
valves, filters, and chip detectors. Other parts are
oil coolers, oil jets, gauge connectors, vents, and
oil system seals.
Oil Tanks
The oil tank stores the system oil supply. An
oil tank may be a simple sealed container (similar
to a car’s fuel tank) where oil is gravity fed to the
engine. Older low-performance aircraft engines
could use this simple tank design. Today’s highperformance aircraft require a more complicated
pressurized type of oil tank; this assures positive
lubrication during all flight conditions.
The dry-sump oil system uses an oil tank
located either in the airframe or mounted on the
engine. See figure 5-4. Oil tanks mounted on the
airframe are normally located within or near the
engine compartment. Additionally, designers
place it high to gain as much advantage as possible
from gravity flow to the oil pump inlet.
A view of a representative oil tank is shown
in figure 5-5. It shows a welded aluminum tank
with an oil capacity of 3.25 gallons and a
0.50-gallon foaming space. The tank is designed
to furnish a constant supply of oil to the engine
in any attitude, and during negative g loading or
forces. This is done by a swivel outlet assembly
mounted inside the tank, a horizontal baffle
mounted in the center of the tank, two flapper
check valves mounted on the baffle, and a 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 tends to rush to the top
of the tank. This happens during decelerations
and inverted flight. Oil trapped in the bottom of
the tank is picked up by the swivel fitting without
any interruption in the flow of oil.
All oil tanks provide an expansion space. This
allows for oil expansion from heat and oil foaming. Some tanks also have a deaerator tray
for separating air from the oil. Usually these
deaerators are of the can type, with oil entering
Figure 5-4.-Dry-sump oil tanks; (A) engine mounted;
(B) engine or airframe mounted.
at a tangent. The air released is carried out
through the vent system in the top of the tank.
The vent system inside the tank is so arranged that
the airspace is always vented. This includes times
when the aircraft is decelerating and oil is forced
to the top of the tank. However, most oil tanks
have a pressurized oil tank to assure a positive
flow of oil to the oil pump inlet. The tank is
pressurized by running the vent line through an
adjustable check relief valve.
Other features common in oil tanks area sump
with drain and shutoff valves in the bottom of
the tank. The drain valve permits oil to be drained
for oil changes. An oil shutoff valve is a motoroperated, gate-type valve attached to the oil sump.
This valve can be operated electrically or manually
Figure 5-5.-Representative oil tank.
to shut off the oil supply to the engine in
emergency conditions.
Some oil tanks include an oil temperature bulb
in the outlet line. These bulbs send the temperature of the oil to indicators in the instrument
panel. Oil quantity units or sight gauges are also
located on the oil tank. A sight gauge gives a
visual indication of oil level. A quantity indicator
connects electrically to a gauge in the instrument
panel. A quantity indicator uses a float-type unit
located in the tank and an electrical transmitter
on the outside of the tank.
Oil Pumps
The oil pump supplies oil under pressure to
engine points that require lubrication. Most
lubrication pumps have both a pressure supply
element and a scavenge element. However, some
oil pumps serve a single function; they either
supply or scavenge the oil. The number of
pumping elements, both pressure and scavenge,
depends largely on the type and model of the
engine. For instance, the axial-flow engines have
a long rotor shaft and use more bearings than a
centrifugal-flow engine. Therefore, the oil pump
elements for both supply and scavenging must be
of larger capacity or more of them.
It is common to use small individual scavenge
pumps in the remote sections of an engine. This
assures proper scavenging of the lubricating oil.
In all types of pumps, the scavenge elements have
a greater pumping capacity than the pressure
element. This is to prevent oil from collecting in
the bearing sumps.
The pumps may be one of several types, each
type having certain advantages and limitations.
The three most common oil pumps are gear,
gerotor, and piston types. The first being the most
used and the last the least used. Each of these
pump types have several different designs. This
makes it impractical to try to completely cover
each type. However, a pump representative of
each of the three types is discussed.
GEAR-TYPE OIL PUMP.— The gear-type oil
pump shown in figure 5-6 has only two elements
(one for pressure oil and one for scavenge).
However, this type of pump could have several
Notice in figure 5-6 a relief valve in the
discharge side of the pump. This valve limits the
Figure 5-6.-Cutaway view of gear-type oil pump.
A set of gerotor pumping elements is shown
in figure 5-7. Each set of gerotors is separated by
a steel plate, making each set an individual
pumping unit. Each set consists of an inner
element and an outer element. The small starshaped inner element has external lobes fitting
within and matching with the outer element,
which has internal lobes. The small element
pinned to the pump shaft acts as a drive for the
outer free-turning element. The outer element fits
within its steel plate, having an eccentric (not
having the same center) bore. In some engine
models, the oil pump has four elements, one for
oil feed and three for scavenging. In other models,
pumps have six elements, one for feed and five
for scavenge. In each case, the oil flows as long
as the engine shaft is turning.
output pressure of the pump by bypassing oil back
to the pump inlet. Also notice the location of the
shaft shear section, which will allow the shaft to
shear if the gears should seize.
GEROTOR OIL PUMP.— The gerotor pump
usually has a single element for oil feed and several
elements for scavenging oil. Each of the elements,
pressure and scavenge, is almost identical in
shape. However, the capacity of the elements is
controlled by varying the size of the gerotor
elements. The pressure element has a pumping
capacity of 3.1 gpm (gallons per minute) compared to a 4.25-gpm capacity for the scavenge
elements. So the pressure element must be smaller
since the elements are all driven by a common
shaft. Engine rpm determines oil pressure, with
a minimum pressure at idling speed and maximum
pressure at maximum engine speed.
PISTON OIL PUMP.— The piston lubrication pump is always a multiplunger type. Output
of each piston supplies a separate jet. Oil drained
from the points of lubrication is scavenged by a
separate pump element. The piston-type pump
(fig. 5-8) is used less than either of the other types.
Valves control the pressure and flow of oil in
the lubrication system. There are three types of
valves common to oil systems that are discussed
in this text. They are relief valves, check valves.
and bypass valves.
Figure 5-7.-Gerotor pumping element.
Figure 5-8.-Axial piston pump.
pressure relief valve limits the maximum pressure
within the system. The relief valve is preset to
relieve pressure and return the oil to the inlet side
of the lube pump. This valve is important if the
system has an oil cooler, since the cooler’s thin
walls rupture easily.
CHECK VALVES.— Check valves installed
in the oil supply lines or filter housings prevent oil in the reservoir from seeping (by
gravity) into the engine after shutdown. Check
valves prevent accumulations of undue amounts
of oil in the accessory gearbox, rear of the
compressor housing, and combustion chamber.
Such accumulations could cause excessive loading
of accessory drive gears during starts, contamination of the cockpit pressurization air, internal oil
fires, and hot starts.
The check valves are usually of the springloaded ball-and-socket type, constructed for free
flow of pressure oil. The pressure required to open
these valves will vary. Most valves require from
2 to 5 psi to permit oil to flow to the bearings.
maintain proper oil temperature by varying the
proportion of the total oil flow passing through
the oil cooler. A cutaway view of a thermostatic
bypass valve is shown in figure 5-9. This valve
consists of a valve body (having two inlet
ports and one outlet port) and a spring-loaded
thermostatic element valve.
The valve is spring-loaded so the valve will
open (bypassing the cooler) if the pressure through
the oil cooler drops too much because of denting
or clogging of the cooler tubing.
The filters are an important part of the
lubrication system, since they remove most foreign
particles in the oil. Without some type of filter
in the oil system, dirt or metal particles suspended
in the oil could damage bearings, clog passages,
and cause engine failure.
The filter bypass valve allows oil to flow
around the filter element if it becomes clogged.
The bypass valve opens whenever a certain
pressure differential is reached because the filter
became clogged. When this occurs, the filtering
action is lost, allowing unfiltered oil to be pumped
to the bearings. This is a dangerous situation;
however, unfiltered oil is better than no oil.
Thermostatic bypass valves are included in oil
systems using an oil cooler. Their purpose is to
Figure 5-9.-Thermostatic bypass valve.
(last-chance filters) for straining the oil just before
it passes through the spray nozzles onto the bearing surfaces.
There are several types of filters used for
filtering the lubricating oil. The filtering elements
come in a variety of configurations. The parts of
a main oil filter-include a-housing, which has an
internal relief valve and a filtering element.
Chip Detectors
The magnetic-chip detector is a metal plug
with magnetized contacts, and is placed in the
DISK-TYPE FILTER.— The disk filter (fig,
5-10) consists of a series of spacers and screens.
The screens and spacers are stacked alternately
in the housing. The spacers direct oil through the
screens as it flows through the assembly. The
screen mesh (usually measured in microns)
determines the size of foreign matter allowed to
pass through the filter.
filter is similar to the cartridge filter used on a
car’s oil filter, as shown in figure 5-11. It uses
either a paper or metal cartridge type of oil filter.
The paper filtering element is removed and
replaced, while the metal type is cleaned and
Each of the oil filter types mentioned in the
preceding paragraphs has certain advantages. The
filter types just discussed are generally used as
main oil filters. These filters strain the oil as it
leaves the oil pump. In addition to main oil filters,
there are also secondary filters throughout the
system. For instance, there may be a finger screen
filter to trap large metal pieces before the magnetic
drain plug. Also, there are the fine mesh screens
Figure 5-11.-Filtering assembly.
Figure 5-10.-Spacers-and-screens oil filter.
scavenged oil path. There are two types. Both
types have magnetized contact points to collect
metal particles. When enough metal particles
collect on the magnetized contacts, one type
completes a circuit between the contacts. This
illuminates a warning light in the cockpit, advising
the pilot of metal contamination. This indicates
that one of the engine gears, bearings, or other
metal parts might have failed. The other chip
detector gives no cockpit indication. It is removed
and inspected at regular intervals for metal
Oil Coolers
Oil coolers reduce the temperature of the oil
through either an air-oil or a fuel-oil type cooler.
These coolers keep the temperature of the oil
within the proper range.
AIR-OIL COOLER.— The air-oil type cooler
(fig. 5-12) is installed at the entry end of the engine
as an integral part of the engine. This type of
cooler is usually an aircraft part conforming to
the inlet duct design of the airframe. This
cooler is made of rectangular-sectioned aluminum
tubing, spirally wound between two end flanges
and formed, by welding, into a cylinder. Two
bosses, located on the horizontal center plane, are
provided for oil inlet and outlet connections. This
type of cooler acts as an inlet air duct; therefore,
a cooling effect occurs when the engine is
operating. The cooling capacity of each of the oil
cooler assemblies also depends upon the amount
of air allowed to pass through the cooler. Some
aircraft use a controllable oil cooler door, which
restricts the opening of the oil cooler exit duct.
FUEL-OIL COOLERS.— The fuel-oil cooler
or heat exchanger (fig. 5-13) cools the hot oil and
preheats the fuel for combustion. Fuel flow to the
engine must pass through the heat exchanger.
However, a thermostatic valve controls the oil
flow, so the oil may bypass the cooler if no cooling
is needed. The fuel-oil heat exchanger consists of
a series of joined tubes with an inlet and an outlet
port. The oil enters the inlet port, moves around
the fuel tubes, and goes out the oil outlet port.
The heat-exchanger type of cooler has the
advantage of allowing the engine to keep its small
frontal area. Since the cooler is flat and mounted
on the bottom side of the engine, it offers little
drag. This type of cooler is an engine part.
Figure 5-12.-Air-oil type of oil cooler.
The most common methods of oil temperature
indicators are a thermocouple-type fitting or an
oil temperature bulb.
Vents are lines or openings to the atmosphere
in the oil tanks or accessory cases of the engine.
The purpose of the vent in an oil tank is to keep
the pressure within the tank from rising above or
falling below that of the outside atmosphere.
However, the vent may be routed through a
pressure relief valve that keeps pressure on the oil
system to assure positive flow.
In the accessory case, the vent (or breather)
is a screen-protected opening that allows
accumulated air pressure to escape. The scavenged
oil carries air into the accessory case, and this air
must be vented. Otherwise, the pressure buildup
within the accessory case would stop the flow of
oil draining from the No. 1 bearing. This oil
would be forced past the rear bearing oil seal and
into the compressor housing. Oil leakage could
cause any of several problems, including high oil
consumption, cockpit air contamination, or a fire.
An oil leakage around the combustion area
or turbine area could cause burning and engine
The screened breathers are usually located in
the front center of the accessory case to prevent
oil leakage through the breather. Some breathers
have a baffle to prevent oil leakage during flight
A vent that leads directly to the No. 1 bearing compartment may be used in some engines.
This vent equalizes pressure around the front
bearing surface. Then the lower pressure at the
first compressor stage will not force oil past the
No. 1 bearing rear carbon oil seal and into the
Figure 5-13.-Fuel-oil heat-exchanger type of cooler.
Oil Jets
The oil jets (or nozzles) are located in the
pressure lines next to or within the bearing
compartments and rotor shaft couplings. The oil
from these nozzles is delivered as an atomized
spray. Some engines use an air-oil mist type of
nozzle spray. This air-oil mist is produced
by tapping high-pressure bleed air from the
compressor and mixing it with the oil. This
method is adequate for ball and roller type of
bearings; however, the solid oil spray method is
Some engines have “expendable oil” jets to
lubricate the bearings supporting the turbine rotor
shaft. The air-oil mist from such jets is not
returned to the tank, but is discharged overboard.
Gauge Connections
Gauge connections are used in the oil system
for oil pressure and oil temperature. The oil
pressure gauge is usually a necessity in all systems
to measure the pressure of the lubricant. This is
done as it leaves the pump on its way to the oil
jets. Since oil pressure is the best indication that
the system is operating properly, the oil pressure
gauge is vital.
Oil System Seals
Any system containing fluids need some type
of seal to prevent fluid loss. The importance
of oil seals cannot be overemphasized! An
improperly installed or leaking seal in the oil
system could cause bearing failure, fire, or cockpit
fumes. This could result in loss of aircraft or
LIFE. There are three types of seals used in jet
engine oil systems—synthetic, carbon, and
The oil pressure gauge connection is always
located in the pressure line between the pump and
the various points of lubrication. The oil
temperature gauge connection may be located in
either the pressure or the scavenge line. However,
the scavenge line is preferred, since this location
permits a more accurate indication of the actual
bearing temperatures, as the temperature of the
oil is shown shortly after it leaves the bearings.
SYNTHETIC SEALS.— Seals, packings, and
O-rings are used where metal-to-metal contact
prevents proper sealing. These seals come in many
different shapes and sizes and are not reusable.
It is important to use the proper seal (identified
by correct part number) for the specific installation. NEVER use a seal, packing, or O-ring
because they look alike. A seal designed to have
excellent sealing characteristics in one environment could be hazardous when used in another.
Some seals swell when contacted with MIL-L-7808
oil, while others deteriorate completely.
LABYRINTH SEALS.— Labyrinth seals
contains series of knifelike, soft metal edges that
ride very close to a steel surface. A certain amount
of air, taken from the compressor, forced between
the steel surface and soft metal edges prevent oil
leakage between sections. These seals were used
as main bearing seals in earlier engines. These seals
are made of very soft metal and used at main bearing areas. Small nicks in the seal can cause major
oil leaks and premature engine changes.
CARBON SEALS.— Carbon seals are used to
contain the oil in the bearing areas. Carbon seals
form a sealing surface by having a smooth carbon
seal rub against a smooth steel surface (faceplate).
All carbon seals are preloaded. Preloading means
the carbon seal is held against the steel surface.
Three common methods of preloading carbon
seals are spring tension, centrifugal force, and air
The engine oil system shown in figure 5-14 is
a representative engine of a self-contained,
pressurized, recirculating, dry-sump system. It
consists of the following systems and parts:
1. Tank
2. Oil pressure and scavenge pump
Figure 5-14.-Engine oil system schematic.
3. Oil filter and condition monitoring system
4. Oil coolers
5. Chip detector
At this location it flows through the scroll vanes
that function as an air/oil cooler. This further
cools the oil and heats the vanes for full-time antiicing. The vanes discharge oil into the oil tank.
If the oil cooler pressure becomes too high, a relief
valve will open to dump scavenge oil directly into
the oil tank.
Oil Tank
The oil tank and air/oil cooler are integral
parts of an aluminum casting. Tank capacity is
7.6 U.S. quarts. The filler port is on the right side
of the engine, and the filler design make it
impossible to overservice the tank. Oil flows to
the oil pump through a screen. The oil level is
shown by a sight gauge on each side of the tank.
The scavenge pump returns oil from the sumps
and accessory gearbox to the oil tank through six
scavenge screens.
Engine Chip Detector
The chip detector is on the forward side of the
accessory gearbox. It consists of a housing with
an integral magnet and electrical connector, with
a removable screen surrounding the magnet. If
there are chips, the completed circuit illuminates
the appropriate number engine CHIP light.
Oil Pressure System
Oil suctioned through the pressure element of
the pump is pressurized and flows through the oil
filter. The oil then flows into passages in the
accessory gearbox and to the six main bearings
in the sumps. A cold-start relief valve downstream
of the filter protects the system by dumping any
extra oil into the accessory gearbox case. Air jets
blow across the oil jets to provide continuous oilmist lubrication. The engine has two sets of oil
jets to provide each main bearing with oil for
cooling and lubrication. Scavenge to the scavenge
elements of the pump flows through screens at
the pump inlet, and then through the electrical
chip detector. The oil then flows through the oil
cooler, main frame, scroll vanes, and into the oil
tank. If the oil pressure drops below 24 psi, the
appropriate ENGINE OIL PRESS caution light
lights in the cockpit.
Maintenance of the oil system is an item of
major importance to the Aviation Machinist’s
Mate. It consists mainly of adjusting, removing,
cleaning, and replacing various parts. To
troubleshoot and repair oil systems effectively,
you should be thoroughly familiar with both the
external and internal oil systems.
The immediate location of any leak or defect
within the oil system of any aircraft engine is
important. The life of the engine is in its oil
supply. Whenever a leak develops or the oil flow
is restricted, a part failure or loss of the engine
may result.
Locating leaks in the external oil system is
easy. Often a visual inspection shows a loose line
or leaking gasket. Although, you should never
assume that an obvious corrective action is all that
is needed.
Oil Filter
Oil discharged from the oil pump is routed to
a disposable element. The element is a 3-micron
filter located on the forward, left-hand side of the
AGB. As the pressure differential across the filter
increases, the first indication will be a popped
impending bypass button. As the pressure increases, the OIL FLTR BYPASS caution light will
illuminate at the same time the filter bypass
A large portion of the maintenance involved
is the replacement of parts and repair of various
oil leaks. Much of this maintenance requires the
use of new gaskets, seals, and packings.
New seals are packaged to prevent damage.
These packages are identified from technical
information printed on the package. This
information identifies the use and qualifications
Oil Coolers
Scavenge oil is cooled before it returns to the
tank by a fuel/oil cooler. After passing through
the oil cooler, oil enters the top of the main frame.
Figure 5-15.-Seal packing information.
Figure 5-17.-Two types of magnetic drain plugs (A) Flakes
from a failed bearing; (B) normal fuzz accumulation.
Figure 5-16.-Removing the oil filter assembly.
of the packing (fig. 5-15). Beside the part number,
the manufacturers cure date is one of the most important items listed on the package. Refer to NAVSUP Publication 4105 for shelf life of preformed
packings. Most synthetic rubbers are not damaged
by several years of storage under ideal conditions.
However, they deteriorate quickly when exposed
to heat, light, moisture, and various other conditions. This is why it is important to keep them in
their original envelopes. Damage also occurs to
packings when improperly stored, such as flattening or creasing from storage under heavy parts.
Before using the parts, inspect new seals for
damage (nicks, scratches, flattening, overage). Do
not use overaged, damaged or nonidentifiable
seals (seals removed from original envelopes).
The difficulty encountered whenever a gasket,
seal, or packing is being replaced is in proper
installation. Always check that the mating
surfaces are clean, and that the new gasket, seal,
or packing is correctly installed. Seals or O-rings
are comparatively soft, so you should use care to
prevent nicks and scratches. Do not use sharp
instruments during installation; they could nick
the seals. Always refer to the applicable maintenance instructions manual for the correct
procedures, tools, and lubricants used during
but consult your specific engine MIMs before
making adjustments.
After any adjustments, you must recheck the
pressure with a direct reading gauge at the
recommended oil temperature and engine rpm.
To identify defects in the oil system that are
attributable to either high or low oil pressure, refer
to table 5-3.
The following procedures are general procedures. You should refer to the correct
maintenance information manuals before you
remove or replace oil filters on your engine. Oil
filters are removed and inspected at regular
intervals. They are also inspected when the cockpit
indicator (chip light) for the magnetic drain plug
warns of possible failure.
1. Provide a suitable container for collecting
oil and remove the filter (fig. 5-16).
2. Inspect the filter for metal contamination.
NOTE: The screen and spacer-type filters
require a special holding fixture for removing filter elements to prevent stacked
screens and spacers from flying off
Before making any oil pressure adjustments,
you should first check the oil pressure with a direct
reading gauge. Record the oil pressure when
running the engine at the recommended oil
temperature and engine rpm. Oil pressure
adjustments are made with the adjusting screw on
the oil pressure relief valve of the oil pump. Turn
the adjusting screw clockwise to increase and
counter clockwise to decrease oil pressure.
3. After inspection, clean the filters. Most
filters are routed to AIMD for ultrasonic cleaning.
4. Install clean or new filters on the oil filter
NOTE: The screen and spacer-type filters
require a special holding fixture for
replacing (buildup) the filter elements. Be
sure the screen and spacers are the correct
number and in proper order.
5. Install the filter assembly using new O-rings
and gaskets. Torque nuts to recommended values.
Some engines prohibit decreasing oil
pressure; the oil pump must be changed
instead. High oil pressure could indicate
blocked oil passages and lowering the oil
pressure could result in an inadequate oil
supply to some bearings.
Magnetic drain plugs are usually removed and
inspected at the same time as the main oil filters.
Remove magnetic drain plugs carefully so
contaminants will not be disturbed until inspected.
Figure 5-17 shows two types of magnetic drain
Oil pressure adjustments vary between different pumps or on pumps with worn parts. One
full turn of the adjusting screw equals about 3 psi,
Table 5-3.-Oil Pressure Troubleshooting Chart
50-50 solder (50-percent tin and 50-percent lead).
A tin particle dropped on the soldering iron will
melt and fuse with the solder.
3. Aluminum (Al). Aluminum particles can
be determined by their reaction with hydrochloric
acid. When a particle of aluminum is dropped into
the hydrochloric (muriatic) acid, it will fizz, and
the particle will gradually disintegrate. Aluminum
particles will also dissolve rapidly and form a
white cloud in a strong caustic solution (sodium
or potassium hydroxide). Silver and copper
(bronze) do not noticeably react with hydrochloric
4. Silver (Ag) and copper (Cu). Silver and
copper (bronze because of its high copper
contents) may be differentiated by their respective
reactions in nitric acid. When a silver particle is
dropped into nitric acid, it will react with the acid,
slowly producing a whitish fog in the acid. When
a particle of copper (bronze) is dropped into the
nitric acid, it will react rapidly with the acid. This
reaction produces a bright, bluish-green cloud in
the acid.
5. Chromium (Cr). These particles may be
determined by their reaction to hydrochloric acid.
When a chromium particle is dropped into
concentrated hydrochloric acid, the acid will
develop a greenish cloud.
6. Cadmium (Cd). Cadmium particles will
dissolve rapidly when dropped into a 5-percent
solution of chromic acid.
7. Tin cadmium. These particles will dissolve
rapidly when dropped into a 5-percent solution
of chromic acid. The tin content will cause a
clouding of the solution.
plugs indicating (A) engine bearing failure and (B)
normal buildup of fuzz.
NOTE: Fuzz consists of fine, hairlike
particles resulting from normal wear. Fuzz
accumulation may be more noticeable on
new engines during the first. 100 hours of
operation. Always refer to your specific
aircraft and engine MIM for contamination and serviceability limits. Rejection
criteria for one engine may be only an oil
flush and oil component replacement on
another type engine.
Metal particles found on the oil strainer
screens and mag plugs indicate a possible failed
part or impending engine failure. The presence
of metal particles on the oil screen or on the mag
plug does not mean that the engine must be
replaced. The type (steel, bronze), shape (flakes,
chunks), and quantity determine the source and
dictate whether or not an engine is serviceable.
The metals usually found are steel, tin, aluminum,
silver, copper (bronze), chromium, nickel, and tin
cadmium combinations. With some experience
you can make a visual inspection as to color and
hardness, and it will help you to identify the metal
particles. The particles of metal found in an engine
may be a granular, flake, or chunk.
When a visual inspection does not positively
identify the metal, the kind of metal may be
determined by a few simple tests. These tests are
performed with a permanent magnet and an
electric soldering iron. You also need about
2 ounces each of concentrated hydrochloric
(muriatic) acid, concentrated nitric acid, chromic
acid, and sodium or potassium hydroxide.
Make sure the metal particles found in the oil
are of an acceptable quantity for the engine to
remain in service. Always refer to the applicable
maintenance instruction manual for the limits of
metal particles for each particular engine.
Always use the appropriate protective
clothing and equipment, and use extreme
care when handling acids.
The Navy Oil Analysis Program (NOAP)
provides a diagnostic technique to monitor and
diagnose equipment or oil condition. This is done
without the removal or extensive disassembly of
the equipment. It is mandatory for all activities
operating aeronautical equipment to participate
in this program. Type commanders or the
cognizant field activity (CFA) are the only ones
to relieve you from this requirement. The CFA
provides information on the sampling points,
The following test procedures help determine
different types of metal particles:
1. Iron (Fe) and nickel (Ni). Use a permanent
magnet to isolate these metal particles.
2. Tin (Sn). Tin particles can be distinguished
by their low melting point. Use a clean soldering
iron, heated to 500°F (250°C) and tinned with a
techniques, and intervals for all Navy equipment.
The CFA also establishes and maintains sampling
information for the maintenance requirements
cards (MRCs). It also maintains maintenance
instruction manuals (MIMs) for the respective
equipment or weapon systems.
Spectrometric oil analysis is a diagnostic
maintenance tool used to determine the type and
amount of wear metals in lubricating fluid
samples. Engines, gearboxes, and hydraulic
systems are the types of equipment most frequently monitored. The presence of unusual
concentrations of an element in the fluid sample
indicates some abnormal wear of the equipment.
Once the abnormal wear is verified and pinpointed, the equipment may be repaired or
removed from service. This is done before a major
failure of the fluid-covered part occurs. This
philosophy enhances personnel safety and
material readiness at a minimum cost, and serves
as a decisive tool in preventive maintenance
action. Thus, worn parts may be replaced prior
to a catastrophic failure.
Wear Metals
Wear metals are generated by the motion
between metallic parts, even though lubricated.
For normally operating equipment, the wear metal
is produced at a constant rate. This rate is similar
for all normally operating equipment of the same
model. Any condition that changes the normal
relationship will accelerate the rate of wear and
increase the quantity of wear metal particles
produced. If the condition is not corrected, the
deterioration will increase and cause secondary
damage to other parts of the assembly. This can
result in the final failure of the entire assembly
and loss of the equipment. New or newly
overhauled assemblies tend to produce wear metal
in high concentrations during the initial break-in
concentration of wear metals in the lubricating
fluid. This analysis is accomplished by subjecting
the sample of fluid to a high-voltage spark. This
energizes the atomic structure of the metal
elements and causes the emission of light. The
emitted light is then focused into the optical path
of the spectrometer and separated by wavelength.
It is then converted to electrical energy, and
measured. The emitted light for any element is
proportional to the concentration of wear metal
suspended in the lubricating fluid.
2. Atomic Absorption—The atomic absorption spectrometer is an optical instrument. It is
also used in determining the concentration of
wear metals in the lubricating fluid. The fluid
sample is drawn into a flame and vaporized. The
atomic structure of the elements present become
sufficiently energized by the high temperature of
the flame to absorb light energy. Light energy
having the same characteristic wavelength as the
element being analyzed is radiated through the
flame. The resultant light is converted to electrical
energy and measured electronically. The amount
of light energy absorbed by the elements in the
flame is proportional to the concentration of wear
NOTE: The spectrometric fluid analysis
method is effective only for those failures
that are characterized by an abnormal
increase in the wear metal content of the
lubricating fluid. This is particularly true
of failures that proceed at a rate slow
enough to permit corrective action. This
is done after receipt of notice from the
The value of a spectrometric analysis is based
on the assumption that the oil sample is representative of the system from which it is taken.
Occasionally, samples from one part may be
substituted for another, resulting in a false
appearance of a developing wear condition. A
sudden increase of wear metal in one part and a
decrease in another should be considered as a
problem related to sample error; for example,
misidentifying a sample as an engine sample when
it was actually a transmission sample.
Identification of Wear Metals
The wear metals produced in fluid lubricated
mechanical assemblies can be separately measured. This is done in extremely low concentrations, by spectrometric analysis of fluid samples
taken from the assembly. Two methods of
spectrometric oil analysis are currently used to
measure the quantity of various metals.
Oil Sampling Techniques
Sampling’ intervals should be as close as
possible to specified times without interfering with
scheduled operations. Generally, the sampling
intervals should not vary more than ±10 percent
1. Atomic Emission—The emission spectrometer is an optical instrument used to determine the
Each operating activity participating in the
NOAP must take routine samples properly
and at the prescribed intervals. In addition to
the routine samples, each operating activity is
required to submit special samples under the
following conditions:
of that time specified. This requirement must be
considered when equipment is scheduled for
detachments or missions away from the home
base. Oil samples will still be due while away. The
customer (squadron or detachment) is responsible
for coordinating oil analysis support at mission
or transit site(s).
1. When samples are requested by the CFA
or by the laboratory.
NOTE: Refer to the applicable scheduled
maintenance or periodic inspection document for the specific routine sampling
interval. Also look for specific sampling
instructions for each type/model/series of
equipment being sampled.
2. When the activity is so directed by the unit
maintenance officer to check out suspected
3. When abnormal conditions exist, such as
malfunction of the oil lubricated part, damage to
the oil lubricating system, excessive engine oil loss,
or zero oil pressure.
4. Before and after the replacement of major
oil lubricating system parts.
5. During and at the completion of a test cell
run. If the repaired or suspect unit is operated on
oil previously used in the test cell system, a sample
must be taken. This is done before and after the
completion of the test cell run.
6. After the final test on an aircraft that is
undergoing rework or scheduled depot-level
maintenance or after installations of new/overhauled engines or engines repaired by AIMD.
7. Following all accidents, regardless of cause
and resulting damage. These samples must be
taken by any means possible to get a representative
There are two basic methods of taking a fluid
sample: (1) the dip tube technique, and (2) the
drain technique.
DIP TUBE SAMPLING.— The following
procedures should be followed when using the dip
tube method for getting a fluid sample:
1. Remove the filler cap from the oil tank and
open the sample bottle.
2. Use a sampling tube of the correct length.
Hold the tube at one end and lower it into the
tank through the filler neck until only the upper
end protrudes. (See fig. 5-18, views A and B.)
3. Allow the lower end of the tube to fill with
oil, then close the upper end with your thumb or
finger. Withdraw the tube and drain the trapped
oil into the sample bottle. (See fig. 5-18, views
C and D.) Repeat this operation until the bottle
Figure 5-18.-Dip tube oil sampling.
has been filled to about one-half inch from the
Do not use mouth suction to fill the
sampling tube. Many oils and fluids are
highly toxic and may cause paralysis or
4. Replace the bottle cap and tighten it to
prevent leakage of the sample. Replace the cap
on the tank and discard the sampling tube.
5. Reduce the chance of misidentifying
samples by marking all oil samples with equipment/system identification as soon as possible
after sampling.
DRAIN SAMPLING.— When using the drain
sampling method for getting a fluid sample, you
should use the following procedures:
1. Open the sample bottle.
2. Open the drain outlet in the bottom of the
tank, sump, case, or drain port, and allow enough
oil to flow through. This washes out accumulated
sediment. (See fig. 5-19, view A.)
3. Hold the sample bottle under the drain and
fill to about one-half inch from the top. (See
fig. 5-19, views B and C.) Close the drain outlet.
4. Replace the bottle cap and tighten it enough
to prevent leakage.
5. Reduce the chance of misidentifying
samples by marking all oil samples with equipment/system identification as soon as possible
after sampling.
NOAP Forms and Logbook Entries
Activities are also responsible for completing
appropriate forms and making entries in the
equipment logbook.
Proper completion of the Oil Analysis Request
(DD Form 2026) is vital (fig. 5-20). Maintenance
actions or recommendations (table 5-4) are based
on information provided by this form and the oil
sample. Incomplete information (oil added since
last sample, hours since overhaul, etc.) could
result in an invalid oil analysis and recommendations. The operating activity must also provide
special reports or feedback information requested
by the oil analysis laboratory or the CFA.
Logbook entries are necessary when starting,
stopping, or changing the monitoring laboratory
for oil analysis. A specific notation is also made
Figure 5-19.-Oil drain sample technique.
Figure 5-20.-Oil Analysis Request (DD Form 2026).
Table 5-4.-NOAP Lab Recommendations
of the NOAP analytical status when transferring
equipment. For complete information concerning
the NOAP, refer to NAVMATINST 4731.1
series. The Joint Oil Analysis Program Laboratory
Manual, NAVAIR 17-15-50, provides instructions
for oil sampling and filling out the sample request
form (DD Form 2026). These instructions include
procedures for submitting samples to the assigned
supporting laboratory, evaluation criteria, and for
getting special technical help.
After completing this chapter, you will be able to:
Recognize the operating principles of
hydraulic systems.
Recognize the different types of ignition
Identify the sources and prevention of
Identify the types and operation of jet engine
starters, and recognize the procedures for safe
operation of aircraft starting equipment.
Recognize systems using fuel for hydraulic
Recognize aircraft power plant electrical
systems and their relationship to other aircraft
ADs deal with a large variety of aircraft
systems. You need a knowledge of hydraulics and
electricity because of these different systems. You
must be familiar with such systems as ignition,
start, bleed-air, and auxiliary power unit systems.
This chapter introduces you to basic hydraulics,
electricity, and the related systems that the Aviation Machinist’s Mate regularly maintains.
Recognize the purpose of the bleed-air
Recognize the use of the auxiliary power unit
systems. Hydraulics apply to fuel and oils systems
too, so a knowledgeable AD must be familiar with
hydraulic principles.
Pascal’s law states that “any force applied to
a confined liquid transmits undiminished in all
directions.” This pressure acts at right angles to
the walls of the container and exerts equal forces
on equal areas. A 100-pound force will result from
5 pounds per square inch of pressure exerted
against a 20-square-inch area. Figure 6-1 shows
a simple hydraulic mechanism that demonstrates
these principles in operation.
Hydraulics is the science of liquid pressure and
flow. In its application to aircraft, hydraulics is
the action of liquids under pressure used to
operate various mechanisms. All modern naval
aircraft use hydraulic systems and hydraulic
The word hydraulics is from the Greek word
for water. Hydraulics originally meant the study
of physical behavior of water at rest and in
motion. Today the meaning includes the physical
behavior of all liquids. A liquid is any fluid whose
particles have freedom of movement among themselves but remain separate. In aviation, hydraulics
usually means the “red fluid” used to operate
landing gear and flight control or propeller
Figure 6-1.-Simple hydraulic mechanism.
Abrasive particles contained in the system are
not flushed out. New particles are continually
created as friction sludge acts as an effective
catalyst to speed up oxidation of the fresh fluid.
A catalyst is a substance that, when added to
another substance, speeds up or slows down
chemical reaction. The catalyst itself is not
changed or consumed at the end of the reaction.
A liquid has a definite volume but no definite
shape. If you put a liquid into a container, it
assumes the shape of that container. Since liquids
are almost incompressible, they transmit pressure
well. Although the application of large forces will
cause a small decrease in the volume, this decrease
is negligible. For more detailed information on
the principles of hydraulics, study the training
manual Fluid Power, NAVEDTRA 12964.
Origin of Contaminants
The contaminants in hydraulic systems can be
traced to four major sources.
Petroleum-based liquids are the most widely
used fluids in hydraulic systems. Refined
hydraulic fluid is clear in color. Red dyes are
added to this fluid so that hydraulic system leaks
are easier to find and identify. Special petroleumbased fluids are used for certain applications. For
example, MIL-H-83282B is the hydraulic fluid
approved for use in, and in the servicing of, Navy
aircraft hydraulic systems. MIL-H-6083C is the
approved hydraulic fluid for the preservation,
packaging, and use in hydraulic test benches.
1. Particles originally contained in the system.
These particles originate during fabrication of
welded-system components, especially reservoirs
and pipe assemblies. Proper design and cleaning
reduce the presence of these particles. For
example, parts designed using seam-welded,
overlapping joints reduce contamination. Parts
designed using arc welding of open sections
increase contamination. Parts designed with
hidden passages, beyond the reach of sandblasting, are the main source of core sand
2. Particles introduced from outside forces.
Particles enter hydraulic systems at points where
the liquid or working parts of the system are in
temporary contact with the atmosphere. Struts
and piston rods are constantly exposed to the
atmosphere. The most common danger areas are
at the refill and breather openings and at cylinder
rod packings. Contamination results from carelessness during servicing and cleaning. Particles
of lint from cleaning rags can cause abrasive
damage in hydraulic systems, especially to closely
fitted moving parts. Rust or corrosion present in
a hydraulic system usually can be traced to
improper storage of materials or parts. Proper
preservation of stored parts helps to reduce
3. Particles created within the system during
operation. Contaminants created during system
operation are of two general types—mechanical
and chemical. Mechanical particles are formed by
the wearing of parts in frictional contact, such as
pumps, cylinders, and packing gland parts. These
worn particles can vary from large chunks of
packings to steel shavings of microscopic size,
which system screens cannot filter.
Experience has shown that trouble in a
hydraulic system occurs whenever the hydraulic
fluid becomes contaminated. The nature of the
trouble—whether a simple malfunction or the
complete destruction of a component—depends
to some extent on the type of contaminant.
Two general classes of contaminants are
abrasives and nonabrasives. Abrasives include
core sand, weld splatter, machining chips, and
rust. Nonabrasives are contaminants resulting
from oil oxidation and the soft particles that are
worn or shredded from seals and other organic
components. Oil-oxidation products, usually
called sludge, have no abrasive properties. Nevertheless, sludge may prevent proper operation of
a hydraulic system by clogging the valves, orifices,
and filters.
The mechanics of the destructive action by
abrasive contaminants is clear. When the size of
the particles circulating in the hydraulic system
is greater than the clearance between moving
parts, the clearance openings act as filters.
Hydraulic pressure pushes these particles into the
softer materials. This results in blocked passages
or scratches on finely finished surfaces from
movement between parts. These scratches result
in internal component leakage and decreased
The chief source of chemical contaminants in
hydraulic fluids is oxidation. Chemical contamination forms as a result of the high pressure
and temperatures acting with the catalytic action
of water, air, and copper or iron oxides. Oiloxidation products appear first as organic acids,
asphaltenes, gums, and varnishes. These products
combine with dust particles and appear as sludge.
Oxidation products that dissolve in liquid increase
a liquid’s resistance to flow. Products that do not
dissolve in liquid form sediments and precipitates,
especially on colder elements such as heat
exchanger coils. A precipitate is a solid substance
that was chemically separated from a solution.
Liquids containing antioxidants have little
tendency to form gums under normal operating
conditions. However, as the temperature increases, resistance to oxidation diminishes.
Hydraulic fluids that are subjected to high
temperatures (above 250°F) will break down,
leaving particles of asphaltene suspended in the
liquid. The red fluid changes to brown, and is
referred to as decomposed liquid. This explains
the importance of keeping the hydraulic fluid
temperature below specified levels.
The second chemical reaction that can produce
impurities in hydraulic systems allows liquids to
react with certain types of rubber. This reaction
causes the structure of the rubber to change,
turning it brittle, and causing the rubber to fall
apart. Make sure that system fluids are compatible
with the seals and hoses, and that those parts are
appropriate for the system.
open containers. Hydraulic fluid absorbs dust and
grit from the air, resulting in contamination
problems. Keep hydraulic parts and servicing
equipment clean. Should the system become contaminated, minimize damage by taking prompt
maintenance action.
Hydraulic system maintenance consists of
inspecting for leaks, contamination, and replacing
parts. External leaks, where fluid is escaping from
a cylinder, valve, or fitting, are usually easy to
A pinhole leak in a 3,000-psi hydraulic
system can force fluid through your skin.
Do not use your hand to feel for a leak.
Inspect the area around the leak, A leak may
not be directly above the accumulation of fluid.
Fluids often follow the structure or tubing to a
lower point before dropping off. When you notice
leaks, trace them to the source, and then repair
or replace the bad unit or part.
Internal leaks are caused by fluid under
pressure slipping past an unseated valve or worn
packing ring. Normally, fluids flow into the return
line back to the reservoir. The signs of internal
leakage are sluggish operation of an actuating
system or a drop-off in system pressure. A drop
in gauge pressure or an indication of insufficient
pressure on the gauge may be caused by an
internal leak. When internal leakage is suspected
or known to be in the hydraulic system, the
symptoms should be noted to aid in locating the
leak. Follow the aircraft technical instruction
for specific troubleshooting and maintenance
procedures. For more information on hydraulic
maintenance procedures, consult the Aviation
Hydraulics Manual, NAVAIR 01-1A-17.
4. Particles introduced by foreign liquids.
Water is the most common foreign fluid contaminant, especially in petroleum-based hydraulic
fluid. Water enters through condensation of
atmospheric moisture and normally settles at the
bottom of the reservoir. Fluid movement in the
reservoir disperses the water into fine droplets.
These water droplets form an oil-water-air
emulsion because of the mixing action created in
the pumps and passages. This emulsion normally
separates during the rest period in the system
Contamination Control
Filters provide adequate control of the contamination problem during all normal hydraulic
system operations. Contamination control from
outside sources are the responsibility of
maintenance personnel. Therefore, take all
precautions to be sure contamination is held to
a minimum during service and maintenance. Do
not reuse fluid drained from hydraulic systems or
equipment. Do not use hydraulic fluid stored in
Fuel, and the means to regulate fuel pressure,
is readily available within the aircraft. It is an ideal
fluid to use for hydraulic control of systems.
Especially those systems controlling certain engine
functions, such as compressor guide vane variable
geometry actuation, and operation of the afterburner variable nozzles.
Figure 6-2.—B1eed-air and vane control system.
Figure 6-3.—Fan and compressor variable geometry systems.
variable IGVs and variable stators. These systems
are the fan variable geometry (FVG) system and
the compressor variable geometry (CVG) system.
Compressor IGV and stator angle are changed by
setting the CVG pilot valve to direct fuel pressure
to two CVG actuators. A torque motor in the
single FVG actuator sets the CVG pilot valve.
See figure 6-3.
Combination Inlet Guide Vane
and Bleed Valve System
The TF41 engine is a good example of a system
that uses fuel as a hydraulic fluid. Fuel pressure
hydraulically operates the inlet guide vanes (IGVs)
and the bleed valve ring. The vane control and
the bleed-air system prevent compressor surges
during low rpm range operation. The system uses
engine fuel pressure as control and hydraulic force
to vary the angle of the high-pressure compressor
variable IGVs. Fuel pressure is used to operate
the high-pressure compressor bleed air valve. See
figure 6-2. This action decreases the airflow
through the rear stages of the compressor.
Compressor surge and choking caused by increased air velocity is prevented.
Variable-Area Exhaust Nozzles
The convergent-divergent geometry of the
variable-area exhaust nozzle provides the optimum
exhaust throat area. The action of the variablearea nozzle is achieved by close tolerance overlapping seals, which bridge adjacent leaf segments
to provide a relatively smooth surface contour.
The TF30-P-414/414A installed in the F-14 aircraft uses the hydraulic action of fuel to position
the nozzles. See figure 6-4.
Variable Inlet Guide Vanes and Stators
The F404-GE-400, installed in the F/A-18 aircraft, uses two different systems to operate the
Figure 6-4.-Variable area exhaust nozzle actuation.
the circuit closes. When the circuit is open, the
current cannot flow and its value, or voltage,
drops to zero until the circuit closes again.
Alternating current is current that changes
direction in the circuit, flowing first in one
direction and then in the other. A cycle is two
complete alternations within a period of time. The
hertz (Hz) indicates one cycle per second. For
example, one cycle per second is 1 hertz. The
standard unit of electricity used in the United
States is 60 Hz ac.
The electromotive force is the force that causes
electrons to flow from atom to atom in a
conductor. Electrons flow from atoms with an
excess of electrons to atoms with less electrons.
The practical electrical units are the volt, the
ampere, and the ohm. The volt is the unit of
electromotive force necessary to cause electrons
to flow in a circuit. The ampere is the rate of flow
of these electrons in a conductor. The ohm is
called the unit of electrical resistance. This
resistance varies according to the kind of material
used as a conductor, the length of the conductor,
and the cross-sectional area of the conductor.
Resistance also varies with temperature. Other
factors being equal, the resistance increases if the
length of the conductor is increased, and decreases
if the cross-sectional area of the conductor is
increased. Resistance increases as the temperature
You should be familiar with the aircraft
electrical system in general. You must become
familiar with the different engine electrical systems
and components that support your type of engine.
Some of the electrical systems that ADs maintain
include ignition, starting, thermocouple, temperature control, and constant-speed drive (CSD)
systems. To understand the operation of these
systems, you must understand basic electricity.
The following paragraphs cover some of the basic
facts and laws that will be helpful in understanding electrical principles.
Electricity is a form of energy. Energy is the
ability of a body to do work. It does not occupy
space, but it can be measured. There are two
forms of electricity-–static and dynamic. Static
electricity is electricity at rest. It is produced by
friction, which causes one body to give up
electrons to another. The body that lost the
electrons will have a positive charge. The body
that gained the electrons will have a negative
charge. As a result, the positively charged body
will try to gain electrons. The negatively charged
body will try to cast off its surplus electrons to
an oppositely charged body or a neutrally charged
body. Lightning is an example of static electricity
during the discharge of electrons. Dynamic
electricity is electricity y in motion. This is the most
useful form of electricity, and it is produced,
controlled, and measured with relative ease. It is
called electricity in motion because it will flow
along a definite path, called a circuit. Static
electricity will remain in the body containing it
until discharged.
There are three methods of producing electricity. These methods are heat, chemical, and
mechanical means. When two dissimilar metals
in contact with each other are heated, an electron
flow takes place between them. Thermocouples
are a good example of this flow. The storage
battery is a good example of converting chemical
energy into electrical energy. The mechanical
means of producing electricity will be of the most
interest to you. The generator and magneto are
two methods used to mechanically produce
There are three types of current that these
mechanical devices produce. They are direct,
pulsating direct, and alternating current. Direct
current is current that always flows in the same
direction. Pulsating direct current is direct
current that is interrupted by a set of breaker
points. The current will flow in one direction when
As an Aviation Machinist’s Mate, your
primary responsibility is to maintain power plants
and related systems. You should know that engine
systems support the entire aircraft. Engine systems
support more than just the engine. This means
that the maintenance of these systems are the
responsibility of more then one work center.
Electrical or pneumatic systems that other work
centers help maintain include ignition, starting,
bleed-air, and auxiliary power units (APU).
Jet engine ignition systems are simple compared to automobile ignition system. They are
simple because jet engine ignition systems require
no ignition timing. Since jet engine combustion
is a self-sustaining process, a spark is needed only
classified as ac or dc systems, and they use either
a high- or low-voltage capacitor.
during the start cycle. After combustion starts, the
ignition system may be turned off. However, some
aircraft use continuously operating ignition systems
to ensure an immediate relight in case of flameout. Pressure switches or mechanical linkages that
automatically reactivate the ignition system are
also used in some aircraft for the same reason.
The ignition systems on all jet engines are
basically the same, but terminology varies between
engine manufacturers. The part that goes in the
combustion chamber to supply spark is called a
spark plug, an igniter plug, or a spark igniter.
They may look a little different and maybe called
by different names, but they all do basically the
same job. Modern jet engines require an ignition
system with a high voltage and high heat spark.
The high-energy, capacitor-discharge ignition
system is the most widely used ignition system.
It provides a high-tension spark capable of
blasting carbon deposits and vaporizing large
amounts of fuel. This high-energy system makes
starts with carbon-fouled igniter plugs possible,
and it also helps in air restarts at high altitude.
High-energy, capacitor-discharge systems are
High-Energy, Capacitor-Discharge
Dc Ignition System
The ignition exciter gets its input from the lowvoltage dc supply of the aircraft electrical system.
See figure 6-5. The ignition system has three
major components. The system consists of one
ignition exciter and two lead assemblies. The
exciter unit is hermetically sealed to protect
internal components from moisture, foreign
matter, pressure changes, and adverse operating
conditions. This type of construction eliminates
the possibility of flashover at high altitude due
to pressure change and ensures positive radio noise
shielding. The complete system, including leads
and connectors, is built to ensure adequate
shielding against leakage of high-frequency
voltage. High-frequency leakage would interfere
with radio reception of the aircraft. The system’s
primary purpose is to supply energy to two spark
Figure 6-5.-Typical sealed ignition exciter box.
Figure 6-6 is a functional schematic of the
system. You should refer to this figure when
studying the theory of operation of a capacitordischarge system. This schematic shows a camoperated breaker point. Most modern systems
have had all mechanical parts replaced with
electronic solid-state devices. System operation is
discussed in the following paragraphs.
A 24-volt dc input voltage is sent to the input
receptacle of the exciter. The voltage first goes
through a noise filter. This filter blocks conducted
noise voltage from feeding back into the aircraft
electrical system. This input voltage operates a dc
motor, which drives one 16-lobe and one 1-lobe
cam. The input voltage is also sent to two breakers
actuated by the 16-lobe cam.
From the breakers, a rapidly interrupted
current is sent to the autotransformer. When the
breaker closes, the flow of current through the
primary winding of the transformer generates a
magnetic field. When the breaker opens, the flow
of current stops. The collapse of the field induces
a voltage in the secondary windings. This voltage
causes a pulse of current to flow into the storage
capacitor through a rectifier. This rectifier limits
the flow to a single direction. With repeated
pulses, the storage capacitor thus assumes a
charge, up to a maximum of approximately 4
joules. One joule per second equals 1 watt of
The storage capacitor connects to the spark
igniter through the triggering transformer and
through a normally open contactor. When the
charge on the capacitor has built up, the contactor
closes by the mechanical action of the single-lobe
cam. A portion of the charge flows through the
primary of the triggering transformer and the
capacitor connected in series with it.
This current induces a high voltage in the
secondary, which ionizes the gap at the spark
igniter. Thus, when the spark igniter conducts,
the storage capacitor discharges the remainder of
its accumulated energy through it. Energy also
comes from the charge from the capacitor that
Figure 6-6.-Functional schematic diagram of a capacitor discharge ignition system.
is in series with the primary of the triggering
The spark rate at the spark igniter varies in
proportion to the voltage of the dc power supply.
This varying voltage affects the rpm of the motor.
Since both cams are geared to the same shaft, the
storage capacitor always accumulates its store of
energy from the same number of pulses before
The use of the high-frequency triggering
transformer, with a low-reactance secondary
winding, holds the discharge time duration to a
minimum. This concentration of maximum
energy in minimum time achieves an ideal spark
for ignition. This spark is capable of blasting
carbon deposits and vaporizing globules of fuel.
The capacitor, constructed integrally with the
exciter unit, is sealed separately in its own case.
All high voltage in the triggering circuits is isolated
from the primary circuits. The complete exciter
is sealed against the escape or entry of air. This
type of construction protects all parts from
adverse operating conditions and eliminates
flashover at altitude. This design also shields
against leakage of high-frequency voltage that
could interfere with radio reception in the aircraft.
High-Energy, Capacitor-Discharge
Ac Ignition System
The ignition system is an automatic, intermittent duty, at-powered, electronic capacitor
discharge system. See figure 6-7. It is used for
Figure 6-7.-Jet engine electronic ignition system.
initiating engine combustion during aircraft
armament firing, during starting, and for
automatic reignition in case of engine flameout.
The ac ignition system consists of an ignition
exciter, a control amplifier, leads and ignition
plugs, and an alternator stator.
is a dual-circuit and dual-output unit that supplies
a high-voltage, high-energy electrical current for
ignition. The exciter consists of a radio frequency
interference filter. The exciter also contains
power, rectifier, storage, and output elements.
The exciter is on the forward part of the compressor section of the engine.
amplifier is the electronic control center of the
engine. It controls the function of the ignition
system as well as other engine functions. The
amplifier is on the compressor section aft of the
engine front frame.
PLUGS.— The ignition leads are high-tension
cables that “transmit electrical current from the
exciter to the igniter plugs. The igniter plugs are
located in the combustion chamber housing.
alternator is an engine-driven, single-phase, ac
electrical-output unit on the engine accessory gearbox. It supplies the engine with electrical power
independent of the aircraft electrical system. It
contains three sets of windings. Two windings
supply electrical power to the ignition exciter, and
the third winding supplies electrical power to the
control amplifier.
ignition windings, through the ignition exciter,
to the igniter plugs. Current flows from the
alternator, through the control amplifier, to the
ignition exciter. At the ignition exciter, current
increases and discharges as a high-voltage output,
which is conducted through the igniter cables to
the igniters. Current crossing the gaps in the
igniters produces a continuous high-intensity
spark to ignite the fuel mixture in the combustion chamber. When engine speed reaches 8,500
rpm Ng and interturbine temperature (ITT)
reaches operating range, a T5 signal is generated.
This signal goes from the T5 temperature
detectors through the T5 circuit, and to the control
amplifier ignition logic circuit. The control
amplifier ignition relay opens and ignition stops.
Combustion then continues as a self-sustaining
Ignition is automatically reactivated when
either a flameout occurs or when aircraft armament is fired. When T5 temperature drops in
excess of 800°F (427°C) from T5 selected by PLA,
a signal transmitted from the T5 detectors causes
the control amplifier T5 flameout logic to close
the amplifier ignition relay. This activates ignition system operation. Ignition continues until
engine operating temperature is again attained and
the 800 “F temperature error signal is canceled.
This action causes the control amplifier to end
ignition operation.
An armament-firing protection circuit
prevents flameout from armament gas ingested
by the engine. When the aircraft armament is
fired, the armament-firing logic circuit is activated
by a signal from the armament trigger switch. The
amplifier logic then causes ignition operation to
be activated. Ignition operation ends after a
1-second time delay in the amplifier logic following release of the armament firing trigger.
Ignition Operation
Ignition System Maintenance
The operation of an ignition system is
explained in the following paragraphs. Refer to
figure 6-7.
Starting procedures call for the ignition switch
ON, the engine cranking for starting, and the
throttle advanced to the 10-degree power lever
angle (PLA) position. At that time, current flows
from the alternator stator to power the control
amplifier. Simultaneously, the PLA ignition
switch in the fuel control closes. The gas generator
speed (Ng) logic circuit will close the ignition relay
to provide ignition whenever the Ng is within the
10 to 48 percent Ng range. With the relay closed,
a circuit is completed from the alternator stator
The only maintenance performed at the
organizational level is cleaning and replacement
of ignition parts. Ignition parts are sealed units
and must be replaced as complete assemblies.
Degrease spark igniters and clean the outer
shell with a wire brush. If deposits exist on
the ceramic tip and on the center and ground
electrodes, remove them by light abrasive blasting.
However, this abrasive blast should not be used
on the ceramic barrel surface. Clean the ceramic
barrel of the igniter with a soft swab and a suitable
solvent. Dry the igniter with compressed air.
Visually inspect the barrel and shell threads. If
necessary, clean the barrel and shell threads with
a die. Visually inspect the exposed ceramic
section. Any cracks are cause for rejection.
The functional testing of the capacitor
discharge ignition system is a simple operation.
You should exercise caution while performing this test. Do not come into contact
with the igniter plugs or leads while the
power is on in the ignition system. Some
systems have voltages up to 28,000 volts
or more. These high voltages could cause
serious injury or death. Prevent fuel or fuel
fumes from gathering under the engine
while the igniter plugs are being ground
Perform a dry run of the engine (operate the
ignition system). Listen at the tailpipe. You can
determine if the unit is working by listening for
the spark. Another way is to remove the spark
igniters, leave them hanging on the high-tension
leads, and operate the ignition system. The spark
can be seen at the plug if the unit is working. The
spark should be brilliant and accompanied by a
sharp report.
Starting a jet engine requires rotating the
compressor fast enough to begin the engine
combustion cycle. Starting systems must be
capable of providing both high starting torque and
high speed. High starting torque is required to
overcome the large amount of weight of the engine
rotor. High speed is required to increase rotor rpm
until the rotor is self-sustaining. There are several
ways to accomplish these objectives. The following paragraphs describe four methods. They are
the air turbine starter, the direct turbine impingement starter, the electrical starter, or the hydraulic
Figure 6-8.-Air turbine starting system.
Compressed air, supplied to the scroll inlet,
is sent to the turbine wheel by the nozzle in the
scroll assembly. The reduction gear system
transforms the high speed and low torque of the
turbine wheel to low speed and high torque at the
output shaft. An overspeed switch mechanism is
used to limit maximum rotational speed. When
the desired starter rotational speed is reached, the
fly weights in the governor assembly will open the
limit switch. This section sends a signal that shuts
off the supply of air. At a higher, predetermined
rotational speed, the overrunning clutch assembly
disengages the output shaft from the rotating
Air Turbine Starter
The air turbine starter is a lightweight unit for
starting engines with compressed air. The starter
is a turbine air motor equipped with a radial
inward-flow turbine wheel assembly, reduction
gearing, splined output shaft, and a quickdetaching coupling assembly. See figure 6-8.
Turbine Impingement Starter
Some naval aircraft are started by means of
low-pressure air directed onto the turbine or
compressor blades. This little used method is
called impingement starting. The number of air
outlets and the air pressure required for impingement starts vary with the size, weight, and design
parameters of the engine to be started. Refer to
figures 6-9 and 6-10.
Starting air is supplied by an external starting
unit, and is controlled by an air shutoff valve in
the external air supply unit. The shutoff valve is
electrically controlled by the applicable engine
starter switch through the start relay. The starting
air is delivered through a flexible hose to the
starting manifold connection. The air is then
ducted through the impingement manifold and
directed onto the second-stage turbine blades.
A check valve in the starting manifold prevents
loss of gases after the engine is started. The
external starting unit electrical connector must be
plugged into the receptacle corresponding to the
engine being started. In addition to the air supply,
external electrical power must be applied to the
aircraft before the engines can be started. The
function of the turbine impingement starting
system is to start and sustain engine rotation at
a speed at which the fuel-air mixture is satisfactory
for light-off. The system must also assist the
engine to increase rpm until it is self-sustaining.
Electrical Starters
Electrical starters are 28-volt dc series wound
motors, designed to provide high starting torque.
Their use is limited to small engines because of
the high current drain on the electrical source
and their heavy weight. Electrical starters develop
a lot of heat while cranking. Starter damage
from heat is prevented by observing maximum
cranking time and time intervals between start
Figure 6-9.-Turbine impingement starting.
Figure 6-10.-Turbine impingement starting system flow schematic.
Hydraulic Starters
Hydraulic starters, like electrical starters, are
energy-limited starting systems. Energy-limited
starting systems are designed to start the engine
in a short time period, and are limited to small
engines. They make ideal starters for auxiliary
power units (APUs).
It is important to understand what is meant
by bleed air and where bleed air comes from.
Bleed air is tapped off of pressurized air from the
engine compressor section. On some engine
configurations, the air is bled from more than one
area of the compressor. This design gives a source
for high- and low-pressure air to suit whatever
requirements a particular system may have. Other
engines have only one source of bleed air available
from each engine. That source is often tapped
from the last stage of compression on each engine.
Components include a high-pressure accumulator and a variable displacement motor.
The variable displacement motor permits high
torque to be applied without exceeding cutoff speed limits. A small electric motor or
hand pump charges the accumulator. The
accumulator then supplies power to the starter.
The engine has a number of different uses for
the air pressure it generates while operating.
Figure 6-11.—Bleed-air system.
Besides thrust, bleed air can be used for engine
starting, seal pressurization, and anti-icing.
Oil and Seal Pressurization System
Air pressure, bled from the compressor,
controls oil leakage on nonrubbing labyrinth or
clearance-type bearing seals. These bearings use
the differential in sump pressures to keep oil loss
to a minimum. The sump scavenge pump capacity
is greater than that of the oil system pressure
pump. Not only does the sump scavenge pump
scavenge all the oil in the sump area, it also
scavenges air in the sump, creating a lower air
pressure than that of the area surrounding the
bearing sump. This action allows the compressor
bleed air external from the sump to flow from
outside the sump area across the bearing seal,
preventing oil leakage in the opposite direction.
The airflow also helps cool the bearings. See
figure 6-12.
Cross-bleed Air Engine Starting System
Most aircraft in the fleet using two or more
engines employ a cross-bleed starting system.
See figure 6-11. This system provides regulated
air pressure from one engine to start the remaining engine(s). The first engine must be started by
an external source of air pressure. External
sources may be auxiliary power units or ground
support equipment. Subsequent engines can then
be started using bleed air from the running engine.
Opening the cross-bleed air valves allows regulated
bleed air from the running engine to supply air
to the other engines’ starter.
Anti-icing System
Compressor bleed-air valves reduce the load
on the compressor, making it easier for the starter
to turn the compressor. During starts, air is bled
from the compressor through ports on the compressor housing. The bleed valves are held open
by compressor air pressure until the engine starts.
After starting, the speed-sensitive valve directs
compressor discharges air to close the bleed
The guide vanes of a turbine-powered engine
are used to direct the flow of inlet air into the
compressor section. The air is coldest at this point,
and is subject to icing. The biggest problem
resulting from ice forming at this point is the
blockage of inlet air, which causes air starvation,
and thus engine failure. Another problem is the
Figure 6-12.-Engine bleed-air distribution.
possibility of inducting chunks of ice into the
engine. Engine anti-icing systems prevent these
problems if turned on prior to entering an icing
condition. Icing will not normally occur in supersonic flight because heat caused by the friction
of the aircraft passing through the air is sufficient
to prevent ice from forming.
Many types of anti-icing systems are in use
today. All systems use bleed air from the engine
to perform the anti-icing function. The use of
bleed air causes engine power loss. Anti-icing will
be used only when absolutely necessary. Some aircraft use a reversible electric motor to open and
close an air valve to supply the needed air. Other
aircraft use an electrical solenoid to control a
pneumatic anti-icing valve. See figure 6-13.
When the aircraft routinely flys in adverse
weather conditions, a fail-safe system may be used
in the system. The solenoid-actuated air valve is
electrically actuated closed. If the switch is turned
on, or if electrical power is lost, the valve is spring
loaded to the open position. Some systems antiice the complete inlet duct, while in other systems
only the guide vanes are anti-iced.
There are a several airframe systems that rely
on engine bleed air to operate. See figure 6-14.
Figure 6-13.-Inlet guide vane anti-icing system.
These systems include air-conditioning and
pressurization, electronic equipment cooling,
windshield washing, anti-icing, and anti-g
systems. The bleed-air system also pressurizes fuel
tanks, hydraulic reservoirs, and radar waveguides
on several types of aircraft.
In addition to supplying aircraft systems with
bleed air, some aircraft manufacturers use it to
provide extra lift to the wings. Design engineers
devised a system to duct engine bleed air across
the leading edge of the wing to increase the lift,
generating airflow. This system decreases the aircraft stall speed and increases its slow flight
capability, which is desirable during aircraft
carrier landings.
Airframe Deicing and
Anti-icing Systems
On foggy days (visible moisture in the air), ice
can form on aircraft leading edge surfaces at
altitudes where freezing temperatures start. Water
droplets in the air can be supercooled to below
freezing temperature without actually turning into
ice. Ice forms when these droplets are disturbed
in some manner. This unusual occurrence is partly
due to the surface tension of the water droplet
not allowing it to expand and freeze. However,
when the aircraft surfaces disturb these droplets,
they immediately turn to ice on the aircraft
surfaces. The ice may have a glazed or rime
appearance. The glazed ice is smooth and hard
to detect visually. The rime ice is rough and easily
Frost forms as a result of water vapor being
turned directly into a solid. It can form on
aircraft surfaces in two ways. It can collect
on aircraft parked outside overnight when the
temperature drops below freezing and the proper
humidity conditions exist. It can form on aircraft
surfaces when the aircraft descends rapidly into
warm, moist air after flying at higher cold
altitudes. In this case, frost forms because of the
Figure 6-15.-Deicer boot location/operation.
cold air coming off the aircraft structure as it
warms up.
prevent the formation of ice. The deicing systems
are common to older aircraft. The anti-icing
systems are common to newer aircraft. Both
systems use bleed air to accomplish their
As shown in figures 6-15 and 6-16, the
deicing system uses bleed air to inflate the
rubber boots along the leading edge of the wing.
The cells or tubes of the deicer boots are inflated
and deflated alternately by pressure and suction,
causing a wavelike motion, which cracks the
formed ice and allows it to be carried away by
the airstream. The system is pneumatically
operated, electrically controlled, and regulated by
a pressure regulator and relief valve. Suction and
pressure gauges provide a means of monitoring
the system operation.
The shape of the wing may be changed
drastically because of the formation of ice or
frost. Control of the aircraft may become difficult
as the lift characteristics of the wing change.
Uneven distribution of ice generates an unbalanced aircraft condition. Enough ice to cause
an unsafe condition can form in a very short
period of time. Thus, some method of ice removal
or prevention is necessary.
Presently, there are two methods for eliminating or preventing ice. One method, deicing,
employs a mechanical system to break up and
remove the ice after it has formed. The second
method, anti-icing, uses heated bleed air to
Figure 6-16.-Deicing system.
Internal tanks are pressurized anytime the
engine is running, provided that electrical power
is on, the refueling probe is retracted, the tad hook
is up, and weight is off the wheels.
Fuel transfer from external tanks to the
main airframe tanks using bleed air is also
available, if the above conditions are met. During
emergencies, troubleshooting, or checking fuel
transfer after installing external tanks, an OVERRIDE switch is installed that defeats all conditions
except that the tailhook must be up.
The anti-ice system shown in figure 6-17 is a
combination of both deice and anti-ice systems,
and is called an ice protection system. Using bleed
air, temperature sensors, thermostatic switches,
and various types of valves, ice protection
requirements are met.
Airframes Fuel Systems
The pressurization and vent system provides
regulated bleed air pressure to all fuel tanks. This
prevents fuel boil-off at altitude and provides a
means to transfer fuel between tanks. This system
also provides pressure relief of the fuel tanks
during ascent and vacuum relief of the tanks
during descent if the pressurization system fails.
Several types of aircraft now in the fleet
have an onboard auxiliary power unit (APU).
Figure 6-17.-Ice protection system.
Figure 6-18.-Typical onboard auxiliary power unit (APU).
APUs are small, self-contained jet engines that
are started either electrically through onboard
batteries or hydraulically through a hydraulic
starter motor. See figure 6-18.
In the past, APUs were too large and too
heavy for practical use in tactical combat aircraft.
Their use was limited to the larger land-based aircraft with missions such as patrol, cargo,
transport, or special projects. Advancements
in technological design and metallurgy have
produced small, lightweight, yet efficient APUs.
These advancements have enabled newer carrierborne aircraft, such as the S-3 and the tactical
F/A-l 8, to operate aboard ship without the
requirement for flight deck support equipment.
This places less demand on the flight deck crews,
and makes the flight deck a somewhat safer
working environment.
The use of an APU makes the modern jet aircraft completely self-sufficient. Aircraft having
air turbine starters can use compressed air from
the APU to start engines. They also supply
electrical and hydraulic power, as well as air
conditioning during ground maintenance. The aircraft is independent of the need of ground power
units to carry out its mission. There are many
types and configurations of gas-turbine units.
Maintenance performed on APUs can be
almost as extensive as that performed on aircraft
engines. The major difference, other than the size,
is that most APUs use centrifugal flow jet engines
instead of axial flow. The nomenclature of many
of the components may be the same. However,
the component itself may not look or operate in
the same manner because of the basic function
of the unit.
Authorized repairs for organizational activities
include minor component replacements and
adjustments. Common repairs include inlet/
exhaust door actuator maintenance, bleed-air
shutoff valve maintenance, and generator replacement. Maintenance on the ignition system,
generator, oil tank, oil pressure switch, oil cooler,
and scavenge oil filter, or major inspections may
require APU removal because of the APU
location. Major APU inspections and repairs
are performed by the supporting intermediate
maintenance activity (IMA) under the Auxiliary
Power Unit and Support Equipment Gas Turbine
Engine Management Program.
After completing this chapter, you will be able to:
Identify helicopter flight characteristics.
Identify maintenance procedures for
helicopter transmission and rotor systems.
Recognize the purpose and major engine
components of turboshaft engines and
helicopter transmission and rotor systems.
The lift generated by a rotating wing enables
the helicopter to accomplish its unique missions
involving hovering and operating in confined
areas. It also creates some unusual operating and
control problems. Since rotor aerodynamics are
the main difference between helicopters and fixedwing aircraft, let’s first examine the rotor in detail.
We will then look at helicopter controls, types of
helicopters, engines, and finally, transmission and
rotor systems.
The helicopter has become a vital part of naval
aviation. Helicopters have many uses. Some of
these uses are antisubmarine warfare (ASW),
search and rescue, minesweeping, amphibious
warfare, and the transferring of supplies and
personnel between ships. Transferring of supplies
and personnel is made through internal loading
or vertical replenishment (VERTREP). The
advantage the helicopter has over conventional
aircraft is that lift and control are relatively
independent of forward speed. A helicopter can
fly forward, backward, or sideways, or it can
remain in stationary flight above the ground
(hover). Helicopters do not require runways for
takeoffs or landings. The decks of small ships or
open fields provide an adequate landing area.
The rotor is subject to the same physical laws
of aerodynamics and motion that govern fixedwing aircraft flight. However, the manner in
which the rotor is subjected to these laws is much
more complex. Fixed-wing flight characteristics
depend upon the forward aircraft speed and
control surface movements. In a helicopter, the
rotational speed and pitch variations of the rotor
blades determine the flight characteristics. Since
flight is independent of forward speed, a
helicopter is able to move in any direction at a
controlled low speed.
The main difference between a helicopter and
an airplane is the source of lift. The airplane gets
lift from a fixed airfoil surface (wing) while the
helicopter gets lift from a rotating airfoil (rotor).
The word helicopter comes from the Greek words
meaning relating wings. You may find it easier
to understand how a helicopter operates by
imagining the following: Remove the wings from
a conventional aircraft and install them above the
airplane. Rotating the wings causes a low-pressure
area to form on the wings’ upper surfaces and
provides lift. This low-pressure area and resulting
lift is, the same as that formed by fixed wings on
an aircraft.
Rotor lift is explained by either of two
theories. The first theory uses Newton’s law of
momentum. Lift results from accelerating a mass
of air downward. This lift is similar to jet thrust
that develops by accelerating a mass of air out
the exhaust. The second theory is the blade
element theory. The airflow over the rotor blade
acts the same as it does on the wing of a fixedwing aircraft. The simple momentum theory
determines only lift characteristic while the
blade element theory gives both lift and drag
characteristics. This comparison gives us a more
complete picture of all the forces acting on a rotor
The blade element theory divides the blade
into parts (blade elements), as shown in figure 7-1.
Engineers analyze the forces acting on each blade
element. Then the forces of all elements are added
to give the rotor characteristics. Each rotor blade
element has a different velocity, and possibly a
different angle of attack. These differences make
analysis a complicated problem.
If the helicopter hovers in a no-wind condition, the rotors plane of rotation is parallel to the
level ground. This attitude also makes the relative
wind parallel to the ground. The angle of attack
is the same on any blade element throughout the
cycle of rotation. The lifting force is perpendicular
to the plane of rotation.
If the helicopter is rising, there is a component
of velocity parallel to the axis of the rotor. Then
the relative wind is the result of the rotational
velocity and the vertical velocity of the helicopter.
Lift acts perpendicular to the relative wind. The
relative wind is no longer parallel to the plane of
rotation. Lift is not acting perpendicular to the
plane of rotation. The vertical thrust then, or the
force acting to overcome gravity, is slightly less
than the lifting force.
So far the discussion has been about the forces
in the vertical direction. These forces support the
helicopter, but do not give it any horizontal
motion. Rotational and vertical velocities have
already entered the picture. In any discussion of
the principles of flight of the helicopter, the
different velocities being considered must be
specified. This also applies to such factors as
torque, drag, and other forces.
Helicopters are subject to several rotarywing aerodynamic effects. These forces act
independently. Their cumulative sum are factors
that affect helicopter flight.
Although torque is not unique to helicopters,
it does present some special problems. As the main
rotors turn in one direction, the fuselage may
rotate in the opposite direction. Newton’s third
law of motion states that “every action has an
equal and opposite reaction.” This tendency for
the fuselage to rotate is known as torque effect.
Since torque effect on the fuselage is a direct result
of engine power, any change in power changes
the torque. The greater the engine power, the
greater the torque. There is no torque reaction
when an engine is not operating. Therefore, there
is no torque reaction during autorotation.
The usual method of counteracting torque in
a single main rotor is by a tail (antitorque) rotor.
This auxiliary rotor is mounted vertically on the
outer portion of the tail boom. The tail rotor and
its controls serve as a means to counteract torque,
and it provides a means to control directional
Dissymmetry of Lift
Dissymmetry of lift is the lift difference
existing between the advancing blade half of the
disk and the retreating blade half. The disk area
is the area swept by the rotating blades. It is
Figure 7-1.-Blade element theory.
created by horizontal flight or by wind in a hover.
You should be aware that hovering in a 20 mph
headwind is the same as flying forward at 20 mph.
When hovering in a no-wind condition, the speed
of the relative wind is the effective speed of the
rotor. However, the speed is lower at points closer
to the rotor hub, as shown in figure 7-2. When
the helicopter moves forward, relative wind over
each blade becomes a combination of the rotor
speed and forward movement. The advancing
blade is then the combined speed of the blade
speed and helicopter speed. While on the opposite
side, the retreating blade speed is the blade speed
minus the speed of the helicopter. Figure 7-3
shows dissymmetry of lift at 100 mph forward
During forward flight, lift over the advancing blade half of the rotor disk is greater then the
retreating half. This greater lift would cause the
helicopter to roll unless something equalized the
lift. One method of equalizing the lift is through
blade flapping.
Figure 7-3.-Dissymmetry of lift.
Blade Flapping
Blades attached to the rotor hubby horizontal
hinges permit the blade to move vertically. The
blades actually flap up and down as they rotate.
The hinge permits an advancing blade to rise, thus
reducing its effective lift area. It also allows a
retreating blade to settle, thus increasing its
effective lift area. The combination of decreasing
lift on the advancing blade and increasing lift on
the retreating blade equalizes the lift.
Blade flapping creates an unbalanced condition, resulting in vibration. To prevent this
vibration, a drag hinge allows the blades to move
back and forth in a horizontal plane. A main rotor
that permits individual movement of blades in a
vertical and horizontal plane is known as an
articulated rotor.
Coning is the upward bending of the
blades caused by the combined forces of lift
and centrifugal force. Before takeoff, due to
centrifugal force, the blades rotate in a plane
nearly perpendicular to the rotor hub. During a
vertical liftoff, the blades assume a conical path
as a result of centrifugal force acting outward and
lift acting upward.
Coning causes rotor blades to bend up in a
semirigid rotor. In an articulated rotor, the blades
move to an upward angle through movement
about the flapping hinges.
Gyroscopic Precession
The spinning main rotor of a helicopter acts
like a gyroscope. It has the properties of
gyroscopic action, one of which is precession.
Gyroscopic precession is resulting action occurring
Figure 7-2.-Symmetry of lift.
90 degrees from the applied force. Applying a
downward force to the right of the disk area will
cause the rotor to tilt down in front. This
downward tilt is true only for a right-to-left
(counterclockwise) rotor rotation.
The cyclic control applies force to the
main rotor through the swashplate. To simplify
directional control, helicopters use a mechanical
linkage, which places cyclic pitch change 90
degrees ahead of the applied force. Moving the
cyclic control forward (right-to-left turning rotor)
places high pitch on the blades to the pilot’s left.
Low pitch is then found on the blades to the
pilot’s right. Since every pitch change causes a
flap, reaching its maximum at 90 degrees, this
flapping causes the disk area to tilt forward.
If offset linkage were not used, the pilot would
have to move the cyclic stick 90 degrees out of
phase. The pilot would have to move the stick to
the right to tilt the disk forward, and forward to
tilt the disk area to the left, and so on.
Factors that affect rotor blade lift are the rotor
area, pitch of rotor blades, smoothness of rotor
blades, and density altitude.
Rotor Area
One assumption in figuring the lift of a rotor
is that lift is dependent upon the entire area of
the rotor disk. The rotor disk area is the area of
the circle. The radius of the rotor disk is equal
to the length of the rotor blade. The lift of a rotor
increases not in direct proportion to the length
of the rotor, but in proportion to the square of
the length of the rotor. The use of large rotor disk
areas is readily apparent. The greater the rotor
disk area, the greater the drag created. This drag
results in the need for greater power requirements.
Pitch of Rotor Blades
If the rotor operates at zero angle of attack
or zero pitch, no lift would result. When the pitch
increases, the lifting force increases until the angle
of attack reaches the stalling angle. To even out
the lift distribution along the length of the rotor
blade, it is common practice to twist the blade.
Twisting the blade causes a smaller angle of
attack at the tip than at the hub.
Ground Effect
When a helicopter hovers close to the ground,
the rotor directs air downward faster then it can
escape. This builds up a cushion of dense air
beneath the helicopter known as ground cushion
or ground effect. It is effective to a height of onehalf the rotor diameter. Ground cushion effect
does not occur at airspeeds greater then 10 mph.
Smoothness of Rotor Blades
Tests have shown that the lift of a helicopter
increases by polishing the rotor blades to a mirrorlike surface. By making the rotor blades as smooth
as possible, the parasite drag reduces. Any dirt,
grease, or abrasions on the rotor blades may be
a source of increased drag, which will decrease
the lifting power of the helicopter.
Autorotation occurs when the main rotor
turns by air passing up through the rotor system
instead of by the engine. The rotor disengages
automatically from the engine during engine
failure or shutdown. During autorotation, the
rotor blades turn in the same direction as when
engine driven. Air passes up through the rotor
system instead of down. This causes a slightly
greater upward flex or coning of the blades.
Density Altitude
In formulas for lift and drag, the density of the
air is an important factor. The mass or density of
the air reacting in a downward direction causes the
upward force or lift that supports the helicopter.
Density is dependent on two variables. One
variable is the altitude, since density varies from
a maximum at sea level to a minimum at high
altitude. The other variable is atmospheric
changes. The density of the air may be different,
even at the same altitude, because of changes in
temperature, pressure, or humidity.
Power Settling
Stalling, as applied to fixed-wing aircraft, will
not occur in helicopters. However, power settling
may occur in low-speed flight. Power settling is
the uncontrollable loss of altitude. Heavy gross
weights, poor density conditions, and low forward
speed all contribute to power settling. During low
forward speed and high rates of descent, the
downwash from the rotor begins to recirculate.
The downwind flows up, around, and back down
through the effective outer disk area. The recirculating air velocity may become so high that
full collective pitch cannot control the rate of
The mechanical flight control system consists
of mechanical linkage and controls. See figure 7-4.
angle changes result from both collective pitch
stick and rudder pedal movement. The auxiliary
servo irreversible transfer of collective pitch
motion will act to displace the rudder. The rudder
pedal motion will not affect rotary wing collective
pitch blade angle.
The rotary rudder control system controls
helicopter heading by moving control rods and
bell cranks connected to the auxiliary servo
cylinder. The auxiliary servo cylinder connects to
the mixing unit by control rods. At the mixing
unit, a control rod operates a forward quadrant.
From the forward quadrant, cables operate a rear
quadrant in the aft fuselage. A control rod from
the rear quadrant connects, at the pylon hinge
line, to the control rods, bell cranks, and pitch
control shaft. These units are found in the tail
gearbox. A hydraulic pedal damper in the servo
cylinder prevents sudden movements of the
control pedals from causing rapid changes in
blade pitch, which might damage the helicopter.
A typical helicopter flight control system
consists of four subsystems. They are the cyclic
pitch control sticks for directional flight, the
collective pitch control sticks for vertical flight,
the directional heading control (rudder) pedals,
and the throttle. The throttle may be a motorcycle-type grip mounted on the collective pitch
stick. The throttle may be lever-type mounted on
the center overhead console. Also included are an
auxiliary servo cylinder, a mixing unit, primary
servo cylinders, and mechanical linkage.
The cyclic pitch control stick controls forward,
aft, and lateral helicopter movements. The
collective pitch control stick controls vertical
helicopter movement. The directional control
pedals control helicopter headings. See figure 7-4.
Movement of the control sticks is sent by
mechanical linkage to a hydraulically operated
auxiliary servo cylinder for power boost. The
boosted input is sent through a mixing unit for
coordination with any heading directional control
inputs. Stick movement is finally sent through the
hydraulically actuated primary servo cylinders to
the rotary wing head. At the rotary wing head,
blade pitch changes.
These linkages and controls transmit force to
primary and auxiliary hydraulic system servo
cylinders and to the rotary rudder. The
mechanical flight controls have two independent
systems. They are the rotary wing flight control
and the rotary rudder flight control systems.
The rotary wing flight controls consist of cyclic
and collective pitch control sticks, a mixing unit,
a balance spring, and linkage. The cyclic controls
give forward, aft, and lateral movement of the
helicopter. The collective controls give vertical
control of the helicopter.
The rotary rudder flight controls consist of
pedals, pedal switches, and pedal adjusters for the
pilot and copilot. Other system components are
a negative force gradient spring and mechanical
linkage. The rotary rudder compensates for the
torque of the rotary wing. The controls provide
a way to change the heading (direction) of the
The cyclic pitch control system controls the
forward, aft, and lateral movements of the
helicopter. Control comes from the pilot’s or
copilot’s cyclic stick. Control rods and bell cranks
connect the stick to the auxiliary servo cylinders,
then to the mixing unit, and three primary servo
cylinders. The primary servo cylinders control
movement of the rotary wing blades through the
swashplate. The swashplate changes the pitch of
the blades.
The collective pitch control system provides
vertical control of the helicopter. Control comes
from the pilot’s and copilot’s collective pitch
control sticks. Control rods and bell cranks
connect the stick to the auxiliary servo cylinder,
then to the mixing unit. At the mixing unit,
movements of the collective stick are sent to the
primary servo cylinders and the swashplate. The
swashplate increases or decreases the pitch of all
blades equally and simultaneously. A balance
spring on a control rod helps to balance the weight
of the collective stick when the auxiliary servo
system is off.
A collective-to-cyclic pitch coupling (fore-andaft) is found in the mixing unit. The coupling
automatically applies a nose-down pitching correction when the collective pitch stick is raised.
When the collective pitch stick is lowered, a
nose-up pitching correction is made. This action
provides attitude control during transitions,
especially during automatic stabilization equipment (ASE) system transitions.
A collective-to-yaw coupling provides
automatic rotary rudder pitch changes to adjust
for collective pitch changes. Rotary rudder blade
Helicopters are of two basic types. They are
the single-rotor and multi-rotor types. The single
main rotor with a vertical tail rotor is the most
common type of helicopter. The SH-60 or SH-2,
shown in figure 7-5, are examples of single-rotor
to the main transmission gearbox through a coaxial main drive shaft. The main drive on the shaft
is on the rear of the T58 and on the front of the
T700 engines. See figures 7-6 and 7-7.
helicopters. Multi-rotor helicopters are classified
into different categories according to their rotor
configuration. Two different types are the coaxial
rotor and tandem rotor.
The single-rotor configuration requires the use
of a vertical tail rotor to counteract torque and
provide directional control. The advantages of
this configuration are simplicity in design and
effective directional control. Coaxial rotors are
two rotors mounted on the same mast and turning
in opposite directions. The torque produced by
the two rotors balance each other out. Coaxial
rotor systems have good ground clearance and are
easy to maneuver. Their arrangement and controls
are more complicated then single rotor systems.
In the tandem rotor design, one rotor is located
forward and the other located aft. Sometimes the
rotor blades are in the same plane. The blades may
or may not intermesh. The design offers good
longitudinal stability since the fuselage hangs at
two points, fore and aft. Like the coaxial rotor,
the tandem rotor has little torque to overcome
since these rotors rotate in opposite directions.
Most of the Navy helicopters are a twinturbine engine powered, single rotor design like
the SH-60 and SH-3. Some small trainer
helicopters like the TH-57 have only one engine,
and large helicopters like the CH-53 Super Stallion
have three turbine engines. The maintenance
procedures and examples of components used are
those of the representative SH-60 and SH-3
helicopters. Both types are used to compare
similarities and differences between a modular and
nonmodular design.
NOTE: The T700-GE-401 turboshaft
engine is one of a new generation of
aviation propulsion systems with modular
construction. Modular construction allows
intermediate maintenance activities to
replace major engine components, such as
turbine sections, with basic hand tools. The
ability to replace parts at the lowest level
of maintenance (third-degree intermediate)
increases aircraft availability by decreasing down time.
The T58-GE-10 engine is a compact turboshaft
engine with high power-to-weight ratio. The T58
engine consists of two basic sections. These sections are the gas generator and the power turbine
Gas Generator Section
The gas generator section is divided into five
subsections. These subsections are the front
frame, accessories, compressor, combustion, and
gas generator turbine.
FRONT FRAME.— The front frame provides
provisions for engine mounting and supports the
accessory section. It provides for engine air inlet
and anti-icing for the inlet air.
NOTE: While procedures for specific
engines or aircraft are included, many
pertinent or mandatory references are left
out. For this reason always refer to the
applicable maintenance instruction
section consists of the centrifugal fuel purifier,
accessory drive gear casing assembly, lubescavenge pump, fuel pump, main fuel control, and
the pilot valve. It mounts to the bottom of the
front frame section and extends under the
compressor section. Power to drive the accessory
section is received from the gas generator rotor.
The SH-3 uses two T58-GE-10 turboshaft
engines. The SH-60 uses two T700-GE-401
turboshaft engines. See figures 7-6 and 7-7.
Both of these turboshaft engines use the free
turbine principle for power takeoff to the main
transmission gearbox. Power takeoff comes from
the power turbine section. This section is
mechanically independent from the gas generator.
Exhaust gases from the gas generator turbine drive
the power turbine. The power turbine is connected
COMPRESSOR SECTION.— The compressor is a 10-stage axial flow unit that consists of
a rotor and a stator. The inlet guide vanes and
the vanes in stages 1 through 3 are automatically
positioned by the stator vane actuator using fuel
pressure for the motive force.
COMBUSTION SECTION.— The combustion section is a singular annular type of unit. It
free-power turbine. Some features of this engine
include an integral inlet particle separator and
self-contained systems incorporating modular
provides a mounting place for a fuel manifold
block, fuel drain, bleed air ports, and igniter
The two-stage axial flow turbine section drives the
compressor. It consists of the turbine stator and
turbine rotor.
The engine consists of five major sections.
These sections are the inlet, compressor, combustor, turbine, and exhaust sections.
Power Turbine Section
Inlet Section
The power turbine section consists of the
exhaust casing assembly, rotor assembly, and
accessory drive system. The power turbine section
may be positioned on the gas generator section
in two different positions (exhaust pointing to the
right or left), depending on the engine location.
The inlet section includes all the parts forward
of the compressor. It directs airflow into the
compressor and provides a mount for the
accessory gearbox (AGB). The engine inlet uses
an integral particle separator (IPS) to prevent
foreign objects from entering the compressor. The
IPS includes a swirl frame, scroll case, and enginedriven blower. The 12 fixed-swirl vanes impart
rotation to the airflow. Any particles are thrown
into the collection scroll and dumped overboard
by the blower. See figure 7-8.
THE T700-GE-401
The T700-GE-401 is a front-drive turboshaft
engine featuring a single-spool gas generator
section. The engine has a five-stage axial and a
single-stage centrifugal flow compressor. It has
a flow-through annular combustion chamber. The
engine also has a two-stage, axial-flow gas
generator turbine and a two-stage, axial-flow,
Compressor Section
The compressor section increases the mass
airflow delivered to the combustor through a sixstage compressor (five axial, one centrifugal). The
Figure 7-8.-Inlet particle separator (IPS) airflow.
inlet guide vanes and first two rows of stators are
a variable geometry design.
mount between the gas generator and power
Combustor Section
Exhaust Section
The combustor section houses an annular
combustion liner. This section also contains
12 fuel injectors and 2 spark igniters.
The exhaust section is aft of the turbine section
and contains two power turbine speed (Np)
sensors. The exhaust section directs the hot
exhaust gases to the atmosphere.
Turbine Section
The turbine section has four axial flow turbine
wheels. The first two wheels drive the compressor
and AGB. These turbines are known as the
engine’s gas generator turbines. The last two
wheels drive the main gearbox. These turbines
are known as the engine’s power turbines.
The turbine section also houses the seven turbine
gas temperature (TGT) thermocouples, which
Because components vary in function and
complexity on different models of helicopters, we
will discuss only representative units. Refer to the
aircraft MIM for details on components for a
specific helicopter.
Figure 7-9.-Engine control quadrant.
manually controls Ng and Np. In this mode, the
electrical control unit (ECU) is disabled. The only
automatic function NOT de-energized is the Np
overspeed protection. To return to automatic
engine control, move the power lever to the IDLE
position, then back to the FLY position.
An engine control quadrant is shown in
figure 7-9. It consists of two engine power
control levers, two engine fuel system selector
levers, and two engine emergency off T-handles.
It also has a power control lever rotor brake
interlock. Each power control lever has a starter
button and four selectable positions. The positions
The transmission system takes combined
power from two engines, reduces the rpm, and
transfers it to the main and tail rotors. The
secondary function is to provide a drive for
electrical and hydraulic power generation. The
transmission system of a typical helicopter consists
of the main transmission gearbox, an intermediate
gearbox, a tail gearbox, and drive shafts. Most
systems also include an oil cooler, blower, and
rotor brake system. Figure 7-10 shows the SH-3
Movement of the power control lever to the
OFF position moves a cable to shut off the fuel.
Movement of the lever between IDLE and FLY
sets the available gas generator turbine speed (Ng).
Move the lever to the FLY position for flight rotor
speeds. If demanded, this setting gives the highest
available power. When moved to the LOCKOUT
position momentarily, the power control lever
Figure 7-10.-SH-3 power transmission system.
transmission system, and figure 7-11 shows the
SH-60 transmission system.
if one pump fails. The single planetary gear
reduces engine rpm to drive the rotary wing and
accessories. See figure 7-13.
NOTE: Always refer to the applicable
MIM when attempting repair on helicopter
A freewheeling unit, at each engine input to
the main gearbox, permits the rotary wing to
autorotate without engine drag. The action
occurs in case of engine (or engines) failure or
when engine rpm decreases below the equivalent
of rotor rpm. The freewheeling unit also provides
a means of disengaging the rotary wing head while
providing power to operate accessories.
The main gearbox mounts above the cabin aft
of the engines. Its main purpose is to interconnect
the two engines in order to drive the rotary wing.
The main gearbox accessory section is at the
rear of the main gearbox lower housing. See
figure 7-12. It drives the primary, utility, and
auxiliary hydraulic pumps. It drives the highpressure torque meter oil pump, the No. 1 and
No. 2 generators, and the rotary wing tachometergenerator. Dual oil pumps are on the accessory
section. These pumps increase reliability through
better lubrication and permit continued flight
The accessory drive rotor lockout permits the
pilot to use engine power to drive the accessory
section of the main gearbox. This action occurs
on the ground without rotating the rotary wing
head. Power comes from the No. 1 engine only.
A switch allows the pilot to position the linear
actuator and divert power to either the accessory
section or rotary wing. This switch, located on
the forward overhead switch panel, is marked
Figure 7-11.-SH-60 power transmission system.
Figure 7-12.-SH-3 main gearbox.
Figure 7-13.-Main transmission schematic.
of the engines. They each contain an input bevelpinion and gear, and a freewheel unit. The
freewheel unit allows engine disengagement during
autorotation. If an engine fails, the freewheeling
unit allows the main transmission to continue to
drive the accessory unit. The input module
provides the first gear reduction between the
engine and the main module.
The main transmission gearbox drives and
supports the main rotor. See figure 7-14. Notice
that the SH-60 main transmission, like the engine,
is of modular design. The gearbox consists of five
modules. The modules are a main module, two
input modules, and two accessory modules. The
left- and right-hand input and accessory modules
are the same, and they are interchangeable. The
main gearbox is pressure lubricated and uses the
standard oil indicating systems.
Accessory Module
Each accessory module provides mounting
and drive for an ac generator and a hydraulic
pump package. Accessory modules mount on the
forward section of each input module. Identical
and interchangeable sensors are mounted on
them. A rotor speed sensor mounts on the right
module. The left module provides a mount for
a low oil pressure sensor.
Main Module
The main module provides mounting for the
two input and accessory units. It has oil pressure, oil temperature, low oil pressure, high
temperature warning, and chip detector systems
A rotor brake mounted on the tail takeoff
provides the capability to stop the rotor system.
The rotor brake disc also provides the means to
position the main rotor blades for folding.
The intermediate gearbox transmits torque
and changes the angle of drive from the main
transmission gearbox to the tail gearbox. The
intermediate gearbox has an input housing and
gear, a spring-loaded disconnect jaw, a center
housing, and an output housing and gear. The
Input Module
The input modules mount on the right and left
front of the main module and support the front
Figure 7-14.-SH-60 main transmission gearbox.
cooled. See figure 7-16. An electromagnetic chip
detector/drain plug is in the bottom of the input
housing. It also has an oil level sight gauge and
a filler plug. Access to the tail gearbox is through
the tail gearbox access fairing at the top of the
disconnect jaw permits folding and unfolding of
the pylon without manually disconnecting the
transmission. A transmission lock prevents the
rotary rudder from windmilling when the pylon
is folded. The lock is automatically unlocked
when the pylon is unfolded (flight position). The
input and output housings contain similar bevel
gears to change the angle of drive up along the
pylon. The center housing has an oil level sight
gauge. It has a chip detector/drain plug on
the bottom, and the filler plug at the top. See
figure 7-15. The electromagnetic chip detector, in
addition to normal chip detector function, also
detects overheating of the intermediate gearbox.
The gearbox is splash-lubricated and air cooled.
The tail gearbox, mounted at the top of the
tail pylon, supports and drives the rotary rudder.
It reduces shaft speed and changes the direction
of drive by 90 degrees. A pitch change shaft,
controlled by the rudder flight controls, operates
a pitch change beam. Linked to the sleeve of
each rotary rudder blade is a pitch change
beam. The gearbox is splash-lubricated and air
The main drive shaft is a dynamically balanced
shaft that transmits torque from the power turbine
to the main gearbox. The SH-3 drive shaft
assembly consists of a high-speed shaft, Thomas
coupling, engine adapter, and match-marked nuts
for engine attachment. The shaft assembly has
match-marked bolts and nuts for attachment to
the main gearbox coupling. The Thomas coupling
and engine adapter secure to the shaft by special
bolts and hi-lock collars to prevent disassembly.
The shaft is made of tubular steel, and it is about
13 1/2 inches long. It is flanged at both ends for
connection to the Thomas coupling, adapter, and
power turbine coupling flange, and to the main
gearbox coupling flange. The drive shaft is within
and protected by the aft engine support. The
direction of shaft rotation is clockwise (when
Figure 7-15.-Intermediate gearbox.
Figure 7-16.-Tail gearbox.
viewed from the aft end of the engine). To remove
the main drive shafts, both remountable power
plants must be removed.
The tail drive shaft runs from the rear cover
of the main gearbox to the disconnect coupling
at the intermediate gearbox. See figures 7-10 and
7-11. The shaft between the intermediate gearbox
and the tail gearbox is known as the pylon drive
The primary purpose of the tail drive shaft is
to send engine power to drive the rotary rudder.
On some installations it also provides a means to
drive the main transmission gearbox oil cooler and
blower fan. On the SH-60, there are five sections
of drive shafting from the main transmission to
the intermediate gearbox. There is one section of
tail drive shaft between the intermediate and tail
gearbox—making a total of six sections. Each
section connects by a flexible steel coupling called
a Thomas coupling, eliminating the need for
universal joints. Each coupling has flexible
stainless steel discs stacked together. Flats assure
that the stack is in correct alignment. The grain
in one disc runs parallel to the flats, and the grain
in the other disc runs perpendicular to the flats.
The shaft sections are supported by viscousdamped bearings. Each viscous-damped bearing
support is a ball bearing enclosed by a thick
rubber-type bag. Heavy silicone oil in the bag
dampens vibrations in the tail drive shaft
assembly. The pulley and belts that drive the oil
cooler fan are attached to the oil cooler drive
shaft. These are found between sections I and II
of the tail drive shaft. See figure 7-11. The
section VI drive shaft is in the tail pylon. It sends
power from the intermediate gearbox to the tail
The rotor brake system permits applying the
main rotor brake manually or automatically.
It uses a master brake cylinder, pressure
gauge, panel package, rotor brake, rotor brake
accumulator, check valves, and pressure switches.
Operation of the system is in conjunction with the
operation of the automatic blade folding system.
The rotor brake at the rear of the main gearbox
is hydraulically actuated. Its purpose is to stop
the rotation of the rotary wing head and rotary
rudder. See figure 7-14. Actuation is manual by
means of the rotor brake master cylinder in the
cockpit. Operation is automatic during blade
folding by the blade positioner control valve. The
rotor brake consists of a rotor brake disc and
When performing inspections, removal,
repairs, or installation maintenance functions on
parts, you must follow procedures specified in the
appropriate MIMs.
Besides inspecting for corrosion, and treating
the main transmission outer surface for corrosion,
there are some areas that require more attention.
Visually inspect these areas for signs of overstress
or beyond torque limit capabilities. The main
areas are the barrel nuts, the forward bell crank
support bridge mounting pad, and the main
module mounting feet.
Barrel Nut Inspection
On the H-60 helicopter there are three
different part numbers for barrel nuts. The
procedure for checking each one of them is the
same. Install a mount bolt onto the barrel nut
until there are two threads exposed beyond the
nut. Using a torque wrench, back out the bolt.
If the breakaway torque is less than that specified
in the MIM for that specific part number nut,
discard the nut. Replace the nut with a new one,
and repeat the procedure.
The main gearbox oil cooler and blower
unit of the SH-3 helicopter is in the aft rotary
wing fairing. It consists of a cooler (radiator),
blower, and duct. The cooler is belt-driven by
the tail drive shaft. If the temperature of the
oil is less than 70°C, the oil is bypassed
to the return line by a thermostatic regulator.
Oil returning from the radiator or the bypass is
forced through the lubricating jets located in the
Bell Crank Support Mounting Pad
Following all safety procedures, clean all
traces of paint and sealing compound from the
mounting pad with dry-cleaning solvent. Repair
minor nicks, gouges, scratches, and corrosion
pitting on the surface of the mounting pad.
Measure the depth of damage to make sure that
blending out the damage does not go deeper than
allowable repair limits. Remove corrosion by light
sanding with aluminum oxide abrasive cloth.
Inspect all blended areas using fluorescent
penetrant. Crack indications or damage greater
than specified requires replacement of the main
The main transmission installation procedures
follow the removal procedures in reverse. There
are some precautions to follow during the installation. To prevent damage, carefully guide the main
module into place to clear all parts on the
helicopter. Apply antiseize compound to the
mounting bolts.
Antiseize compound may contain lead. Do
not smoke, drink, or eat when handling it.
Wash thoroughly after use. If accidently
swallowed, do not induce vomiting. Seek
medical attention. Sealing compounds and
adhesives are toxic. Use rubber or
polyethylene gloves and goggles. Wash
hands thoroughly with soap and water
before eating or smoking. Avoid breathing
vapors during mixing, lay-up, or curing.
Avoid breathing dust from sanding or
Main Module Mounting Feet
Visually inspect the main module mounts daily
for cracks radiating outward from the center of
the bushings. Check for cracks extending down
from the sides of the mounting feet. If no cracks
are found either in the mounting feet or the sealing compound, further inspection is not required.
Crack indications will require replacement of the
main module. Repair minor damage as long as
the repaired area stays within the specified
repairable limits.
To prevent corrosion damage to locator pins
and supports on the main module, seal pins and
supports from moisture accumulation. Lower the
main module into position on the helicopter
mounting surface. To ease installation, first line
up the front mounting holes and insert the mount
bolts. Next, line up the rear mounting holes and
insert the bolts. Follow the remainder of the
installation procedures as specified in the MIMs.
To make the main transmission accessible for
removal, the removal of other items is necessary.
First, remove rotor blades and head, swashplate,
and engine air inlets. Next, remove the left and
right input modules with accessory modules
attached. Remove all electrical connections and
harnesses, and all oil and hydraulic lines.
Disconnect and support the tail rotor drive shaft.
To prevent damage to the tail rotor and drive
shafts while the tail drive is disconnected, make
sure the tail rotor does not rotate. Remove fore
and aft bell crank support mounting nuts, bolts,
and washers. Install a lifting eye on the main
rotor shaft. Remove main module mount bolts,
and raise the main module clear of the helicopter. Lower the module onto an adapter and
secure with the appropriate bolts, washers, and
Inspect the shaft for scratches. Blend out
scratches with crocus cloth. Replace the shaft
section if the damaged or blended area exceeds
specified limits. Inspect flexible couplings for
nicks or dents. Blend out dents and nicks on the
edges of individual disc, and replace the coupling
if the damaged or blended area exceeds specified
limits. Inspect each flexible coupling for disc
separation. Each disc assembly may have disc
separation or buckling, provided there are no
kinks, sharp bends, or cracks. Replace the entire
coupling if specified limits are exceeded. Inspect
the plates in the coupling washer area for wear.
Replace the entire coupling if any wear is found.
Inspect the couplings for cracks in the disc. Check
areas close to bolts and special washers. Replace
the coupling if cracks are found.
NOTE: The main transmission main
module weighs about 750 pounds. With all
equipment installed, it weighs 1,200
pounds. The main transmission with the
rotor head and equipment installed weighs
about 2,749 pounds. Be sure to use a hoist
with suitable weight capacity.
Remove paint, dirt, and grease from the
damaged area. Use a cloth moistened with methyl
Figure 7-17.-Rotary wing head assemblies; (A) Self-lubricating head; (B) Grease-lubricated head.
ethyl ketone. Protect bearings and Thomas
couplings from solvent and dirt. Blend out the
damaged area using crocus cloth. The blend radius
must not exceed specified limits, Wipe the area
with a clean cloth moistened with methyl ethyl
ketone. Prime the area with zinc chromate primer,
and touchup with matching paint.
NOTE: Limit your repair of the drive shaft
to the outside diameter of the tubular
shaft. Repair must be done in the scratched
area only.
The rotor system consist of the main rotor
head, main rotor blades, tail rotor head, and tail
rotor blades. This section on rotor systems
discusses the general components and identifies
some of the different features or materials being
The fully articulated rotary wing head is
splined to the rotary wing shaft of the main gearbox. The head is supported by the rotary wing
shaft. The head supports the rotary wing blades,
and is rotated by torque from the main gearbox.
The head transmits movement of the flight
controls to the blades. Its design permits
automatic folding of the blades from a control
panel in the cockpit.
There are two types of rotary wing heads in
use. They are the grease-lubricated and selflubricating heads. See figure 7-17. The greaselubricated rotary wing head contains grease
fittings for lubrication. The self-lubricated rotary
wing head has fewer grease fittings, because oil
reservoirs (tanks) are used on the hub plate at the
vertical hinges. These reservoirs lubricate upper
and lower hinge bearings and the stack bearings
in sleeve spindles.
Principal parts of the head are the hub and
swashplate. The hub consists of a hub plate, lower
plate, and hinges between each arm of the plates.
The hub contains sleeve spindles that are attached
to the hinges and damper-positioners. The
swashplate consists of a rotating swashplate and
a stationary swashplate. The rotary wing head
has antiflapping restrainers, droop restrainers,
adjustable pitch control rods, and the rotating and
stationary scissors. The swashplate and adjustable
pitch control rods permit movement of the flight
controls to be transmitted to the rotary wing
blades. The hinges allow each blade to lead, lag,
and flap. The damper-positioners restrict lead and
lag motion and position the blades in preparation
for folding. Sleeve spindles allow each blade
rotation on its spanwise axis to change blade pitch.
The antiflapping restrainers and droop restrainers
restrict flapping motion when the rotary wing
head is slowing or stopped. The control lock
cylinder unlocks to permit the sleeve to turn about
the sleeve spindle axis. The stationary scissors are
connected to the stationary swashplate and main
gearbox upper housing. The rotating scissors are
connected to the rotating swashplate and lower
hub plate.
The rotary wing blades provide the lift
necessary for flight. See figure 7-18. The blades
are of the nitrogen-pressurized spar type. The
Figure 7-18 .-Rotary wing blade.
black that shows depends on pressure in the
spar. Remove from service any blade on
which the pressure indicator shows any black
color. The blade may be put back into service
when the unsafe (black) indication is found
and corrected. Replace a malfunctioning indicator, but only if the spar pressure is within
permissible limits.
blade has an air valve in the spar back wall near
the root end and a cylindrical pressure plate. The
root end plate is attached to the inboard end of
the spar. A seal plate is found inside the spar tip
end. Both are sealed for pressurization. Pressure
loss in the spar shows impaired integrity of the
spar or a seal leak. The cuff provides the means
for attaching the blades to the rotary wing head
sleeve spindles. Nickel-plated or titanium abrasion
strips bonded to the spar leading edges prevent
Older blades consisted of a hollow extruded
aluminum spar and aluminum pockets. They have
a tip cap, a root cap, and a steel cuff. Newer
blades consist of a pressurized titanium spar,
honeycomb core, and fiber glass graphite
skin. The newer blades are often repairable at
organizational-level maintenance instead of depot
level. They are repairable at the lower level of
repair because of their honeycomb and fiberglass
design. Both types of blades are statically and
dynamically balanced to permit individual replacement and interchangeability of the blades. In
addition to balancing, manufacturers and depot
repair facilities stencil blades with a pretrack
number to aid in blade tracking.
The rotary rudder head provides for attachment of the rotary rudder blades and counteracts
the torque of the main rotor head. See figure 7-20.
It also serves as a rudder for directional control
of the helicopter. The rotary rudder head
is driven by the tail gearbox. Blade pitch
changes by the action of the pitch change
shaft. The pitch change shaft moves through
the center of the output gear shaft of the
tail gearbox. As the shaft moves outward
from the gearbox, the pitch of the blades
decrease. The pitch beam is connected by
adjustable pitch change links to the forked
brackets of the blade sleeves. The flapping
spindles permit flapping of the blades in each
The pressure indicator, usually known as
a BIM or a blade inspection method, compares
built-in reference pressure with blade spar
pressure. See figure 7-19. When pressure
in the spar is within the required service
limits, three white stripes show in the indicator.
If pressure in the spar drops below the minimum
permissible service pressure, the indicator will
show three black stripes. The amount of
Tail rotor blades are built around a spar
that mounts on the tail rotor. The SH-3 has
five all-metal, single aluminum pocket blades
bonded to a C-shaped spar. The SH-60 blades
are built around two graphite composite spars.
Honeycomb paddles are then bonded to the
spars with fiber glass to form the blade.
Figure 7-19.-Blade pressure indicator (BIM).
Organizational maintenance of the helicopter
rotor system includes periodic inspection, lubrication, rigging, and adjustment. Organization
maintenance includes the cleaning of the rotary
wing and rudder blades, and the removal and
replacement of malfunctioning components.
Vibration of the rotary mechanisms can result
in work hardening of metals and later fatigue
failure. Nondestructive testing of special parts of
the rotary wing and the rotary rudder at specified
intervals is necessary to prevent failures. The
rotary wing head and rotary rudder assemblies are
high-time removal items, as listed in the periodic
maintenance information cards (PMIC).
Cleaning of the rotary wing and rotary rudder
should be accomplished as necessary, using
approved cleaners mixed with water. The concentration of the mixture will vary, depending on the
surface condition.
Figure 7-20.-Rotary rudder head.
is necessary when the blades, the main gearbox,
or the main rotor head assembly have been
replaced. Unless the blades are in proper track,
vibrations will occur in the helicopter with every
revolution of the main rotor. At high rpm settings,
these vibrations could cause serious structural
Both the rotary wing and rudder blades
have areas that are joined by bonding
adhesives. Never use solvents or cleaners
not specifically authorized in the MIM.
Never use lacquer thinner, naphtha, carbon
tetrachloride, or other organic compounds
for cleaning in these bonded areas. The
bonding will be weakened by the solvent
and may result in blade failure.
Tracking the blades is necessary to be sure that
all of the blades rotate in the same horizontal
plane (track). This is accomplished by pretrack
rigging of the rotary wing head and by the use
of pretracked blades.
Although ADs remove and replace rotary
wing parts, the airframes work center normally
performs the rigging checks. Rigging checks and
adjustment involve coordinating the cyclic pitch
control stick, collective pitch control stick, and
pedal positions with the correct rotary wing and
rotary rudder blade angles. Rigging checks are
necessary to ensure that the flight controls are
operating under normal friction loads. At the
completion of rigging, a flight test is performed
by a qualified pilot. This includes a check of blade
Pretrack rigging of the rotary wing head
involves adjusting the pitch control rods until
an exact sleeve angle (within 1 minute) is
attained on all sleeve spindles. When this exact
angle is established, a micrometer type decal
is affixed on the adjustable pitch control
rods. The decal becomes a permanent reference
at the overhaul activity. A pretrack number
is stenciled on each blade at manufacture or
overhaul, based on the effective angle of the
blade. Install any pretracked blade on the
helicopter simply by setting the adjustable
pitch control rod to the pretrack number stenciled
on the blade. The blade tracking is then checked
with either Strobex or electronic tracking equipment.
You should perform blade tracking whenever
the helicopter has been rerigged. Blade tracking
A Strobex blade tracker permits blade tracking from inside the helicopter in flight or on the
ground. See figure 7-21. The system uses a highly
concentrated stroboscopic light beam flashing in
sequence with the rotation of the rotary wing
blades. When aimed at a fixed target on the blade
tips, the tips appear to stop.
To synchronize the strobe light and rotary
wing rotation, a soft iron sweep attached to the
rotating swashplate passes close to a magnetic
pickup on the stationary swashplate. This causes
a once-per-revolution pulse used for synchronization. Each blade has a retroreflective target
number attached under the blade in a uniform
location. Blade tracking is then determined by the
relative vertical position of the fixed target
Figure 7-22.-Electronic tracking.
Figure 7-21.-Strobex blade tracker.
connecting cables. This assembly provides
continuous meter readings automatically on outof-track conditions. See figure 7-22. The meter
readings are an average of many electro-optical
samplings of each blade, thus eliminating erratic
readings caused by wind gusts.
The electronic blade tracker has a electrooptical pickup (scanner), an electronic conversion
unit (computer), and a magnetic phase detector
and pulse bracket. The tracker has sun shields and
After completing this chapter, you will be able to:
Recognize the principles of turboprop
operation and identify the major components of the turboprop engine.
Recognize power lever movement through the
alpha and beta ranges, and the effects of those
movements on the propeller.
Identify the propeller designation system,
operating principles, and basic propeller parts.
Recognize the procedures and cautions used
in the removal, cleaning, and reinstallation of
the propeller assembly and subassemblies.
Recognize propeller operating principles.
more complicated and heavier than a turbojet
engine of equal size and power. The turboprop
delivers more thrust at low subsonic airspeeds.
This advantage decreases as flight speed increases.
In normal cruising speed ranges, the propulsive
efficiency of a turboprop decreases as speed
increases. In a turbojet, the propulsive efficiency
increases as speed increases. The ability of a
propeller to accelerate a large mass of air at low
airspeed results in the unusual high performance
of a turboprop during takeoff and climb. This
low-speed performance also enables a turboprop
aircraft to carry heavier payloads, making them
ideal cargo aircraft. At about Mach 1 airspeed,
the turboprop engine can deliver more thrust than
the turbojet engine of the same gas turbine design.
For a given amount of thrust, the turboprop
engine requires a smaller gas turbine with lower
fuel consumption than the turbojet engine.
The turboprop engine produces thrust indirectly through the propeller. A characteristic of
the turboprop is that changes in power do not
change engine speed. Changes in power change
the turbine inlet temperature (TIT). During flight,
the propeller maintains a constant 100-percent
engine speed. This speed is the design speed where
power and maximum efficiency is obtained.
Changes in fuel flow affect power changes. An
increase in fuel flow causes an increase in turbine
There are a variety of turboprop aircraft in
the Navy inventory. The C-130 Hercules, a cargo
transport aircraft, is the workhorse of naval
aviation. The E-2 Hawkeye is the fleet’s airborne
early warning aircraft. The C-2 Greyhound is a
fleet logistics support aircraft. The P-3 Orion is
our fleet antisubmarine warfare (ASW) aircraft.
See figure 8-1.
In this chapter, the T56 engine and the
54H60-77 model propeller are examples of a
common turboshaft engine and propeller system.
There are differences in the turboprop aircraft
mentioned above, but the basic operation,
assemblies, and maintenance are similar.
The turboprop engine section of this chapter
discusses the operating principles, parts, and
systems unique to turboprop engines. After
learning about the turboprop engine, we will
discuss propellers. The propeller section describes
basic propeller parts, operating principles, and
maintenance procedures.
If the exhaust gases from the basic part of a
turbojet rotates an additional turbine that drives
a propeller through a speed-reducing system, it
is a turboprop engine. The aircraft turboprop is
Figure 8-1.-Four types of turboprop aircraft.
These changes occur through coordination
between the propeller governor and the turboprop
engine fuel control. Together they establish the
correct combination of rpm, fuel flow, and
propeller blade angle to create the propeller thrust
required to provide the requested power.
inlet temperature and a corresponding increase in
energy available at the turbine. The turbine
absorbs more energy and sends it to the propeller
in the form of torque. The propeller, in order to
absorb the increased torque, increases blade angle
to maintain constant engine rpm.
Power Section Assembly
The turboprop engine consists of three major
assemblies. They are the power section assembly,
the torquemeter assembly, and the reduction gear
assembly. See figure 8-2. We will discuss the
power section assembly first.
The power section assembly is essentially a
constant-speed turbojet engine. See figure 8-3.
The major assembly consists of an axial flow
compressor assembly, a can-annular combustion
section, a turbine assembly, and an accessory drive
Figure 8-2.-Turboprop engine major assemblies.
Figure 8-3.-Power section assembly.
The pickup assembly consists of electromagnetic pickups mounted radially over the teeth
of the torque and reference shaft flanges. These
pickups produce electrical impulses at the passage
of each exciter tooth. The pickups are displaced
so that the reference flange impulse from its
pickup and the torque flange impulse from its
pickup are slightly out of phase at zero load.
Because zero torque indications are not at the
electrical zero of the indicator, both positive and
negative torque conditions are measured.
housing. The power section assembly contains oil,
fuel, ignition, control, and cooling air systems.
It also has a compressor extension shaft to which
the torquemeter attaches.
Torquemeter Assembly
The torquemeter assembly is located between
the power section and reduction gear box
assemblies. Its purpose is to transmit and measure
the shaft output from the power section to the
reduction gearbox assembly.
The torquemeter operates on the principle of
accurate measurement of torsional deflection
(twist) that occurs in any power transmitting shaft.
This torsional deflection is detected by magnetic
pickups. The deflection is measured electronically,
and displayed on the cockpit instrument panel in
terms of inch-pounds of torque, or shaft
horsepower (SHP). The principle parts of the
torquemeter assembly are shown in figure 8-4.
Two concentric shafts make up the torquemeter assembly. The inner shaft (torque shaft)
carries the load and produces the measured twist.
The outer shaft (reference shaft for measuring
purposes) is rigidly connected to the torque shaft
at the drive input end only. There are separate
flanges on both the torque and reference shafts
at the reduction gear assembly end. Rectangular
exciter “teeth” are machined in line on each
flange, which enable the pickups to detect the
relative displacement of the two flanges.
The torquemeter housing serves as a rigid
lower support between the power unit and the
reduction gear assembly. It provides a mounting
for the pickup assembly at the reduction gear end.
Reduction Gear Assembly
The reduction gear assembly changes the high
rpm, low torque of the turbine section to low rpm,
high torque necessary for efficient propeller
operation. This change is made through a twostage reduction system of sun and planetary gears.
See figure 8-5. The two stages of reduction
provide an overall speed reduction of 13.54 to 1;
for example, when power section rpm is 13,820,
the propeller shaft rpm is 1,020. The reduction
gear case also provides the drive and location to
mount the propeller and accessories. Accessories
mounted on the case include a starter, generator,
engine-driven compressor (EDC), oil pump, and
tachometer generator. The reduction gearbox
assembly also uses safety systems that we will
discuss next.
The complexity of the turboprop configuration brought about the possibility of certain
hazardous in-flight situations. Safety features
have been designed into the system to activate
Figure 8-4.-Torquemeter assembly.
Figure 8-5.-Reduction gear and torquemeter assemblies.
The thrust sensitive signal system is a safety
device used during takeoffs. The TSS automagically initiates propeller feathering and shuts
down the turboprop engine in case of power loss.
This allows the pilot to concentrate on flying the
aircraft during the critical takeoff period. Feathering the propeller reduces the yawing action (caused
by drag) and asymmetric flight characteristics on
multiengined aircraft. The TSS system is on the
reduction gearbox (RGB), and it is armed through
a switch in the flight station.
torque conditions. When a predetermined
negative torque is applied to the reduction gearbox, a stationary (nonrotating) ring gear moves
forward against spring force. This action results
in a rod moving forward through the reduction
gear nose case. The rod positions the feather valve
in the valve housing to an increased blade angle.
When the propeller blade angle has increased
enough to relieve negative torque conditions, the
plunger retracts and the propeller returns to
normal operation.
The NTS system functions in flight during
temporary fuel interruptions, air gust loads on
the propeller, normal descents with lean fuel
scheduling, or high compressor bleed-air conditions at low-power settings.
Negative Torque Signal (NTS)
Safety Coupling
The negative torque signal system momentarily prevents the propeller from driving the
engine during in-flight conditions. The NTS
system is mechanically locked out during engine
operation in the ground range. The system’s
mechanical linkage and plunger are inside the
front case of the reduction gearbox. They work
with the propeller valve housing assembly to
increase propeller blade angle during negative
The safety coupling is an automatic mechanical device that decouples (disconnects) the power
section from the reduction gearbox assembly when
negative torque exceeds the setting of the safety
coupling. The safety coupling is between the main
input pinion gear shaft on the reduction gearbox
and the outer member of the torquemeter drive
shaft. Any transfer of torque, positive or negative,
between the power section and the reduction gear
automatically whenever a system-related hazard
occurs. The following text discusses some of the
hazards and their related safety features.
Thrust Sensitive Signal (TSS)
assembly transmits through the safety coupling.
Positive torque occurs when the power section
drives the propeller through the reduction gear
assembly. Negative torque occurs when the
propeller drives the power section. The safety
coupling backs up the NTS system to prevent
engine compressor and turbine damage. If the
NTS system fails to limit negative torque, the
safety coupling protects the engine from extensive damage by decoupling.
train. The propeller brake has three positions.
They are the released, applied, and locked
The propeller brake is a friction-cone type,
consisting of a stationary inner member and a
rotating outer member. During normal engine
operation, reduction gear oil pressure holds
the brake in the released position. This is
accomplished by oil pressure, which holds the
outer member away from the inner member.
When the engine is shut down, reduction gear oil
pressure drops. A spring force moves the outer
member into contact with the inner member or
to the applied position. The propeller brake locks
when it is moved in a direction opposite normal
rotation. When locked, it acts upon the reduction
gearbox primary stage reduction gearing to
prevent movement along with the friction-cone
Propeller Brake
The propeller brake is used when a turboprop
engine is not in operation. Since the compressor
and turbine rotors can rotate easily when the
engine is shut down, a propeller brake is needed.
The propeller brake prevents the propeller from
windmilling on the ground or when it is feathered
in flight. It also decreases the time for the
propeller to come to a complete stop after engine
shutdown. The brake is in the reduction gearbox
assembly between the rear case and rear case inner
diaphragm. It is part of the accessory drive gear
The control of a turboprop engine involves the
control of engine speed, temperature, and torque.
Figure 8-6.-Control schematic.
The turboprop fuel control and the propeller
governor are connected and operate in coordination with each other. Together they establish the
correct combination of rpm, fuel flow, and
propeller blade angle to create sufficient propeller
thrust to provide the desired power.
Manual control of the system is provided by
the power levers and the emergency shutdown
handles mounted in the flight station. See
figure 8-6. The control systems are divided into
two operational ranges. They are the flight control
range (alpha) and ground handling range (beta).
For airborne operation, the propeller blade
angle and fuel flow for any given power lever
setting are governed automatically according to
a predetermined schedule. Below the flight idle
power lever position, the coordinated rpm blade
angle schedule becomes incapable of handling the
engine efficiently. Here the ground handling or
beta range is encountered. In this range of the
throttle quadrant, the propeller blade angle
is controlled by the power lever position. Next,
we will discuss the engine’s control system
components—power levers, fuel control, and
Power Levers
The power lever controls power delivered by
the engine. There are six positions marked on the
power lever quadrant. See figure 8-7. Listed below
Figure 8-7.-Coordinator quadrant markings.
4. Control engine
the limits of
governor. This
speed ground
speed ground
are the corresponding degrees as read from the
engine coordinator.
Max reverse — 0°
and propeller speed outside
operation of the propeller
includes reverse thrust, lowidle, flight idle, and highidle.
Ground idle — 9°
5. Provide a measure of engine protecion during overspeed conditions by reducing fuel
flow and turbine inlet temperature.
Flight idle — 34°
Air start
— 90° (max power)
6. Provide a starting fuel flow schedule
that, with the temperature datum valve,
avoids overtemperature and compressor
7. Compensate for changes in air density due
to variations in compressor inlet air
temperature and pressure.
The two ranges of operation of the power lever
quadrant are alpha and beta ranges. The beta
range is the portion of the quadrant from max
reverse (0°, to flight idle (34°). With the power
lever set at ground idle, the propeller is at a
5-degree negative blade angle to compensate for
power section thrust. Movement of the power
lever toward flight idle increases fuel flow and
blade angle. Movement from ground idle towards
max reverse increases fuel flow, but causes the
propeller to schedule reverse blade angle. In the
flight range (alpha), the power lever schedules fuel
flow only.
8. Provide a means of cutting off fuel flow
electrically and manually.
The fuel control senses compressor inlet pressure, compressor inlet temperature,
and engine speed. Using these three factors
and the setting of the power lever, the fuel
control meters the proper amount of fuel.
Pressure and temperature compensating systems
are designed to maintain constant turbine inlet
temperature as the compressor inlet conditions
Fuel Control
The fuel control includes a cutoff valve
for stopping fuel flow to the engine. This
valve operates both electrically and manually.
Electrical control is made possible by moving
the fuel and ignition switch to the OFF position.
This signals the cutoff valve motor to close
the valve. Actuation of the emergency shutdown
control closes the valve mechanically through a
cable and control rod system from the flight
A turboprop fuel control is similar to a
turbojet fuel control. The difference is the
turboprop fuel control operates in conjunction
with a propeller governor. The fuel control
mounts on the engine accessory drive housing. It
is mechanically linked to the coordinator. The fuel
control is designed to perform the following
1. Provide a means of varying fuel flow
to permit a selection of power that is
coordinated with propeller blade angle and
engine speed.
Mounted on the rear face of the fuel control
is the engine coordinator. The function of
the engine coordinator is to coordinate the
propeller, temp control, and fuel control. A
system of levers, bell cranks, and linkages
connect the coordinator with the fuel control and
the propeller control. Refer to the system
schematic in figure 8-7.
2. Regulate the rate of fuel metering during
acceleration to prevent excessive turbine inlet temperature.
3. Control the rate of decrease of fuel metering during deceleration to prevent flameout.
The propeller converts the power output of the
engine into forward thrust to move the aircraft
through the air. A propeller is essentially a
“rotating wing, ” or airfoil. When the aircraft
engine turns the propeller, relative motion is
developed between the wing-like propeller blades
and the air. As it pulls itself through the air, the
propeller carries along anything that is attached
to it, within the limitations of the power
developed. The faster the propeller spins, within
certain limits, the greater the resulting pull or
There are different propeller manufacturers
and many varied designs. These designs include
the experimental multicurved blade for propellers.
All propellers have the same basic parts. Terms
for the parts of one propeller are applicable to
other propellers. The basic parts of a propeller
are as follows:
1. BLADE. One arm of a propeller from the
butt to the tip. Propellers usually have two or
more blades. See figure 8-8.
2. BLADE BACK. The surface of the blade
as seen by standing in front of the propeller.
See figures 8-8 and 8-9.
3. BLADE FACE. The surface of the blade
as seen by standing directly behind the propeller.
See figure 8-9.
Figure 8-9.-A four-blade propeller.
4. SHANK. The thickened portion of the
blade near the hub of the propeller. The shank
is sometimes referred to as the root. See figure 8-8.
5. TIP. The portion of the blade furtherest
from the hub. See figure 8-8.
6. HUB. The central portion of the propeller
that is fitted to the propeller shaft, securing the
blades by their roots. See figure 8-8.
7. LEADING EDGE. The forward or
“cutting edge” of the blade that leads in the
direction the propeller is turning. The other edge
(rear edge) is called the TRAILING EDGE. See
figure 8-8.
nut that locks the propeller hub to the propeller
shaft. It is part of the propeller rather than the
9. BLADE STATIONS. These are reference
lines, usually designed as measurements, made
from the hub. These lines are numbered and locate
positions on the propeller blade. They are usually
designated at 6-inch intervals. The first station is
Figure 8-8.-Propeller basic parts.
normally 12 inches from the hub. Figure 8-10
shows the blade stations of a propeller blade.
10. BLADE ANGLE (PITCH). Blade angle
is the angle formed by the chord of a section of
the blade and a plane perpendicular to the axis
of rotation. The blade angles in figure 8-11 are
representative of standard low- and high-pitch,
as well as the feather angle. These angles will vary
with different propeller installations.
11. BLADE CHORD. Blade chord is the
distance between the leading and trailing edges.
This is an imaginary line extending from the center
of the leading edge to the center of the trailing
edge. It is important for blade balancing.
12. FEATHERING. Feathering is streamlining the propeller blade with the relative wind.
This feature is found in most multiengined
propeller installations. Feathering serves to reduce
the drag caused by a windmilling propeller on a
dead engine and to stop rotation that could cause
further damage. See figure 8-11.
13. REVERSING. A reversing propeller
allows for a negative blade angle. With a negative
Figure 8-11.-Propeller blade angles.
blade angle, a propeller produces reverse thrust
(thrust in a direction opposite to that normally
produced by the propeller in flight). Reverse thrust
produces a braking action used during landing to
reduce the landing roll-out distance. It can also
be used to taxi an aircraft backwards.
The model designation for the propeller
assembly, shown by markings on the barrel,
identifies the type of propeller. The number and
letter group describes the basic model, and the
number group that follows the dash indicates the
number of minor modifications.
A breakdown of the designation of the
54H60-77 propeller is as follows:
5—Indicates the number of major changes
incorporated in the propeller.
4—Indicates the number of blades.
H—Describes the blade shank size. (The use
of a LETTER here also indicates that the
blades are made of aluminum. A NUMBER here would indicate the shank size,
and also that the blades were made of
steel, as in model 24260.)
60—Indicates the spline size of the propeller
77—Indicates minor modifications made to
the propeller.
Figure 8-10.—Blade stations.
the engine load and the propeller takes a larger
bite of air. This function will keep the turboprop
engine at 100-percent rpm.
On constant-speed propellers, the blade angle
must be adjusted to provide the most efficient
angle of attack at all engine and aircraft speeds.
The most efficient angle of attack is very small;
it varies from 1 to 4 degrees positive angle. The
actual blade angle necessary to maintain this small
angle of attack varies with the forward speed of
the aircraft. With constantly increasing aircraft
speeds and high-altitude operations, it is necessary
to have a wide range of blade angle settings. This
range of settings must adapt the propeller to
conditions encountered in takeoff, climb, and
The first propellers were fixed-pitch and
designed mainly to get the aircraft off the ground.
The pitch (blade angle) was small so that the
engine could quickly turn over to its full rpm and
use its full horsepower for takeoff. Once an aircraft with a fixed-pitch propeller of low blade
angle is in the air, forward speed of the aircraft
is limited. The low blade angle allows the propeller
to turn too fast to take a big enough bite of the
onrushing air. As a result, the engine must be
throttled to prevent excessive overspeeding.
The first improvement over the fixed-pitch
propeller was the ground adjustable-pitch type.
On this type, the blade angle (pitch) could be
changed or adjusted on the ground by manually
twisting the blades in the hub to the desired angle.
When the angle was increased to improve cruising
conditions, takeoff conditions suffered. An aircraft taking off from the ground with the
propeller at a high blade angle position is much
the same as setting a car into motion in high gear.
The engine is not able to produce full horsepower
because the high blade angle loads the propeller
too much to enable it to turn over at the full rpm
of the engine.
The next design was the two-position propeller. It enabled a pilot to use a low blade angle
(high rpm setting) for takeoff, climb, and
necessary operational acceleration. The pilot
could then change the propeller blade angle in
flight to a higher blade angle (low rpm setting)
for cruise. With this propeller, full engine rpm
could be developed for takeoff. Aircraft speed
could be increased at cruise with a decrease in
engine power because the high-pitch propeller
takes larger bites out of the air. The two-position
propeller did not, however, produce the most
efficient and economical use of engine horsepower
for all of the numerous intermediate flight
conditions encountered by aircraft.
Constant-speed propellers were eventually
designed to maintain a preselected rpm automatically. Suppose the aircraft is heading into a
gradual climb. The constant-speed propeller maintains the selected rpm automatically by turning
the propeller blades to a lower angle. That is, the
propeller takes a smaller bite of air when the load
on the engine is increased. Now, should the aircraft assume a nose-down attitude, the propeller
blades move automatically to a higher blade angle;
hence, the propeller takes a larger bite of air. In
other words, increase the load on the engine and
the propeller takes a smaller bite of air. Decrease
One of the main requirements of any propeller
is its ability to withstand severe stresses. We will
discuss these stresses, which are greatest near the
hub, in the following paragraphs. Figure 8-12
shows the forces acting on propeller blades.
Figure 8-12.-Natural forces acting upon propeller blades.
and forth while the propeller is turning. Fluttering causes a distinctive noise, which is nearly
drowned out by the exhaust noises of the engine.
Fluttering will weaken the propeller and may
result in structural failure unless detected early and
Centrifugal Force
The greatest force acting upon the propeller
blade is centrifugal force. This force tends to pull
the blade of a spinning propeller out of its hub.
To prevent the blades from breaking into
fragments or flying off into space, the blade is
thicker near the hub. The hub is made from a
strong steel forging.
The propeller system has one primary
function—increasing or decreasing pitch as
required by power lever movement. Safety
features incorporated in the 54H60-77 propeller
system include the automatic mechanical pitchlock, the automatic negative torque control, the
mechanical low pitch stop with a secondary
hydraulic low pitch stop (referred to as the beta
follow-up), and an emergency feathering system.
The complete 54H60-77 model propeller,
shown in figure 8-13, consists of the front antiicing propeller spinner, the hub mounting
bulkhead assembly, the variable pitch aircraft propeller (propeller assembly), the rear deicing
propeller spinner, the air baffle assembly, the
propeller control (integral oil control) assembly,
and the propeller afterbody assembly.
Thrust Bending Force
Thrust bending force causes a rotating propeller to try to pull away from the aircraft.
Because it is held back by the hub and the load
of the aircraft it is pulling, the blade tips, which
are thinner and lighter than the blade shank, bend
forward. The sum of these bending forces on the
blades is carried at or near the hub. Hence,
the section of the blade at the hub must be
proportionately thicker. Centrifugal force
counteracts thrust bending force by its tendency
to pull the blades in a straight line.
Torque Bending Force
Torque bending force is the tendency for a
blade to bend backwards, throughout its length,
in a direction opposite rotation. This bending
force is created by the density of the air.
Spinners and Afterbody Assemblies
The main purpose of the front and rear
spinners is to streamline the airflow around the
outside of the propeller assembly for cooling. The
front and rear spinner assemblies improve the
aerodynamic characteristics of the whole propeller
assembly. They enclose the dome, barrel, and oil
control assemblies. The front spinner has an air
inlet in the middle of it. Cooling ram air enters
to cool the dome, barrel, and oil control
The propeller afterbody assembly streamlines
airflow into the engine air inlet. The afterbody
assembly has a top and bottom half. The two
halves of the afterbody have electrical deicing
wires to prevent ice buildup on the back side of
the propeller assembly.
Aerodynamic Twisting Force
Aerodynamic twisting force tries to rotate the
blades in the hub to an increased blade angle. The
point at which this force is exerted most strongly
on the chord of the airfoil is known as the center
of pressure. During normal cruise conditions, this
center of pressure is nearer the leading edge of
the propeller, so the force tends to rotate the
blades to a higher pitch.
Centrifugal Twisting Force
The centrifugal twisting force on the blades
tends to twist them to a lower pitch angle. This
occurs because all parts of the propeller try to
remain in a plane parallel to the plane of rotation.
Hub Mounting Bulkhead Assembly
and Propeller Assembly
The hub mounting bulkhead is the mounting
surface for the front and rear spinner assemblies.
The variable pitch aircraft propeller (propeller
assembly) has four major subassemblies. They are
the barrel assembly, the blade assembly, the dome
assembly, and the pitchlock regulator assembly.
Propeller Vibration
Sometimes, in the face of these forces, a propeller loses some of its rigidity. The result is a
flutter, which is a type of vibration in which the
tips of the blades attempt to twist rapidly back
Figure 8-13.-Propeller system.
As we discuss the propeller assembly, refer to
figure 8-14.
BARREL ASSEMBLY.— The propeller barrel
assembly serves several functions. It retains the
four propeller blades and also supports the dome
assembly and the propeller control assembly.
Engine torque is transmitted to the propeller by
the barrel assembly, which mounts and secures
to the front of the reduction gearbox propeller
The barrel assembly is a split type; the front
and rear barrel sections are manufactured and
balanced as a matched pair. These sections are
kept together throughout the service life of the
propeller. The high centrifugal blade loads are
carried by the barrel shoulders and lips at each
blade position.
A machined integral extension on the rear
barrel half is splined internally and has seats at
both ends. The front and rear cones are beveled
to match the extension seats for centering and
securing the propeller on the propeller shaft. The
extension is splined externally for driving pumps
in the propeller control assembly. A propeller hub
nut locks the barrel assembly to the reduction
gearbox propeller shaft. The propeller hub nut has
a flange on its inboard end that butts against the
front cone.
BLADE ASSEMBLY.— The broad, lightweight propeller blade is forged from a solid
aluminum alloy, which gives it the strength
necessary to obtain the high thrust capability at
low aircraft speeds. The blade butt is partially
hollow to allow for installation of the blade
bushing and blade balancing assembly. Propeller
balancing is discussed in the balancing section of
this chapter.
The blade shank has a molded fairing that is
composed of a plastic foam material covered with
a nylon reinforced neoprene material. The heater
assembly is bonded to the leading edge of the
fairing. It contains the necessary blade deicer
elements to prevent ice buildup on the blade
assembly. Blade heater element damage, involving
cut or broken heater wires because of weather
corrosion or foreign object damage (FOD) strikes,
can be repaired if no more than four wires are
damaged. If more than four wires are damaged,
the heater assembly must be replaced. The
purpose of the blade fairing (cuff) is to streamline
and direct the airflow to the engine intake.
of the propeller system. The dome assembly
is mounted on the front barrel shelf and
held in position by the dome retaining nut.
The principal components of this pitch-changing
mechanism are the rotating cam, the stationary
cam, the piston assembly, and the low pitch stop
assembly. The low pitch stop assembly is screwed
into the lever sleeve bushing at the front of the
Preformed packings are used throughout the
dome assembly for internal leakage control and
to seal the piston assembly in order to separate
the inboard and outboard hydraulic pressure
necessary for blade movement. Shims are used to
establish the proper clearance between the rotating
cam and the blade segment gears. The dome
assembly is mounted in position on the front
barrel shelf. It is held in place by the dome
retaining nut that is locked in place by a special
head screw.
The low pitch stop assembly screws into the
dome assembly. It sets the desired low pitch
stop blade angle. In the flight range of operation,
the low pitch stop lever assembly prevents
the propeller blade angle from going below
13 degrees. In the ground range, extra high
hydraulic fluid pressure actuates the low pitch
stop assembly, allowing the piston to move
further outboard. This turns the blades from the
low pitch stop position towards the reverse blade
The propeller pitchlock regulator assembly
mounts within the propeller barrel and is splined
to the propeller hub nut. The pitchlock regulator
assembly directs hydraulic pressure to the outboard and inboard sides of the dome piston. It
also serves as a safety feature by preventing a
decrease in blade angle, by pitchlocking, under
certain conditions. Pitchlock occurs if hydraulic
control pressure is lost or during an overspeed of
103 to 103.5 percent.
The pitchlock regulator assembly contains two
ratchet rings that are spring-loaded together, but
are held apart by hydraulic pressure. One ratchet
ring is splined to the rotating cam of the dome
assembly. The other ratchet ring is splined to the
propeller rear barrel half and does not rotate. If
hydraulic pressure is lost, the ratchet rings come
together, and their teeth mesh to prevent a
decreased blade angle. This is referred to as pitchlock, and can only occur between approximately
15 to 60 degrees of blade angle. Pitchlock is
DOME ASSEMBLY.— The propeller dome
assembly is the blade angle changing mechanism
Figure 8-14.-Variable pitch aircraft propeller.
Propeller Control Assembly
(Integral Oil Control Assembly)
mechanically cammed-out below 15 and above 60
degrees. When in pitchlock, the propeller operates
as a fixed-pitch propeller. However, the reverse
rake of the pitchlock ratchet teeth allows rotating
of the propeller into higher blade angles for
feathering or to regain control of the propeller.
The propeller control assembly is a
nonrotating integral oil control mechanism. It
mounts on the rear extension of the propeller
The electrical connections are for the pulse
generator coil, the auxiliary pump motor, the
synchrophasing system, and the anti-icing and
deicing systems.
barrel. See figure 8-15. The control assembly
contains two major components—the pump
housing assembly and the valve housing assembly.
The pump housing assembly contains the
hydraulic reservoirs, pumps, and valves. The valve
housing assembly is where all mechanical and
electrical connections necessary for propeller
operation are made. The mechanical connections
include the linkages between the engine control
system and the negative torque system (NTS).
housing assembly forms the lower part of the propeller control assembly. The pump housing
contains five positive displacement gear-type
pumps (three mechanically driven and two
Figure 8-15.-Propeller control assembly.
electrically driven). An externally mounted ac
electric motor drives the two common shafted
auxiliary pumps. Two hydraulic fluid sumps are
contained in the pump housing assembly. One is
a pressurized sump with a capacity of 6 quarts.
The other is an atmospheric sump with a capacity
of 4.5 quarts. A pressure cutout switch located
in the pump housing serves to terminate the
action of the auxiliary pumps. This action occurs
when the feather blade angle is reached.
The three mechanically driven gear-type
pumps are the main pump, the main scavenge
pump, and the standby pump. The electrical
pumps are the auxiliary scavenge pump and the
feather pump. The feather pump is used for static
ground operations. It also serves to complete the
feather operation in flight. An electrical motordriven pump is needed since the output pressure
of the mechanically driven pumps is reduced in
proportion to the decaying propeller rpm.
When speeder-spring force is equal to the
flyweight centrifugal force, the pilot valve centers
in the pitch change sleeve to block pitch change
hydraulic pressure from the propeller dome. Any
change in the engine speed will change the outward position of the flyweight. The flyweight
shifts the pilot valve to direct pitch change
hydraulic pressure to the dome.
Constant-speed governing is blocked out in the
ground handling range. The propeller blade angle
is coordinated with the position of the power
lever. Interaction of the cams on the alpha and
beta shaft control the position of the pilot valve
in the speed servo governor. When the power lever
is moved, a cam on the alpha shaft positions the
pilot valve to obtain a corresponding blade angle.
As the blade pitch changes, a cam on the beta
shaft returns the pilot valve to a position that will
maintain the blade pitch at the angle scheduled
by the power lever,
Rigging pin holes are located on the valve
housing assembly for rigging the valve housing
to the propeller assembly and power lever.
Adjustments are provided to set the mechanical
governor speed and the reverse and ground
handling blade angles.
valve housing assembly is considered the brains
of the propeller system. It mounts to the upper
part of the pump housing assembly, forming the
propeller control assembly. The valve housing
assembly is the most complex assembly of the
propeller system. Major units of the valve housing
assembly are the speed servo governor assembly,
flyweights, speeder spring, pilot valve, feather
valve, feather solenoid valve, main and standby
regulating valves, high-pressure relief valve, beta
and speed-setting lever assembly, alpha and beta
pinion shafts, linkage support assembly, and the
electrical branch cable.
The two ranges of operation controlled
through the valve housing assembly are the
governing range and the taxi range. The governing
range is commonly called the alpha or the flight
range. The taxi range is commonly called the beta
scheduling or the ground handling range. All
primary propeller operations, except for feathering and unfeathering the propeller, are determined
by the position of the pilot valve in the speed servo
governor. Hydraulic fluid for blade angle change
operation is pumped from the pressurized sump
by the main pump (and standby pump if needed)
to the pilot valve chamber.
Flyweights are geared to the propeller shaft.
This causes their rotation to develop centrifugal
force in direct relation to the engine speed. The
centrifugal force extends the flyweights outward
and pushes the pilot valve toward the increased
pitch position. This movement of the pilot valve
opposes the speeder-spring force, which pushes
the pilot valve toward the decrease pitch position.
In today’s Navy, comprehensive and systematic means of maintaining a multiengined propeller system is essential. You, as an Aviation
Machinist’s Mate, must know the procedures for
day-to-day maintenance. You must know the procedures for removal and installation of a propeller, rigging, adjustment, and troubleshooting
of propeller systems. You should also be familiar
with the procedures for propeller balancing and
leakage test requirements. The modern-day propeller system is a complex and durable system,
and, with proper maintenance, a highly reliable
aircraft system.
NOTE: The information noted here is
general. Refer to applicable technical
instructions for the specific methods used
on a particular model of propeller.
Propeller Inspection
Inspect the blades daily for any gouges, nicks,
scratches, or gross damage. If the propeller has
struck any object (static or rotating), inspect the
blades carefully for damage. For example, yellow
paint marks on a blade indicate that the propeller
could have possible damage from hitting a piece
of support equipment.
the damage. After removal of metal, measure
carefully any blade that has a finished depression.
Send blades with depressions exceeding repair
limits to overhaul.
First, measure the depth of each damaged area
with a dial gauge. Raised edges may interfere with
the setting of the knife edge of the gauge. Using
aluminum oxide paper, remove any raised edges
adjacent to the damaged area. Set the base of the
gauge parallel to the longitudinal axis of the blade.
Gouge depth is the difference between the deepest
point in the damaged area and the adjacent blade
surface. In a reworked area, it is the difference
Propeller Repair
Cuts, scars, scratches, surface cracks, nicks,
etc., are common types of damage. Repair by
forming a “saucered out” depression smoothly
blended into the blade surface. Carefully examine
the area with no less than a three-power magnifying glass. Use the magnifying glass to make certain
the bottom of the damage is completely removed.
Use local etching to be sure there are no cracks
at the bottom of-the rework. To avoid removing
excessive metal, make a local etching check at
regular intervals during the process of removing
Figure 8-16.-Propeller blade damage rework limits.
A line of continuous repairs would materially
weaken the blade structure. For an illustration of
minor repairs to the aluminum alloy propeller
blade, refer to figures 8-16 and 8-17.
between the deepest point in the damaged area
and the reworked surface.
To remove damage from a blade, remove the
damaged metal and the adjacent area with diemakers’ rifflers and aluminum oxide paper. Work
the damage from the blade and smoothly blend
the resulting depression into the surrounding blade
surface. Do not try to move any of the metal in
the damaged area by cold-working. Repairs in a
given area on a blade are unlimited provided their
locations do not form a line of continuous repairs.
NOTE: The blade damage repair limits
shown in figures 8-16 and 8-17 are examples
of repair limits found in maintenance
instruction manuals. Refer to the specific
maintenance instruction manual for the
propeller that you are working on.
Figure 8-17.-Propeller rework guide.
Propeller Removal
It usually takes a crew of three to remove the
propeller and place it on the dolly. Refer to the
applicable maintenance instructions for a list of
tools and procedures used to remove the propeller
assembly. Cycle the propeller to maximum reverse
to aid in the removal of the afterbody assembly.
Maximum reverse is the only position that allows
afterbody removal without causing damage to the
afterbody or propeller blades. After you remove
the top and bottom sections, cycle the propeller
back to the feather position. Secure all electrical
power to the propeller. Disconnect the cannon
plugs and propeller control input linkage, and
stow them away to prevent snagging during
removal. Remove the propeller spinner. Remove
the dome cap and oil transfer tube. Position an
oil shield under the dome to catch oil spillage. See
figure 8-18.
The propeller should NOT be cycled for
at least 10 minutes before removing the
dome cap.
Figure 8-18.-Dome cap removal.
Unscrew the dome retaining nut and remove
the dome assembly. Remove the pitchlock
assembly with the puller. This assembly includes
the pitchlock control cam, pitchlock stationary
ring and pin, and externally splined spacer ring.
Install a dynamometer between the hoist and yoke
Propeller Cleaning
If you disassemble the propeller, clean all the
parts with approved cleaning solution except the
deicing brushes and slip rings. Thoroughly dry all
parts after cleaning. Vapor blasting is NOT
permitted on this propeller. The use of rags or
paper for cleaning or wiping internal parts of the
propeller and control assembly is NOT permitted.
The use of these materials may cause lint or
minute particles to enter the hydraulic system.
Malfunctioning of parts is possible. Unused parts
or parts not reassembled within a reasonable time
should be preserved with a corrosion-preventive
compound. Exact procedures for the cleaning and
prevention of corrosion on propellers are found
in the applicable maintenance instructions.
NOTE: When removing or installing
the propeller, do not turn the No. 1
blade beyond positive 100 degrees or
minus 15 degrees. Movement passed these
limits will damage the propeller. Damage
occurs when the beta follow-up shaft
releases from the beta gear segment.
Damage also occurs when the beta shaft
moves into the control assembly beyond its
Propeller Installation
Remove the propeller and control assembly
from the propeller shaft. Install them on a dolly,
and secure them before removing the hoist.
Remove the rear cone from the engine shaft
and retain it with the propeller assembly. See
figures 8-19 and 8-20.
Installation procedures are the reverse of
removal. Before installing the propeller on the
propeller shaft, install (or clean and inspect) the
torque retainer (drive bracket assembly) and the
negative torque bracket assembly on the engine
Figure 8-19.-Propeller removal.
reduction gearbox. Reclean the propeller shaft and
apply a light coating of hydraulic fluid over the
entire shaft.
Propeller Servicing
The propeller fluid fill cap is under the access
doors of the top half of the spinner afterbody.
The total capacity of the propeller system is about
25 quarts.
DO NOT use engine oil in the propeller
system. If engine oil is added to the
control by mistake, remove the propeller.
Remove both the control and propeller.
Send the assembly to an overhaul activity
for complete seal replacement. Observe
duty cycle for the feather pump to avoid
overheating and damage.
Figure 8-20.-Propeller stowage.
Upon initial installation, fill the propeller with
hydraulic fluid, Add as much propeller hydraulic
fluid as possible to the atmospheric sump and the
pressurized sump. Statically cycle the propeller
several times between feather and reverse. Add
fluid each time the PROP PUMP No. 1 light
illuminates. Repeat this procedure until you add
a total of 25 quarts. After completion of the first
engine runup, inspect the pressurized sump
capacity. Note the fluid level, using the shoulder
within the filler hole as a guide. If necessary,
completely fill the pressurized sump.
If the system was full on the last flight and
is below FULL at the next check, suspect external
leakage. Make a thorough visual inspection of the
propeller and control to locate the site of the
leakage. Statically cycle the propeller before
checking the fluid level. Do not remove the filler
cap for at least 10 minutes after operation of the
propeller or actuation of the auxiliary pump. If
the fluid level is low at the initial check, run the
auxiliary pump. Recheck the fluid level. Operating
the propeller into the feather position and back
should be sufficient. If the fluid level is still low,
add the amount required to bring the fluid level
up to that specified by the MIM.
It is important to check the fluid level in the
control assembly before checking the operating
propeller system. This procedure is especially true
before feathering. Without fluid, the pressure
cutout switch will not operate to stop auxiliary
pump motor operation.
During static system operation, intermittent
operation from about 75 degrees to air or ground
start positions may cause the pitchlock to engage.
Decreased pitch operation cannot continue until
the pitchlock releases. This disengagement is made
by moving the propeller toward the feather
position a few degrees.
publication for the special tool requirements.
Check for freedom of movement for both power
levers and E-handles. There must be NO binding
or interference. Check cable tension with a
tensiometer. Insert rigging pins in slots. Rig pins
should go in with slight finger pressure. Adjust
control rod lengths if needed. Remove rig pins.
Check rigging at max reverse, takeoff, and flight
idle by comparing coordinator readings and
inserting rig pins in appropriate engine or valve
housing slots. After mechanical rigging checks
agree, adjust valve housing assembly for setting
blade angles. See fig 8-21.
After all rigging has been completed, check
for installation of bolts, nuts, and safety wire.
Torque all bolts and nuts, and safety wire all rod
ends, as required by the appropriate technical
manual. At this point, you should remove all
rigging pins and install the valve housing
atmospheric sump filler cap. Now complete propeller and engine checks to test for proper
Feathering Check
Depressing the feather button in the flight
station causes normal feathering. This action
supplies voltage to the holding coil of the feathering switch, auxiliary pump, and feather solenoid.
Hydraulic fluid positions the propeller control
feather valve to feather the propeller. When the
propeller has fully feathered, pressure buildup will
operate a pressure cutout switch. The switch
causes the auxiliary pump and feather solenoid
to become de-energized through a relay system.
Feathering also occurs by pulling the emergency
shutdown handle (E-handle). This action mechanically positions the feather valve and electrically
energizes the feather button holding coil to send
the propeller to feather.
Rigging and Adjustment
Unfeathering Check
During the initial engine/propeller installation,
or whenever a fuel control, coordinator, or
linkage has been replaced, make a complete
rigging check. The final propeller control linkage
rigging and valve housing adjustment is done with
the propeller installed. In most cases, when a T56
engine is ready for issue (RFI) from an aircraft
intermediate maintenance department (AIMD) to
a squadron, the fuel control to coordinator rigging
has been completed.
A minimum of five rigging pins are necessary
to rig and adjust the propeller control linkage and
valve housing. Refer to the appropriate technical
To unfeather the propeller, pull the feather
button and hold in the unfeather position. This
action causes the auxiliary pump to come on.
Fluid pressure flows to the decrease side of the
pitch change piston in the dome. This action
unlatches the feat her latches. The propeller will
start to unfeather. Upon reaching the air start
pitch angle (45 to 48 degrees), the air start beta
switch closes. Closing the switch energizes the air
start control relay. This relay energizes the feather
valve solenoid, which sends the blades back
towards feather. The return of the blades toward
Figure 8-21.-Valve housing assembly (adjustments).
feather opens the air start beta switch to reenergize the air start relay. Now the blades are
cycling around the air start blade angle to stabilize
the propeller speed and engine speed. These
stabilized speeds prevent a NTS condition from
occurring during an air start. The air start blade
angle for unfeathering cuts out ground unfeathering when you depress the pressure cutout override
(PCO) button. This button is usually adjacent to
the feather button. Depressing the PCO button
prevents the blades from cycling around the air
start switch. The PCO button allows the blades
to go to a lower blade angle setting.
Fuel Governor, Pitchlock, and
Reverse Horsepower Checks
The purpose of the fuel governor check is to
be sure that the fuel control governor will limit
the engine speed if the propeller governor fails.
The pitchlock check makes sure that the propeller
pitchlock will engage to prevent the propeller from
going to a lower blade angle. The reverse
horsepower check will ensure that the reverse
horsepower will operate normally.
NOTE: Stopping a decreasing blade angle
during ground operations may cause a
problem. If stoppage occurs above the low
pitch stop, do not, under any circumstances, attempt to decrease the blade angle
any further.
With the engine running at low rpm, place the
NTS/feather valve check switch in the NTS check
position. Turn the fuel/ignition switch to OFF.
NTS action should develop and illuminate the
NTS light. If unsuccessful, maintenance action
is necessary.
NTS Check On Shutdown
Increase the blade angle before trying a
decrease, so the pitchlock teeth will disconnect.
The pitchlock teeth will engage upon release of
the feather button. Failure to increase the blade
angle will cause damage to the pitchlock teeth.
Accomplish all propeller balancing in a
horizontal plane using the propeller balancing
kit 7A100, or its equivalent. See figure 8-22.
Before performing actual propeller assembly
buildup and balancing, you must always refer to
the appropriate technical publication.
Preliminary and final balance has already been
completed on new and overhauled propellers
before they are disassembled and shipped to an
AIMD. Do not perform preliminary balance if
final balance can be obtained first.
NOTE: The final balance check can be
erroneous because of residual hydraulic
fluid in the propeller dome assembly. You
must make sure the dome assembly is completely drained of any residual hydraulic
fluid before installing the dome assembly
for the final balance cheek.
You must obtain horizontal balancing on all
propellers during assembly. Horizontal balancing
must be performed in a room free of air currents
and with the propeller assembly clean and dry.
The plane of the blades must be horizontal, and
the blade pitch must be set at 45 degrees.
Do not install the dome cap, low pitch stop
assembly, pitchlock regulator assembly, propeller
hub nut, hub mounting bulkhead assembly, and
their associated parts. These units are not included
as part of the balancing procedure. Install the
dome assembly without the dome-to-barrel
preformed packing and gear preload shims.
Tighten the dome retaining nut snugly past its
normal locking position. Use masking tape to hold
the dome retaining nut special head screw (without
its cotter pin) in place at its normal locking
Final Balance Check
The final balance check is obtained by adding
bolts, washers, and nuts to balancing holes in the
deicer contact ring holder assembly near the outer
edge. If possible, bolts, washers, and nuts should
be divided equally on each side of the deicer
contact ring holder assembly. Do not disturb
similar bolts, washers, and nuts, which are painted
red and already located in the balancing holes.
They are used for balance of the holder assembly
itself, not the propeller. Use special bolts, washers,
and nuts on the deicer contact ring holder
assembly installed on the propeller. For the
plastic molded holder assembly, use no more
than six AN960-10 washers on one bolt; use no
more than six NAS514P1032-16 bolts and six
MS20364-1032A nuts.
Obtain final balance with the propeller
assembly mounted on the horizontal balance
machine, with the plane of the blades horizontal
and the dome assembly installed. The sensitivity
of the balance machine must be calibrated so that
any unbalance shown by the machine may be
corrected or reversed by applying a restraining
moment of 6 inch-ounces.
If final balance cannot be obtained because
of the maximum limit on the number of bolts,
washers, and nuts that can be added to the deicer
contact ring holder assembly, it is necessary to
obtain preliminary balance first, and then final
balance. Remove the final balance bolts, washers,
and nuts from the holder assembly, if they are
The bolts, washers, and nuts that are
colored red must not be removed. These
are used for balance of the holder assembly
itself, not the propeller.
Preliminary Balance
If final balance cannot be obtained, preliminary balance must be obtained by installing
balance washers on the blade balancing plugs of
the light blades.
With the propeller suspended on the balancing
stand, you should place the balance washers on
the shanks of the light blades next to the outboard
Figure 8-22.-Propeller balancing kit (7A100).
electric contact rings. See figure 8-23, Preliminary
balance has been obtained when the propeller
shows no tendency to tilt, or when tilting maybe
stopped or reversed by the addition of the lightest
balance washer to one or more blades on the light
side of the propeller.
to performing the external and internal hydraulic
leakage test.
After you have determined the amount of
washers to install, remove the propeller from the
balancing stand, Disassemble the propeller until
the light blades have been removed. Install the
required washers on the blade balance plugs.
Before you begin the propeller test, first verify
that the propeller test equipment has been
inspected, serviced, and properly assembled. This
must be done prior to the installation of the
propeller. Install the propeller on the test equipment in accordance with the appropriate technical
publication before beginning the hydraulic leakage
After the hydraulic fluid is warmed up,
exercise the propeller between the high and the
low blade angles several times to purge air from
both the test equipment and propeller system.
Purging will avoid erratic operation during the
external and internal leakage tests.
External and Internal
Hydraulic Leakage Test
Reassemble and reinstall the propeller on the
balancing stand. Recheck the preliminary balance
and obtain the final balance check, as previously
After you obtain final balance, remove the
special head screw taped on the dome assembly.
Remove the dome assembly from the propeller,
using care not to disturb the 45-degree setting of
the rotating cam. Remove the balancing arbor
from the propeller.
Remove the propeller from the balancing
stand. Remove the deicer contact ring assembly
and the packing seal ring with its preformed
packing. The propeller must be reassembled prior
Attempting to initiate a decrease in propeller blade angle when the propeller is in
a range from about 60 degrees to 15
degrees may cause the pitchlock to engage
or cause damage to the ratchet teeth. If it
becomes necessary to stop in this range,
first increase the blade angle to above the
pitchlock range, and then proceed to a
decreased blade angle.
The test equipment used to supply the various
pressures and flow requirements is the hydraulic
propeller test stand GS1221. With the test stand
maintaining 150 psi, cycle the propeller blades
between a low blade angle and a high blade angle
until a total of eight cycles are completed. No
external leakage is permissible during the cycling.
If any external leakage occurs at the junction
of the barrel half seals and the blade packings,
eliminate the leakage by separating the barrel halves and adding zinc chromate putty,
MIL-P-8116, to the junction. You must control
the amount and location of the putty to prevent
it from getting into the barrel cavity. Leakage
from the blade bores can be eliminated by
replacing the blade preformed packing. The
complete external leakage test must be rerun after
Figure 8-23.-Propeller suspended for balancing.
any external leakage corrective work has been
The internal flow and leakage tests are
designed to ensure the proper internal operation of
the propeller system. The test equipment will
supply the various hydraulic pressures to the inboard and outboard side of the dome piston, surge
valve, and pitchlock mechanism ensuring smooth
blade angle movement to the reverse and feather
blade angles.
If internal flow and leakage requirements are
not in compliance with the appropriate technical
publication, you must disassemble the propeller
and inspect all visible packings for damage and/or
proper location. All internal flow and leakage tests
must be completed before you can issue the propeller to an operating activity. The time for
discovering external/internal leakage is NOT
when the propeller has been installed and is ready
for final rigging.
After completing this chapter, you will be able to:
Recognize basic troubleshooting procedures
and common troubleshooting errors.
Identify the types and uses of the various
drawings used in aircraft maintenance publications.
Recognize the characteristics of a multimeter
and the use of the high-voltage insulation
Recognize the purpose of, and the procedures
for, the use of the two types of borescopes.
Recognize the main variables involved when
performing engine tests and the purpose of the
materials used in engine field cleaning.
Recognize the codes and methods for the
identification of various fluid lines.
corrective measure has been taken to correct a
malfunction. Never be hasty in assuming any conclusions. Do not confuse the significance of cause
and effect. For example, low oil pressure, in itself,
is not truly a trouble. It is only an indication that
a problem exists. the true cause of which is the
The preceding paragraph emphasizes the
importance of tracing an abnormal indication to
its source. Do not be satisfied that restoration to
a normal condition is accomplished by some
obvious adjustment. Adjusting the oil pressure
relief valve setting may not be the only correction or the correct step for a low oil pressure
discrepancy. Refer to table 9-1 for possible low
oil pressure causes. Invariably, there is a reason
for an abnormal engine indication. You must
know the reason and make sure that proper
corrective measures are taken before the
discrepancy is signed off. Proper corrective
measures require the use of time-proven methods,
not guesswork.
As previously stated, to troubleshoot intelligently, you must know the system and the
function of each component in the system. You
must have a mental picture of the location of each
component in relation to other components in the
Efficient maintenance of any aircraft depends
upon the operating and maintenance personnel’s
familiarity with the aircraft and its associated
parts. You must have a thorough knowledge of
the engine and know its normal operating conditions. You should know the oil pressures, fuel
pressure, rpm limits, and exhaust gas temperatures
for all of the engine operating ranges. In addition to knowing the normal engine conditions, you
should understand the effects of atmospheric conditions on the engine. You must consider such
things as air temperatures and pressure, fuel
temperatures, and wind effects.
All aircraft maintenance activities are plagued
by repeat discrepancies. These discrepancies show
up repeatedly on the same aircraft. They show up
even after being signed off as previously corrected.
There are many possible causes for repeat
discrepancies. The most common cause is poor
troubleshooting procedures.
It is often unsafe to assume that the
engine is completely repaired when some obvious
Table 9-1.-Oil Pressure Troubleshooting Chart
responsibility to get first aid training, including
CPR. Personal CPR training is available at many
Navy medical facilities. Many local fire stations
and Red Cross agencies also offer this type of
training. It is important for you to be currently
certified in the special skills of CPR when
troubleshooting electrical systems or ignition
system. Studying schematic diagrams of the
system is one way to establish this mental picture.
In the performance of your duties, you may
come across many potentially dangerous conditions and situations. When working, there is one
rule to stress strongly—SAFETY FIRST. Whether
you are working in the shop, on the line, or during
a flight, you should follow prescribed safety
procedures. When working on or near aircraft,
there is the danger of jet blast, intake duct,
propeller, or rotor blades. Because of these
dangers, you need to develop safe and intelligent
work habits. You should become a safety
specialist, trained in recognizing and correcting
dangerous conditions and unsafe acts.
You should study the cardiopulmonary
resuscitation (CPR) training section of Basic
Military Requirements, NAVEDTRA 10054, and
Standard First Aid Training Course, NAVEDTRA 10081. These texts provide valuable information, but they do not replace training. It is your
Do not perform CPR unless you have had
the proper training.
The need for troubleshooting is shown by poor
engine performance or system operation. The
maintenance instruction manuals for an aircraft or engine contains specific information
for understanding and troubleshooting engine
problems. Understanding the cause of common
engine problems will help you avoid them.
Air System
All personnel should clearly understand that
any obstruction or interruption of the air as it
moves through the engine will affect the entire
operation of the engine. Entrance of foreign
matter into the air system is one of the main causes
of service trouble. As the foreign matter passes
rearward, it causes improper combustion, blockage of cooling air passages, and combustion
section damage. The turbine section and the afterburner section may be damaged.
Troubleshooting procedures are similar in
practically all applications. Troubleshooting is the
logical or deductive reasoning procedure used
when determining what is causing a particular
system malfunction. There are six steps involved
in good deductive troubleshooting.
1. Conduct a visual inspection. This inspection should be thorough and searching. Check
all lines, units, mechanical linkages, and parts
for evidence of leaks, looseness, material
condition, and proper installation. During this
visual inspection, check all systems for proper servicing. Check reservoirs for proper servicing
levels, accumulators for specified preload,
Compressor Section
The compressor section of the turbojet engine
is subject to a variety of service troubles from
foreign matter passing through the compressor.
Small pebbles, nuts, bolts, and washers can cause
damage to the compressor rotor blades and stator
vanes. Damage severe enough to warrant complete
overhaul of the engine may result from the ingestion of foreign matter.
The aircraft discrepancy book (ADB) and the
visual information display system/maintenance
action forms (VIDS/MAFs) are invaluable
troubleshooting tools. The type of flight and
previous maintenance actions may give an
indication of the engine’s problem. You may
save many man-hours of troubleshooting an
engine discrepancy by first screening the ADB
for prior maintenance. You may find prior
maintenance that has contributed to the cause of
the new problem. You may find improperly
performed maintenance, and trends of prior
Combustion Chamber and Turbine Section
High temperatures take place in the combustion and turbine sections. Few major service
troubles occur when maximum temperature and
maximum rpm limits are closely monitored. Hot
starts, overtemperatures, and exceeding maximum
rpm will cause extensive service troubles in the
combustion and turbine sections.
2. Conduct an operational check. Check the
malfunctioning system or subsystem for proper
operation. This includes checking the mechanical
movement of throttles and linkages for the
correct movement without binding. Check for the
proper sequence of operation and speed, and
whether a complete cycle was obtained. Sometimes an operational check requires special test
equipment or engine turnup to check out the
system properly.
Fuel System
The fuel system of the gas turbine engine
begins at the engine-driven fuel pump. The system
ends with delivery of fuel to the combustion
chambers through fuel nozzles in the form of
highly atomized spray. Common service troubles
are wrong types of fuel, water or other foreign
matter in the fuel, defective fuel-boost pumps, and
clogged fuel filters.
3. Classify the trouble. Malfunctions usually
fall into four basic categories. They are the
hydraulic, pneumatic, mechanical, and electrical
groups. Using the information gained from steps
1 and 2, you must determine under which
classification the malfunction occurs. Something
affecting normal flow of fluid would be classified
under the hydraulic classification. The flow of
fluid may be affected by external and internal
leakage, total or partial restriction, or improper
Lubrication System
The proper grade of oil and cleanliness when
replenishing the oil supply are absolute requirements to avoid lubrication system troubles. Gas
turbine engines have few moving parts, but the
speed at which the parts move, especially the
rotor, requires constant lubrication to the bearings. Oil starvation for a short period of time can
result in having to remove the engine for overhaul.
Something affecting the normal flow of
compressed gases is classified as a pneumatic
malfunction. This type of malfunction stems from
the same general sources as hydraulic malfunctions mentioned in the previous paragraph.
never use part replacement as a method of
troubleshooting. Each year, thousands of dollars
worth of engine parts are returned to overhaul
activities with labels stating that they are faulty.
Upon completion of test and check by the
overhaul activity, many parts show no defect. This
practice is wasteful and very expensive. Another
problem occurs when mechanics replace a major
assembly. The problem could have been corrected
by replacing a subassembly. The replacement of
an entire unit when only a small part is at fault
reflects poor maintenance practices. When
troubleshooting, remember you are working with
equipment that is both expensive and scarce.
Replace equipment only after it is thoroughly
tested and determined to be faulty.
Many units that operate hydraulically or
pneumatically use mechanical linkage. If a
discrepancy in the linkage exists, it will affect the
system’s operation. Many mechanical discrepancies
are found during visual inspections. Common
mechanical problems are worn linkages, broken
linkages, improperly adjusted linkages, or
improperly installed linkages.
You must be able to determine if the electrical
system is functioning normally. Common electrical problems are complete power failure,
circuit failure, or component failure. Electrical
components have circuit breakers and switches to
operate or control them. Check the obvious things
such as tripped circuit breakers or switches in the
wrong position. Although AEs normally
troubleshoot electrical problems, sometimes you
will work together on a problem. Learn to read
electrical schematics and use a multimeter so you
can trace electrical power requirements
throughout the affected system.
6. Conduct a final operational check. The
affected system must have a thorough check to
determine proper operation. Checking the engine
for proper operation consists primarily of reading
engine instruments. Compare the readings with
those given by the manufacturer for specific
engine conditions, atmospheric pressures, and
Early model engine test procedures used rpm
as the engine operating parameter to establish
thrust. Today’s engine test procedures use engine
pressure ratio (EPR) as the primary thrust indicator. EPR is the ratio of the total pressure at
the front of the compressor to the total pressure
at the rear of the turbine.
4. Isolate the trouble. This step calls for sound
reasoning. A full and complete knowledge of
engine theory, as well as a complete understanding of the affected system is necessary.
During this step you must use knowledge and
known facts to determine where the malfunction
exists in the system. Usually the trouble can
be pinned down to one or two areas. Eliminate
those units that could not cause the known
symptoms and those that can be shown to operate
Using EPR as the thrust indicator means that
on a hot day it is quite possible for the engine rpm
to exceed 100 percent. If you use rpm on a hot
day instead of EPR, you would actually have a
lower amount of thrust at 100 percent, Just the
opposite is true on a cold day; desired thrust
ratings may be reached at something less than 100
percent using EPR. Other variables, such as a
dirty or damaged compressor, will also affect
thrust. These conditions reduce thrust for a given
rpm. Exhaust gas temperature (EGT) is NEVER
used for setting thrust. Monitor EGT carefully to
prevent excessive temperature readings.
Accomplish the isolating and correcting
of engine discrepancies by comparing symptoms
with probable cause. Begin with the most
obvious and proceed to the less likely causes.
A combination of several small maladjustments
or malfunctions may contribute to make a
complex discrepancy.
5. Correct the trouble. Accomplish this step
only after the trouble has been definitely located.
Malfunctions are corrected by servicing, adjustments, and part replacement. Part replacement
is often needed to correct a malfunction, but you
There are certain errors that you should avoid
if you are to become an effective troubleshooter.
and recognize normal readings. Learn how to use
test equipment and read schematics before you
are required to troubleshoot a problem.
These errors may result from faulty procedures
or by the way you think about your work. A
report entitled “A Preliminary Investigation of
Troubleshooting” lists common troubleshooting
practices to avoid. The report indicates that you
should avoid making the following common
1. Checking parts of the system that had
nothing to do with the cause of the symptom.
The diligent use of the applicable maintenance
instruction manual (MIM) is essential in preventing poor troubleshooting procedures. The
MIM for each aircraft provides troubleshooting
aids that cover the six steps listed in the previous
section. These manuals provide a variety of
troubleshooting aids. Table 9-2 shows examples
of troubleshooting chart formats. These charts list
a definite sequence of action for a problem,
according to the probability of failure and ease
of investigation. These charts supply the trouble,
probable cause, and the remedy for some of the
more common malfunctions.
2. Ignoring part of a system because you
didn’t know that the unit was a possible cause of
the symptom.
3. Making a difficult check when a simple one
would have done the job.
4. After isolating the engine trouble between
two points, making further checks beyond,
rather than between these points. Sometimes
making a check between two points despite
the fact that no engine trouble was found between
The MIM also contains schematics and
diagrams for use in troubleshooting. Diagrams
of the various electrical circuits, fuel systems,
and lubrication systems are very useful in
isolating discrepancies. There are many types
of diagrams, but those most important for
troubleshooting include schematics and pictorials.
There are some MIMs that combine different
types of diagrams to provide more information.
To understand how a system or component of the
aircraft functions, you must be able to read these
5. Omitting a check even though you know
it is related to the situation.
6. Attempting to remember a lot of information, This finding points out the necessity of
writing down all information during checks, and
using the maintenance manuals, not your
The report pointed out that troubleshooters
sometimes made errors because of previous
experience in similar situations. When they
analyze and repair the engine trouble successfully, they have a feeling of satisfaction.
They then tend to use the same procedures,
even though the procedures are not relevant
to the present engine trouble. Similarly, if
the mechanic tried certain checks in the past
and they did not fix the problem, the mechanic
was reluctant to use them again, even when they
provided required information to correct the
Schematic diagrams enable maintenance personnel to trace the path of electrical circuits, fuel
systems, and lubrication systems of aircraft.
Schematic diagrams use symbols for a graphic
representation of the assembly or system.
Electrical components or circuits use the standard
electrical symbols shown in figure 9-1. Look at
this figure and notice the electrical symbols for
Some of the mechanical symbols used in
schematics are shown in table 9-3. A complete
listing can be found in the Military Standard,
Mechanical Symbols for Aeronautical, Aerospace,
and Spacecraft Use, Part 2, MIL-STD-17B-2.
To become an expert troubleshooter, you must
know the function of each part and how the
operation of one part or system affects another.
A good way to learn engine systems and the
different relationships is to study schematics and
diagrams. Trace out all the systems you work on,
Schematic diagrams show the relationship of
each part with other parts in the system. They do
not always indicate the physical location of the
Table 9-2.—Samples of Various Troubleshooting Charts
Table 9-3.-Aeronautical Mechanical Symbols
Figure 9-1.-Electrical symbols.
9-2.-Fuel control schematic.
part. Figure 9-2 is a schematic diagram that shows
the fuel flow of a turbojet engine fuel control.
It shows the flow of fuel from the fuel pump
through the fuel control, the fuel distributors, and
the internal fuel tubes. Fuel from the internal fuel
tubes enters the combustion chambers through
fuel lines to the fuel manifold. Each part is
illustrated and identified by name. The arrows
show the direction of flow through each line and
component within the fuel control.
Pictorial diagrams show location, function,
and appearance of parts and assemblies. This
diagram is sometimes called an installation
diagram. Pictorial diagrams are valuable for
locating and identifying parts. See figure 9-3.
Other uses for this type of drawing is the
disassembled or exploded view and the cutaway.
Cutaway drawings show the internal construction
of parts, and exploded views show how the
various parts of a component are assembled.
Figure 9-3.-Pictorial diagram of a fuel distributor.
Figure 9-4.-Turbofan, cutaway view.
Figure 9-5.-Turbofan, exploded view.
Figure 9-4 shows a cutaway drawing, and
figure 9-5 shows an exploded view drawing.
You can identify fluid lines in aircraft by
markers made up of color codes, words, and
geometric symbols. These markers identify each
line as to function, content, primary hazard, and
direction of flow. Figure 9-6 lists the various types
Additional information concerning diagrams
is contained in the training manual Blueprint
Reading and Sketching, NAVEDTRA 10077
Figure 9-6.-Fluid line color codes and symbols.
of fluid lines and indicates the color code and
symbol for each line. Figure 9-7 shows the
markers used for rocket fuel lines.
In most instances, fluid lines are marked by
the use of 1-inch tape or decals, as shown in figure 9-7, view A. On lines 4 inches or larger in
diameter, lines in an oily environment, hot lines,
and some cold lines, steel tags replace tape or
decals. See figure 9-7, view B. On lines in engine
compartments, where there is a possibility that
tapes, decals, or tags can be drawn into the engine
intake, paint is used. See figure 9-7, view C.
As shown in figure 9-7, view A, reading from
left to right, the line function is indicated by color
code (red, gray), name (ROCKET FUEL), and
symbols (four-pointed star and crescent). Content
of the line is indicated by name (HYDRAZINE).
The primary hazard is indicated by the word
TOXIC. Pressure is indicated in pounds per
square inch (125 psi). Direction of flow is
indicated by arrows. Two-headed arrows are used
to show reversible flow. In addition, certain lines
may have a special code for lines that have a
specific function within a system. Some examples
include drain, vent, pressure, and return lines.
There are four general classes of hazards
found in connection with fluid lines. These
hazards are shown by special fluid line markings.
1. Flammable material (FLAM). The hazard
marking FLAM is used to identify lines containing
Figure 9-7.-Fluid line identification group (A) using tape and decals; (B) using metal tags; (C) using paint.
abilities since there are few aircraft systems independent of electrical operations. In addition,
many troubleshooting charts call for continuity
checks that a properly trained AD can easily
materials known as flammable or combustible
2. Toxic and poisonous materials (TOXIC).
The word TOXIC identifies lines that contain
materials that are extremely hazardous to life or
3. Anesthetics and harmful materials
(AAHM). The AAHM markings identify lines
containing materials that produce anesthetic
vapors and all liquid chemicals that are hazardous
to life and property. These materials do not
produce dangerous quantities of fumes or vapors.
4. Physically dangerous materials (PHDAN).
The PHDAN marking identifies lines that carry
a material that is not dangerous within itself but
is asphyxiating in confined areas. The material in
the line may be in a dangerous physical state of
high pressure or temperature.
NOTE: Only authorized personnel may
repair and maintain electronic and electrical equipment because of the chance of
injury, the danger of fire, and possible
material damage. To use electrical test
equipment correctly, you need to have a
working knowledge of basic electricity.
You should adhere to the special and
general rules for the handling of electrical
Aircraft and engine manufacturers are
responsible for the original installation of
identification markers, and maintenance personnel are responsible for replacement as necessary.
Tapes and decals are on both ends of a line and
at least once in each compartment through which
the line passes. In addition, identification markers
are found next to each valve, regulator, filter, or
other accessory within a line. Location requirements are the same for tapes and decals as they
are for tags and paint.
Additional information is listed in Aviation
Hose and Tube Manual, NAVAIR 01-1A-20.
Complete instructions for installing fluid line
identification markers are contained in MILSTD-1247B.
Always connect an ammeter in series—never
in parallel. Always connect a voltmeter in
parallel—never in series. Never connect an
ohmmeter to an energized circuit. On test meters,
select the highest range first, then switch to lower
ranges as needed.
When you use an ohmmeter, select a scale that
gives a near midscale reading. Midscale is the
point where the meter is most accurate. Do not
leave the multimeter selector switch in a resistance
position when the meter is not in use. The leads
may short together, discharging the internal
battery. There is less chance of damaging the
meter if the switch is on a high at-volt setting,
or in the OFF position. Meters that have an OFF
position dampen the swing of the needle by connecting the meter movement as a generator. This
prevents the needle from swinging wildly when
moving the meter.
View the meter from directly in front to
eliminate parallax error. Observe polarity when
measuring dc voltage or current. Do not place
meters in the presence of strong magnetic fields.
Never measure the resistance of a meter or a
circuit with a meter in it. The high current required
for ohmmeter operation may damage the meter.
This also applies to circuits with low-filamentcurrent tubes and to some types of semiconductors.
When you are measuring high resistance, do
not touch the test lead tips or the circuit because
it may cause body resistance to shunt the circuit,
causing an erroneous reading. Connect the ground
lead of the meter first when making voltage
You will find that certain test equipment will
prove invaluable in troubleshooting engine
systems. We will briefly discuss the necessary test
equipment available for maintaining aircraft
systems. More detailed information can be found
in the technical manual for the particular piece
of equipment.
During troubleshooting, you will often
measure voltage, current, and resistance.
Although AEs normally troubleshoot and repair
electrical systems, you should have a knowledge
of basic electricity and test equipment. This
knowledge will increase your troubleshooting
The multimeter contains circuitry that allows
it to be a voltmeter, an ammeter, or an ohmmeter.
A multimeter is often called a volt-ohm-milliammeter (VOM).
The VOM does have two disadvantages—it
can load the circuit under test, and the meter
movement is easy to damage because of improper
testing procedures.
Two common multimeter are the Simpson
260 and the newer digital model Fluke 8100A. See
figures 9-8 and 9-9.
measurements. Work with one hand whenever
The process of fault detection often leads
beyond visual inspection and power checks. Use
a voltmeter to find out if a power circuit is delivering power to the proper place. However, the
voltmeter won’t identify what is wrong. An
ohmmeter is a better instrument for identifying
the problem. With the ohmmeter, you can find
opens, shorts, grounds, or incorrect resistance
values. With schematics you can trace circuits
until you isolate the trouble.
Use an ohmmeter to measure circuit continuity
and total or partial circuit resistance. An ohmmeter is shown in figure 9-10. The ohmmeter’s
Figure 9-8.-Standard multimeter.
Figure 9-9.-Digital multimeter.
pointer deflection is controlled by the amount of
battery current passing through the moving coil.
Before measuring the resistance of an unknown
resistor or electrical circuit, the test leads of the
ohmmeter are first shorted together, as shown in
figure 9-11. With the leads shorted, calibrate the
meter for proper operation on the selected range.
Figure 9-10.-Ohmmeter.
Figure 9-11.-Simple ohmmeter circuit.
When the variable resistor is adjusted properly,
with the leads shorted, the pointer of the meter
will come to rest exactly on zero. This reading
indicates zero resistance between the test leads.
When the test leads of an ohmmeter are separated,
the pointer of the meter will return to the left
side of the scale. This happens because of the
interruption of current and the spring tension
action on the movable coil assembly.
After you adjust the ohmmeter for zero
reading, it is ready to be connected in a circuit
to measure resistance. The power switch of the
circuit to be measured should always be in the
OFF position. Circuit voltage applied across the
meter could damage meter movement.
Connect the test leads of the ohmmeter across
(in series with) the circuit to be measured. The
current produced by the meter must now flow
through the circuit being tested. The reading
obtained from the ohmmeter between zero and
infinity shows the number of ohms resistance in
that circuit. An infinity reading shows there is an
open circuit.
In open circuits, current flow stops. The
cause may be a broken wire, defective switch,
etc. To detect open circuits, you must perform
a continuity test, which will tell if the circuit is
complete or continuous.
Look at figure 9-12. It shows a continuity test
of a cable connecting two electronic units. You
can see that both plugs are disconnected, and the
ohmmeter is connected in series with conductor
A, which is under test. The power should be off.
When checking conductors A, B, and C, the
current from the ohmmeter flows through plug
2 (female), conductor A, B, or C, to plug 1
(female). From plug 1, current passes through the
jumper to the chassis, which is grounded to the
aircraft structure. The aircraft structure serves as
the return path to the chassis of unit 2 and
completes the circuit through the series-connected
The ohmmeter shows a low resistance because
no break exists in conductors A, B, or C.
However, checking conductor D reveals an
open. The ohmmeter shows maximum resistance
because current cannot flow in an open circuit.
With an open circuit, the ohmmeter needle is all
the way to the left, since it is a series-type ohmmeter (reads right to left). When you have finished
using the ohmmeter, turn the ohmmeter switch
OFF to conserve the batteries,
High-voltage Insulation Tester
The high-voltage insulation tester detects or
measures leakage of high-voltage insulation.
Figure 9-13 shows one of the several types of
testers now in use.
The high-tension voltage tester is composed
of four sections. They are as follows:
A vibrator invertor, which converts a
24-volt dc source to 110-120 volts ac.
Figure 9-12.-Continuity test.
The insulation tester operates on high
voltages that are dangerous. Always
discharge the ground circuits before you
touch them. Always connect the ground
lead to a good earth ground before you
turn the tester on. When testing a lead,
connect the tester to the lead before you
turn the test voltage switch on. Be safe.
Operate the equipment and secure the leads
with one hand only.
Figure 9-13.—High-voltage insulation tester.
A voltage control, which controls the
amount of high voltage applied to the test circuit.
Two of the most important factors affecting
turbine engine life are engine speed (rpm) and
engine exhaust gas temperature (EGT). Excess
exhaust gas temperatures of a few degrees reduce
turbine blade life up to 50 percent. Low exhaust
gas temperatures reduce engine efficiency and
thrust. Excess engine speed can cause premature
engine failure.
A jet calibration test unit tests tailpipe
temperatures, engine speed, and other engine
parameters more accurately than the cockpit
gauges. Errors in reading cockpit instruments are
made due to the position or height of the seat or
the angle of viewing the gauge. In addition to
more accurate readings, the new jet test units
provide a printed readout of engine conditions for
trend analysis.
A step-up transformer, which steps up the
low voltage from the vibrator invertor or 110-volt
source to high voltage.
A rectifier that changes ac to dc. The test
leads are connected to the rectifier dc output side.
To calibrate the high-voltage tester, follow
these four steps:
1. Set the microampere and kilovolt meters
to indicate ZERO when the machine is OFF.
2. With the current ON, connect the test
3. Move the voltage control until the needle
is about one-sixteenth inch past 10, 000 volts. Then
set the test voltage limiter gap. If it is properly
adjusted, the neon light will not flash at the test
voltage setting.
4. Adjust the excess current indicator to flash
on and off continually when the microampere
meter shows an excess of 100 microampere
You will find borescope inspection requirements listed in the periodic maintenance inspection cards (PMIC) for scheduled or conditional
inspections. Borescopes provide illumination of
internal areas of aircraft engines or engine parts.
This illumination allows for internal inspections
that require minimum disassembly, such as the
removal of port covers or thermocouples.
When operating the high-voltage tester, be
sure to make a continuity test with an ohmmeter
before you apply the high-voltage insulation test.
To hookup the high-voltage tester, follow these
Types of Borescopes
1. Ground the high-voltage insulation tester
to a good local ground with the grounding lead.
2. Connect the red high-voltage test lead to
the high-voltage side of the part to be tested.
The rigid borescope assembly allows for inspection of internal engine conditions having a
direct access. Another type of borescope has a
flexible probe and is commonly known as a fiberoptic borescope, or fiberscope. The fiber-optic
3. Connect the black ground test lead to the
ground side of the part to be tested.
scopes have a flexible probe, which can be snaked
around, behind, and into areas impossible to
reach with the other scope. See figures 9-14 and
The advantage of the flexible probe is
that many areas can now be inspected without
opening the engine. The probe can be worked
back through the compressor from the front of
the engine. It can also be inserted through
a fuel nozzle opening, and worked back to the
turbine area to as close as one-quarter of an
inch from the surface being inspected. Being able
to assess possible damage at that close range has
a distinct advantage over the rigid type of
borescope. A fiber-optic borescope and an
inspector experienced in its use could make the
difference between the man-hours spent rejecting
a perfectly good engine or obtaining more flight
hours from that same engine after damage was
BoreScope Use
Know your equipment. Borescopes are easily
damaged and have different characteristics. For
example, optical characteristics of the small
borescope magnify and distort all areas other than
the turbine blade leading edge. This often results
in what appears to be an alarming vane condition. Know and be able to locate all inspection
areas or ports on your engine. Figure 9-16 shows
an example of inspection ports and areas that you
can inspect from that port.
Figure 9-14.-Rigid borescope.
Figure 9-15.-Fiber-optic (flexible) borescope.
Figure 9-16.-Engine inspection ports.
Figure 9-17.—Borescope view of thermocouple and turbine blades.
Figure 9-18.—Borescope view of turbine blades.
including stage, area, magnitude, direction, and
adjacent material conditions. Some inspection
cards instruct you to mark or bend an 18-inch
length of lockwire, and then insert the wire
through an adjacent inspection port. You can then
use this wire as a gauge to measure vane cracks
while viewing through the borescope.
Figures 9-17 and 9-18 show different engine
parts seen through a borescope. You must know
the words to describe the damage found. Then
compare this information with serviceability
limits in the maintenance instruction manuals. See
table 9-4.
Establish internal reference points. Pictorial
and cutaway diagrams help establish these
reference points. When the probe is in the inspection port, it is easy to lose your sense of direction. Some borescopes have an index mark on the
eyepiece to show direction.
Scan the inspection area thoroughly and
in an orderly manner. Compressor cracks can be
difficult to detect because of small close parts,
rotating the engine too fast, or because of
deteriorated borescope optics. Note and evaluate
any inconsistencies. Record all information,
Table 9-4.-Examples Of Serviceability Limits
Report your conclusions and recommendations. Carefully read and follow the procedures in your aircraft maintenance manuals
for borescoping. Different procedures are
required for different types of engines to
prevent damage or to inspect properly. For
example, on the turboprop engine, the turbine
blade inspection is made easier by releasing
the propeller brake. This makes it possible
to rotate the turbine smoothly in both directions.
On turbofan engines, you should not borescope if there is a possibility that wind can
rotate the fan rotor.
Water washing consists of an emulsion of
demineralized water and cleaning liquids. Field
cleaning of this nature is accomplished as a
desalination wash to remove salt deposits when
operating in salt-laden air. Field cleaning maybe
done as a performance recovery wash to remove
dirt and other deposits.
Specific steps to follow in cleaning a particular
engine are found in the maintenance instructions
for the particular engine. These steps generally
include blocking some lines and ports and
removing any equipment in the inlet duct. This
is done to prevent damage by the cleaning
material. The engine is then motored over for
specific periods of time while the cleaning
compound is fed into the inlet duct. After cleaning, the engine must be returned to its original
configuration. The engine is dried by running it.
Engine performance for trend analysis is often
checked after engine washing.
Although water washing is not required
before borescoping, it helps clean the engine so you get the best possible evaluation.
Dirt and small carbon particles may obscure
small cracks or pitting that could be missed
in a dirty engine. Water washing also improves engine performance by removing sulfidation.
After completing this chapter, you will be able to:
Recognize the types of repairs accomplished
at the intermediate maintenance level.
Identify the different types of test cells and
their components.
Identify the repair limits for the intermediate
level of maintenance.
Recognize the purpose and entries on the
engine test log sheets.
Identify the different methods of cleaning and
marking engine parts.
The purpose of this chapter is to familiarize
you with the types and authorized repair limits
for intermediate maintenance activities (IMAs).
Intermediate maintenance applies to those maintenance functions normally performed in centrally
located facilities for support of the operating
units. The facilities are designated as aircraft
intermediate maintenance departments (AIMDs).
The primary purpose of the IMA is to
support and supplement the work of organizational maintenance activities. Squadron personnel
assigned to the IMA, ashore or afloat, are
assigned to perform the total work (within their
skills) of the intermediate activity and not just the
work related to support of the squadron from
which they were assigned.
different engines used in naval aviation, the
maintenance procedures in this chapter are
general in nature. Components and repair
limits discussed are representative. Do not refer
to them when working on a specific engine or its
The Gas Turbine Engine Maintenance Program was formed under the three-degree concept
as specified in OPNAVINST 4790.2 (series).
Under this concept, each engine’s intermediate
maintenance manual defines specific engine
maintenance actions as either first-, second-,
or third-degree functions. These functions are
determined largely by degree of difficulty and
frequency of repair.
The Gas Turbine Maintenance Program
defines the repair functions of AIMD power
plants. Repair capabilities are different for each
particular engine and AIMD. It is important that
you become familiar with the repair capabilities
and functions of your intermediate maintenance
First-degree repair is the repair of a damaged
or nonoperating gas turbine engine and its
accessories or components. When the compressor
This chapter covers some of the procedures
and equipment used in an intermediate maintenance department. Because of the number of
rotor is replaceable, the repair includes compressor rotor replacement and/or disassembly.
NOTE: The terms “first-degree repair”
and “complete engine repair” (CER) are
synonymous. CER is used primarily with
older engines that are not included in the
three-degree program.
applicable MIMs or disassembled for further
Many of the chemical solutions and their
components used in cleaning, inspection,
and repair are toxic, flammable, and
extremely corrosive. Improper mixing or
use of these chemical can produce violent
reactions, rapid heat generation, and
explosive/toxic gases. Personnel performing maintenance procedures should consult
the applicable maintenance instruction
manuals and be familiar with the safety
precautions associated with the hazardous
materials or equipment. The warnings in
these technical manuals identify the types
of hazards and precautions to take. The
Material Safety Data Sheets (MSDS) and
your safety office have specific information for hazardous material in your work
Second-degree repair is also the repair of a
damaged or nonoperating gas turbine engine and
its accessories or components. The difference is
that second-degree repair will normally include
the repair/replacement of turbine rotors and
combustion sections. Repairs include afterburners
and the replacement of externally damaged,
deteriorated, or time-limited components, gearboxes, or accessories. Minor repair to the
compressor section is made in second-degree
repair. The repair or replacement of reduction
gearboxes and torque shafts of turboshaft engines
comes under second-degree repair. The repair or
replacement of compressor fans of turbofan
engines also comes under second-degree repair
Good mechanics clean all engine parts
thoroughly before inspecting them. Cleaning and
close inspection make it possible to detect
faults that endanger safe engine operation and
maximum performance. The primary purpose of
engine parts cleaning is to accomplish the
The third-degree repair encompasses major
engine inspections and the same gas turbine engine
repair capability as second-degree maintenance.
Certain functions that require high maintenance
man-hours and are of a low incidence rate are
excluded. The functions described represent broad
generalities. Refer to the appropriate engine
maintenance plan or intermediate maintenance
manual to determine the degree of assignment for
specific repair functions.
1. Permit thorough inspection of components
for flaws, damage, and dimension wear.
2. Prepare surfaces for repair (plating,
welding, or painting).
3. Remove organic or inorganic coatings for
inspection of underlying surfaces or remove
coatings adversely affecting engine performance.
Selection of cleaning materials and processes
for any engine part is determined by the nature
of the soil, the type of metal or coatings, and the
degree of cleanliness necessary for a thorough
inspection and repair.
Generally, engine parts operating in relatively
low temperature ranges (cold section parts) are
cleaned by solvent washing, decreasing tanks, and
vapor decreasing. Cleaning engine parts that
operate in hot sections of the engine (combustion
and turbine sections) require more comprehensive
Once an engine arrives at an AIMD activity,
it is cleaned and evaluated for repair. If inducted
for repair, a major inspection and all repairs
required to place the engine back in ready-forissue (RFI) status are accomplished. The first
steps for inspecting aircraft engines include the
cleaning and the marking of parts. After cleaning, engines are inspected in accordance with
When cleaning plastic parts, you should be careful
to avoid heat buildup. After cleaning thoroughly,
dry the part and apply corrosion prevention
cleaning. This cleaning includes using a heavyduty alkaline cleaner and blasting.
Soft carbon deposits are removed by decreasing and steam cleaning. Decreasing removes
dirt and sludge by immersing or spraying the part
with cleaning solvent. Hard carbon deposits are
removed by decarbonizing, brushing, scraping, or
grit-blasting. The following text provides general
cleaning procedures to familiarize you with the
methods and materials used for cleaning parts.
Always refer to the appropriate maintenance
manual for the latest cleaning procedures.
Constant changes are made in cleaning and
finishing (coatings) materials.
Vapor Decreasing
Vapor decreasing removes oil, grease, and
preservative compound by solvent vapor. A flat
bottom tank with heating coils on the bottom and
cooling coils midway around the tank is required.
The part is suspended in the vapor area below the
cooling coils. Heated cleaning solvent vapor
condenses on the cool part. It dissolves oils,
grease, and preservatives. Cleaning action stops
when the part reaches the vapor temperature. If
further decreasing is necessary, the part must be
cooled before using vapor degreaser again. Vapor
decreasing cannot be used on titanium parts
because recommended solvents cause stresscorrosion at high temperatures.
Recommended solvents for vapor decreasing
are trichloroethane (O-T-620), trichlorotrifluroethane (MIL-L-81302), and perchloroethylene.
These solvents are designed to clean metals by the
vapor cleaning method because their high vapor
density results in small vapor loss.
Degreasing (Solvent Cleaning)
Small accumulations of grease, oil, and dirt
may be removed by hydrocarbon solvent cleaning.
This method is not effective in removing bakedon oil deposits or most surface coatings.
Decreasing solvents are flammable, and
their vapors are toxic. Keep all solvents
away from open flames, and use only in
well-ventilated areas. Avoid solvent contact with skin, eyes, and clothing by wearing rubber gloves, a face shield or goggles,
and an apron or coveralls.
Decarbonizing is the chemical removal of
carbon deposits. Decarbonizing agents are
detergents, sodium silicates, chlorinated hydrocarbons, and various acid solutions. This cleaning
method is effective for paint stripping, rust
removal, and general cleaning of ferrous and hightemperature parts. Parts are soaked in hot or cold
tanks and rinsed with high-pressure water.
Dry-cleaning solvent (P-D-680) is the recommended cleaner. Solvent cleaning uses a tank with
cleaning solvent to soak the part clean. Some
tanks have pumps to provide mechanical agitation to help clean parts. You can also use a softbristle brush to remove stubborn stains. The
cleaning tank should have a hinged, counterweighted cover so it can be covered when not in
use. Since some plastic- and rubber-based
materials are attacked by hydrocarbon solvents,
steam clean these parts to remove contaminants.
Carbon removers require careful handling.
Wear goggles, rubber gloves, and aprons
when using these solutions.
Steam Cleaning
Some carbon removers attack aluminum and
magnesium parts if they are left in the solution
too long. There is also the possibility of a chemical
reaction when aluminum, magnesium, and steel
parts are immersed in the same tank. This practice
often results in damage to magnesium parts, such
as dissimilar metal corrosion.
Upon removal from cleaning solutions, rinse
the parts in a soap-and-water solution or with a
Steam cleaning is a cleaning process used when
you do not want to remove paint and surface
coatings. To properly clean with steam, it is
necessary to add cleaning compounds. Do not
steam clean oil-impregnated parts.
Set steam valve to the proper strength and
force required for the job. Hold the steam gun
about 12 inches from the part at a 45-degree angle.
or temporary. Permanent markings are those
markings that remain during the entire service
life of the part. They are applied during manufacturing or after modification of parts.
Temporary markings maintain identification of
parts or reference positions during ordinary
handling, storage, and final assembly. Temporary
markings ensure parts may be returned to original
assembly position. If a part is going to be cleaned,
inspected, and repaired, temporary markings will
probably be removed by solvents during cleaning.
If part identification needs to be maintained,
attach tags or place parts in separate containers.
petroleum solvent. Change the rinse water
frequently to prevent a buildup of acid or alkaline
in the water. Air-dry the engine parts, and then
coat them with a corrosion preventive if they are
not to be processed further.
Decarbonizing loosens most hard carbon
deposits remaining after decreasing. The complete
removal of all hard carbon deposits generally
requires brushing, scraping, or grit blasting. Use
caution during these procedures to avoid
damaging parts. In particular, do not use wire
brushes or metal scrappers on machine surfaces
or bearings.
Temporary Markings
Abrasive Blasting
Certain materials must be used for temporary
marking during assembly and disassembly. Use
only approved pure dye markers to mark engine
hardware. Using nonapproved markers can leave
harmful elements on engine parts. You may
use layout dye (lightly applied) to mark parts
that are directly exposed to the engine gas path.
Some exposed items are the turbine blades
and disc, turbine vanes, and combustion chamber
Use abrasive blasting to remove hard carbon
or lead deposits, rust, and heat scale. The
type (wet or dry) and size of abrasives vary
for different engine parts. Mask all openings,
identification markings, and other areas as
required before blasting. Grit materials such
as ground corn, apricot or peach pits, walnut
shells, clover seed, and cracked wheat or rice are
in general use.,
NOTE: Do NOT use any temporary
marking method that leaves a heavy carbon
deposit. Do NOT use any marking that
leaves a deposit of copper, zinc, lead, or
similar residue, such as pencil or black
grease pen. These deposits may cause
carbonization or intergranular attack when
the part gets very hot. Parts marked with
unauthorized materials should have all
traces of markings removed before using
Dry grit blasting is sometimes done in a sandblast cabinet. The part must first be degreased or
put through a decarbonizing solution, and then
rinsed and dried thoroughly. After grit blasting,
remove the dust by air blasting, and clean with
petroleum solvent or hot water. Some types
of soft grit leave a light grease film on the
part. Remove this film by decreasing if the part
is to be subjected to fluorescent penetrant
Wet abrasive blasting is an effective method
to remove heat scale, carbon deposits, and
temporary markings, and to produce a uniform
satin finish on metals. This type of blasting
removes metal, but so slowly that dimensions
change very little.
Permanent Markings
Permanent markings should be positioned in
the area of lowest stress. Do not apply markings
within 0.030 inch of corners, radii, fillet, or edges.
Choose an area where markings will not be worn
off or obliterated by contact with another
part. If possible, place new markings next to
original markings. Always refer to applicable
engine manuals and power plant changes for
recommended marking met hods and details.
Some of the methods of permanent markings
include using a metal stamp, vibropeen, blasting,
and acid-etching.
Marking engine parts and assemblies aids in
identification, reassembly, and tracking the
service life of parts. All marks are applied to
produce maximum legibility and durability
without affecting the function or serviceability
of the part. Markings are either permanent
Most intermediate-level maintenance activities
are responsible for the replacement of compressor
sections and the repair of those compressor blades
in the later compressor stages that cannot be
serviced without the removal of the compressor
halves. Compressor repair usually results from
foreign object damage (FOD), although other
failures occur. A number of compressor failures
may be broadly classified under air leaks and
compressor contamination. The AIMD may also
be required to modify the compressor rotor or
stator blades as a result of a technical change or
bulletin. Modifications may include reworking the
components, changing blades by stages, or
replacing the entire rotor assembly.
Before starting disassembly of an engine,
check the applicable technical instructions to
confirm the scope of repair necessary. Use
the step-by-step procedure and repair limits
contained in the current technical publications
for the particular engine. In the process of
engine disassembly, many associated parts
become accessible for inspection. Inspect these
parts as closely as possible to prevent later
Support the engine during disassembly and
buildup to prevent overstressing of the parts being
removed. In addition to preventing stress, it allows
proper alignment of parts being removed or
NOTE: Compressor cases are machined in
matched sets. Damage beyond repair to
one case is cause for rejection of the
opposite case. A new compressor rotor is
not required when replacing the entire case
Immediately upon removing each subassembly
or individual part from the engine, transfer
it to a parts rack. Arrange the part to protect
it or the assembly from damage. Provide proper
covering and supports to protect shafts, gears,
studs, or any projecting piece from being bent,
scratched, or otherwise damaged. Be careful to
prevent the entrance of dirt and other foreign
materials into the engine. Whenever practicable,
use temporary covers to seal all openings in
dismantled engines. Cover the ends of all removed
tubing. Take extreme care not to lay carbon
seals and plates on the sealing surfaces. Provide appropriate containers to hold each part
separately until the time for reinstallation.
During disassembly, examine all parts and
assemblies for cracks, scoring, and burning.
Check the engine for indications of work incorrectly performed during any previous repair
or overhaul.
Compressor Contamination
Accumulation of dirt on the compressor
blades reduces the aerodynamic efficiency of the
blades. Dirt hurts engine performance. The
efficiency of the blades is impaired by dirt deposits
similar to that of an aircraft wing under icing
conditions. High exhaust gas temperature (EGT)
may result when foreign deposits are on
compressor components. On some turbojet
engines, high EGT requires early engine overhaul.
Slow acceleration could also result from foreign
material obstructing the compressor outlet vanes.
This obstruction could result in a needless engine
You must pay particular attention to the
material requirements for the nuts and bolts used
in the turbojet engine. Hot sections require
common hardware (nuts, bolts, safety wire and
cotter pins) that are resistant to high operating
temperatures. Other engine parts may require
special hardware, and it is imperative that all
properly coded parts (if serviceable) be placed in
their original positions.
Compressor Leaks
Air leaking from the compressor results in low
engine performance. Air leakage may occur
between the high and low compressors, or at some
intermediate stage. It may also occur because of
bleed-air valves stuck open or cracks in the
compressor case itself. Air leaks in the compressor
are found through engine monitoring or low
engine performance; for example, when the engine
fails to meet minimum power requirements for
Welding is permissible on some parts of the
turbojet engine. Refer to the applicable technical
instructions before attempting to repair an engine
or component by welding.
Compressor Failures
Loose objects often enter an engine’s compressor
either accidentally or through carelessness. Thousands
of dollars worth of damage to a compressor rotor can
result from pliers being left in the air intake, as shown
in figure 10-1. The nut and bolt holding the pliers
together came loose and went through the compressor,
causing the damage shown in the illustration. A simple
solution to the problem of tools drawn into an engine is
to check the tools against a tool checklist.
Internal mechanical failures, such as a compressor
blade breaking off, result in compressor efficiency loss.
These failures are difficult to detect. Broken blades and
vanes result in high exhaust gas temperatures or an
increase in compressor rpm due to loss of efficiency. Of
course, mechanical failures of compressor blades could
result in severe damage to the compressor, combustion
chamber, and the turbine as foreign object damage
Compressor Blade Damage
and Repair
Figure 10-2.-(A) Blade showing dent and galling;
(B) blade showing a score; (C) blade showing
scratches; (D) blade showing gouge and burr.
Defining and showing examples of damage may help
you to recognize damage and make
Figure 10-3.-(A) Blade showing corrosion
(pitting); (B) blade showing cracks; (C) blade
showing burn; (D) blade showing damage and
repair (blend).
Figure 10-1.-Compressor damage caused by
pliers passing through compressor.
Table 10-1.-Blade Maintenance Terms
You may make minor repairs to compressor
blades, provided the repairs are made without exceeding allowable limits in the prescribed MIM. Well-rounded damage to leading
and trailing edges is acceptable without rework. No rework is necessary provided the
damage is in the outer half of the blade.
The indentation may not exceed values specified
repair of damage easier. Figures 10-2 and 10-3
show examples of possible blade damage to
an axial-flow engine. The damages are enlarged to show detail. See table 10-1 for the
definitions or descriptions and possible causes of
blade damage. If you know by close examination
what causes the damage, you may be able to
reduce the cause.
emery cloth. Use a medium stone for areas
containing small nicks and dents.
in the MIM. Figure 10-4 illustrates representative
repairable limits. Figure 10-5 shows examples of
blade repairs.
Use fine emery cloth and/or a fine abrasive
stone to polish the reworked area. Polish until the
finish looks and feels like the original. If two
blended areas overlap to form a sharp point or
ridge, blend out the point or ridge. Blend the
contour surfaces with a medium stone and finish
with emery cloth and/or a fine abrasive stone. The
finished repair should be as much like the original
finish as possible.
When working on the inner half of the airfoil, you should treat damage with extreme
caution. Make no attempt to remove damage by
straightening. Inspect repaired compressor blades
by dye check, magnetic particle inspection, or
by fluorescent penetrant inspection methods.
Remove all traces of the damage. All surfaces
must be smooth. All repairs must be well blended.
No cracks of any extent are tolerated in any area.
Bowed or bent blades are not reused. If gauges
are not available, the repaired blades are aligned
and compared with a new blade of the same stage.
Front compressor blades that require replacement are replaced by blades having the same
moment balance code. The moment balance codes
are marked on the front face of the root of the
blade. At the original buildup of the compressor
rotor discs and blades, a set of blades coded
according to individual moment were installed.
The installation on the disc minimizes the static
imbalance due to variations in the blade moment.
The blades are numbered in clockwise sequence,
Use a smooth file when removal of considerable amounts of material is necessary. File
or blend at right angles to the width of the blade.
When you cannot reach the damaged area with
a file, use course emery cloth. Use a medium stone
on areas that have been reworked with a file or
Figure 10-4.-Repairable limits and examples of maximum repair.
Figure 10-5.-Compressor blade repair.
compressor stator. The inspection and repair of air
inlet guide vanes and swirl vanes on engines require
the use of a strong light. Attach this light to a rigid
support to enable positioning in hard-to-reach areas.
Inspect all sections of both screen assemblies for
breaks, rips, or holes. Screens may be tin-dipped to
tighten up the wire mesh, provided the wires are not
worn too thin. If the frame strip or lugs have
separated from the screen frames, rebrazing may be
as viewed from the front. Install blades in their
correct numbered position. Make sure new blades are
correctly numbered for the blades they replace.
NOTE: The blade part number follows a
change designation number. Never take
this number as the moment weight
code of the blade. The code letter i s
stamped adjacent to, but not following,
the part number.
Compressor Stator Vanes
Inspect guide and swirl vanes for looseness.
Inspect the outer edges of the guide vanes. Pay
particular attention to the point of con-tact between
the guide and swirl vanes. Check for cracks and dents
due to the impingement striking of foreign particles.
Also, inspect the edges of the swirl vanes. Inspect the
downstream edge of the guide vanes very closely, as
cracks are more prevalent in this area. Cracks which
branch in such a manner that a piece of metal could
fall into the compressor are cause for rejection.
Pitting or corrosion, if within the allowed
tolerance, is not serious on the compressor stator
vanes of axial-flow engines. Do not attempt to repair
any vane by straightening, brazing, welding, or
soldering. Use crocus cloth, fine files, and stones to
blend out damage. Remove a minimum of material,
and leave a surface finish comparable to that of a new
part. The purpose of this blending is to minimize
stresses, which concentrate at dents, scratches, and
cracks. Maximum blend out limit damage greater
than 50 percent of the stator assembly vanes requires
assembly replacement. When vanes are damaged, to
the maximum blend out limit in any 60-degree sector
of one stage, replace the assembly. Send parts
damaged beyond maximum repair limits to overhaul
for repair and replacement of vanes. The use of
prebored compressor vane and shroud assemblies
permits the replacement of one-half of any low-
Blending of hollow vanes on the concave and
convex surfaces, including the leading edge, is
limited, Some small, shallow dents are acceptable.
The damage may be rounded or gradual contour type,
but not a sharp or V-type. Do not allow any cracked
or torn vane material to exist in the damaged area.
Blend the trailing edge if one-third of the weld
figures 10-6 and 10-7. Concave surfaces of rubberfilled vanes may have allowable cracks extending
inward from the outer airfoil. These cracks are
allowable provided there is no suggestion of pieces
breaking away. Using a light and mirror, inspect
each guide vane trailing edge and vane body for
cracks and damage caused by foreign objects. Cracks
in the vane body are cause for rejecting the entire
The combustion section can be removed,
repaired, or replaced in part or entirely depending on
the extent of damage encountered. The combustion
section consists of liners, support duct, outer and
inner case, and the first stage turbine nozzle
assembly. Most repairs to this section are
accomplished by welding or replacement of
Combustion Chamber Liners
Inspect combustion chamber liners for cracks by
using dye penetrant or the fluorescent penetrant
method of inspection. Cracks con-verging so that
metal could break loose or any loose, cracked, or
damaged swirl vanes are cause for rejecting a liner.
Remove liners having buckled areas in a weld seam.
Areas that have more than a three-sixteenths-inch
wave, which does not include the weld seam, require
removal from service. See figure 10-8.
Figure 10-6.-Guide vane trailing edge before
Combustion chamber liners may be retained in
service with some flaws. For example, liners with
cracks less than 0.125 inch long starting from
combustion air holes (no more than three per section) may remain in service. Liners with radial or
circumferential cracks less than 0.750 inch long
extending from or around the crossover tubes or
igniter plug bosses may also remain in service. You
may reuse combustion liners that are burned, except
when burned in the area identified in item 5 of figure
10-8. Rework cracked deflectors to remove the
cracked area by blending or cutting before reuse.
Burning of the cooling louver must not exceed two
tabs totally burned or a total area of two tabs per
Combustion Chamber Support
Figure 10-7.-Guide vane trailing edge after
Cracks in the combustion chamber support can
be repaired by inert-gas fusion welding.
Figure 10-8.-Combustion chamber liner limits.
Repairs must not result in distortion or misfit of
parts at assembly. If a hole is distorted as a result
of welding, use a file to restore it to its original
configuration. The weld bead must be blended by
hand after the welding. See figure 10-9.
1. Maximum length of crack 2 inches.
2. Cracks, maximum of 33 percent and not more
than three adjacent holes.
Combustion Chamber
Inner and Outer Ducts
Operational cracks in the inner and outer
outlet ducts are permissible. The cracks may not
3. Only one crack per section permitted, and it must
be welded.
Figure 10-9.-Combustion chamber support limits.
exceed 3 inches in length. Repair cracks more than
3 inches long and exceeding 75 percent of the
entire circumference of the duct by gas fusion
welding. When repairing a duct, use a silicone
carbide grinding wheel. Grind a 90-degree “V”
groove, 0.035-inch deep, for the entire length of
the crack. It is permissible to remove some of the
parent material when grinding. Adequate depth
is necessary to be sure that the grinder removes
all material from the crack. Repairs must not
result in distortion or misfit of the parts at
assembly. File out any distorted hole, as a result
of welding, to restore it to the original configuration. See figure 10-10 for an example of inner and
outer ducts repair limits.
Inspect turbine sections and components
thoroughly because of the extreme heat encountered in this section. Turbine sections can
be replaced in whole or in part. The turbine rotor
is usually repaired by changing individual blades
or an individual rotor. It is not feasible to describe
all of the damage conditions that may be found
in the turbine. If the damage is within prescribed
limits, but there is doubt about the rework, you
should replace the turbine rotor or part.
Some first-stage turbine blades are coated with
a protective coating to prevent sulfidation. Alpak
is normally used for coating these blades. Without
Figure 10-10.-Combustion chamber inner and outer duct.
Figure 10-11.—Examples of unacceptable sulfidation of turbine blades and vanes.
this coating, these blades are very susceptible to
by a rough or crusty appearance at the leading edge,
on the concave side of the airfoil section, or on the
platform at the root of the air-foil. The rotor should be
replaced if there is evidence of splitting,
delamination, separating, flaking, or loss of material
in any area of the blade. Figure 10-11 shows an
example of unacceptable sulfidation of turbine blades.
Turbine Blade Sulfidation
Sulfidation is high-temperature corrosion.
Sulfidation starts with the excessive levels of sodium
and sulfur in the air and fuel mixture entering the
engine. This type of environment attacks turbine
blades and stator vanes.
Sulfidation first appears
as a rough or crusty surface on the leading edge and
concave surface of the airfoil. It progresses to scaling,
splitting (delamination), and eventual metal loss. The
sulfidation process accelerates with an increase in
sulfur intake and an increase in engine operating
All blades should be inspected for sulfidation.
This form of corrosion is permissible if evidenced only
Turbine Blades
You may inspect turbine blades on axial-flow
engines, and clean them in the same manner as
compressor blades. However, because of the extreme
heat under which the turbine blades operate, they
are more easily damaged. Inspect the turbine blades
for stress rupture cracks and deformation of the
leading edge. See figures 10-12 and 10-13.
Stress rupture cracks usually appear as fine
hairline cracks. These cracks are found on or across
the leading or trailing edge at a right angle to the
edge length. Visible-cracks may range-in length from
one-sixteenth inch upward.
Deformation, due to overtemperature, appears as
waviness along the leading edge. The leading edge
must be straight and of uniform
Figure 10-12.-Stress rupture cracks.
Figure 10-13.-Turbine blade waviness.
thickness along its entire length, except areas
repaired by blending. Stress rupture cracks or
deformation of the leading edge are often
mistaken for foreign material impingement
damage. When any stress rupture cracks or
deformation of the leading edge of the first-stage
turbine blades is found, suspect an overtemperature condition. Check the individual blades for
stretch, and the turbine disc for hardness and
In the following blade removal and replacement procedures, read the numbered steps while
referring to figure 10-14.
NOTE: Before removing any blades,
match mark all blades and baffles. Install
the original or its proper replacement in the
original position.
Steps 1 through 3 refer to figure 10-14, view A.
1. Locate and mark damaged blade.
2. Straighten locking strip tab of damaged
3. Straighten locking strip tab of two
adjacent blades.
NOTE: Install blades removed for detail
inspection or for check of disc stretch in
the same slots from which they were
removed. Number the blades before
Steps 4 through 6 refer to figure 10-14, view B.
Inspect the turbine blade outer shroud for air
seal wear. If you find evidence of shroud wear,
measure the thickness of the shroud at the worn
area. Use a micrometer (or another suitable
measuring device) to be sure of a good reading
in the bottom of the narrow wear groove. If the
remaining radial thickness of the shroud is less
than specified in the appropriate technical
manual, replace the stretched blade.
4. Slide leading blade, baffle, and trailing
blade from slot; weigh damaged blade.
5. Remove remaining loose blades from
6. Remove and discard locking strips. Do
not reuse locking strips.
Steps 7 through 9 refer to figure 10-14, view C.
7. Place new locking strip in blade dots.
8. Position baffles and trailing and leading
blades into slot.
9. Slide leading blade, baffle, and trailing
blade into the last two serration.
Turbine Blade Replacement
There are two categories of moment-balanceweight classifications for turbine blades in axialflow engines. These are individually classified
blades and sets (two or three) of matched blades.
The first category includes individual blades with
permanently marked moment-balance-weight
classifications. The second category includes sets
of blades with the same classifications and
temporary markings.
Figure 10-14, view D, is a hydraulic tab bender
and consists of a hydraulic foot pump and
adjust able tab-bending adapter.
Step 10 refers to figure 10-14, view E.
10. Prebend locking strip tabs using special
tool provided or a common screwdriver. Set
adjusted tab bender onto tab to be bent. Apply
hydraulic pressure with foot pedal until tab is set.
Remove tab bender and check clearance, as shown
in figure 10-14, view F.
NOTE: Never mix blades. Classification of
the two categories does not follow the same
schedule. Always refer to the Turbine
Rotor Disc Assembly Service Record to
determine which category applies.
Turbine Stator Vanes
If visual inspection of the turbine assembly
shows that blade replacement is necessary, use the
following example procedures as a guide. Replace
any number of blades through 100 percent in the
turbine rotor. Replacement is permissible if the
pan weight of the replacement is within 1 gram
of the old blade.
The inspection of turbine vanes is accomplished in the same manner as described in the
previous section for the compressor vanes.
Damage may be greater as turbine vanes
experience extremely high heat. Blend minor nicks
and dents with fine stones or emery cloth. Visually
Figure 10-14.-Turbine rotor blade removal and replacement.
inspect the vanes for signs of bowing. Replace
bowed vanes that are in excess of the allowable
clearance with vanes of the same classification.
Use a straightedge and feeler gauge stock to
measure bowing. See figure 10-15. Damage that
does not crack the metal or reduce the vane
thickness by more than allowed is acceptable
without rework. This is true if the damage is of
a gradual contour shape and not sharp or
V-shaped. Round bottom dents on the leading
edge may be acceptable without rework. Dents
must not exceed allowable limits in depth and
must not crack or tear the vane. Blend sharp
indentations to remove stress concentration.
Determine how many vanes you need to
replace on the particular engine before you
measure the nozzle guide vane area. If the
replaced blades exceed the allowed number and
you must measure the area, it is better to replace
the assembly.
One of the final procedures in the maintenance
of the turbine section of a turbojet engine is to
check for clearances. The service instructions
manual gives the procedures and tolerances for
checking the turbine. Figures 10-16 and 10-17
show clearances being measured at various
locations. Use special manufacturer’s tools to
obtain accurate readings. Also, use the tools
described in the service instructions manual for
specific engines.
The exhaust section of the turbojet engine is
very susceptible to heat cracking. Inspect this
Figure 10-16.-Measuring the turbine blade to shroud tip
section thoroughly, along with the combustion
section and turbine section of the engine. The
exhaust section of an afterburning engine is
subject to extreme heat and pressures. Inspect the
external area of the exhaust cone and tail pipe for
cracks, warping, buckling, and hot spots. Hot
spots on the tail cone are a good indication that
a fuel nozzle or combustion chamber is not
functioning properly. If there is an afterburner,
inspect the afterburner flap segments for burning, warping, or misalignment. Figure 10-18
shows an afterburner duct and nozzle assembly
with flap segments closed. Inspect the afterburner
synchronizing gear segments for worn and missing
teeth and security of installation. Inspect the
nozzle actuation pistons for cracks and/or bent,
chafed, or scored piston rods. Inspect the roller
guides for warpage and the turnbuckle for security
of installation.
Inspect the internal area of the exhaust pipe
and afterburner for evidence of hot streaks,
buckling, and warping, including the flame
holder. Also, inspect all external fuel and
hydraulic lines for evidence of distortion,
buckling, or leakage. Accomplish the repair and
replacement of parts of the exhaust section using
Figure 10-15.-Checking vane bowing.
partial assembly of the parts, and will be
accomplished during assembly of the components.
For a jet aircraft engine to operate properly, it
must be adjusted properly. The checks for proper
adjustment can best be made during controlled
operation of the engine. The various types of test
facilities, or test cells, are designed for service testing
of the jet engine according to procedures and manuals
published by the Naval Air Systems Command
(NAVAIRSYSCOM). Most engine test cells are used
at the intermediate-or depot-level maintenance
Operators of these test facilities are required by
OPNAVINST 4790.2 (series) to be certified by
completion of at least one of the following three basic
methods of operator training:
Figure 10-17.-Measuring the turbine wheel to
exhaust cone clearance.
1. Completion of formal training at a Naval Aviation
Depot (NADEP) school, based on the specific model of
test facility.
the latest technical instructions for that particular
2. Completion of on-site training provided by a Naval
Air Engineering Service Unit (NAESU) engineer.
3. Completion of on-the-job training (OJT) under the
direct supervision of a senior petty officer or civilian
technician who is a certified test
Because of the high rpm, main engine bearings
are critical sections of an engine. The number of main
engine bearings may differ from one model engine to
another. With the engine disassembled and in the
vertical position, all bearings and housings can be
inspected and replaced as necessary.
The accessories drive section contains the
various gearboxes for driving the accessories. These
gearboxes should be inspected for cracks and worn
areas, and the splines should be checked for proper fit
and clearances. When you remove or replace
gearboxes, the splined drive shafts must be carefully
removed or installed to prevent spline damage.
Most mating gears require backlash as well as
end clearance checks. You should carefully inspect
the gear and spline teeth for irregular or excessive
wear, galling, and flaking. Usually, runout
measurements are not required if there is no evidence
of gear teeth spalling. These checks may require
Figure 10-18.-Afterburner duct and nozzle
stand operator and who has been designated by
the activity’s commanding officer to provide this
All certified test cell operators must hold a
valid support equipment (SE) license and ensure
that each particular engine and engine test system
is indicated on the license. Refer to OPNAVINST 4790.2 (series) for training and licensing
Engine testing is accomplished primarily in a
test cell or test house that is fully equipped to
measure all of the desired engine operating
parameters. The building is usually of concrete
construction and contains both the control and
engine rooms, although in some installations only
the control or the instrumentation room is
enclosed. Most of these cells have noise silencers
installed in the inlet stack for noise suppression
and a water spray in the exhaust section for
cooling. Many of the test cells use computers to
automatically record all instrument readings and
correct them to standard day conditions. A typical
enclosed test facility is described in the following
paragraphs. Portable universal engine runup test
systems and the engine test log sheets are also
The test equipment is capable of handling
30,000 pounds of jet thrust during performance
tests. The test facility building configuration has
an engine mass airflow capacity of 180 pounds
per second, or approximately 17,000 pounds of
thrust, including afterburner operation.
The complete test equipment consists of nine
interconnected major assemblies. These
assemblies are installed in the test facility building.
The test facility building is constructed of
reinforced concrete throughout. Removable concrete panels are provided in the primary intake
and exhaust stack systems to permit airflow and
acoustic expansion of the test facility to a
30,000-pound thrust capacity.
The test cell consists of horizontal primary and
secondary air intake stacks, a vertical exhaust
stack, an engine room, and a spray room.
Immediately adjacent to the test cell is a one-story
structure containing an engine control room, a
pump room, and a fuel-filter room.
The pump room contains a cooling water
pump, an air compressor, a 28-volt dc generator
set, and a 115-volt, 400-Hz ac generator set for
operation of the test facility. In addition, the
pump room contains batteries for operation of
the C02 fire-extinguishing system. The fuel-filter
room contains a fuel flow measuring package in
addition to a fuel-filter separator.
The test facility operates on the principle that
the accumulation of sufficient data for comparison with known or desired optimum values
will satisfactorily indicate the relative service of
an engine under test. To accomplish this result,
you compare the engine performance and test cell
environmental data to standard day performance
criteria for the engine model being tested. Various
controlling, sensing, and indicating systems
are provided to accurately measure engine performance and test cell environmental conditions
during the test.
Engine harness and hookups between engine
and connector panel, as well as special test items,
are the responsibility of the using activity.
A brief discussion of the purpose of each of
the major assemblies is given in the following
paragraphs. System description and operating
principles are also discussed.
Variable Height Stand Assembly
The variable height stand assembly, or thrust
bed, is used to support, restrain, and position the
engine in the desired testing attitude. This
assembly is also used to transfer the engine from
the trailer to the thrust bed. The thrust bed
positioning system and the thrust measuring
system are part of the variable height stand.
The thrust bed positioning system is essentially
a pneumatic-powered, hydraulic oil-driven device
capable of raising and lowering the test engine to
any attitude within operating range. Operating
controls are mounted on the left front A-frame
of the thrust bed.
In addition to supplying lifting force for height
positioning of the thrust bed, oil pressure is used
to drive the calibration cell cylinder whenever
controlled thrust pressure is required.
thrust measuring system becomes operative when
the thrust button is acted against by the load cell
as a result of slight forward motion of the engine
under power. If desired, the load cell may be
preloaded by the preload unit. As thrust is
developed by the engine, the pressure between
the button and the cell produces an electric
potential within the cell. The electric current thus
recording instrumentation for monitoring engine and
test facility conditions is located here. An operator
can exercise complete control of the test facility from
this control board.
The control board instrument system consists of
four panels designated as panel A, panel B, panel C,
and panel D. See figure 10-19. The panels contain
instruments and control components required to
properly operate the test equipment to functionally
test turbojet engines.
The electronic counter is an instrument designed
to count specific events during variable intervals of
time. As installed on the control board, it is used to
indicate engine rotor revolutions per minute.
The interlock system is a simple series circuit
consisting of manual and automatic switches, relays,
and fuses. An electric motor-driven test generator
provides all 28-volt dc power used by the interlock
system and the test equipment. It is started with a
28-volt dc motor-generator set switch. The switch
opens and closes the main relay of the generator
output line. The 28-volt engine master switch is the
master input switch for the 28-volt system, including
the interlock system.
established is connected to the thrust measuring
circuit box, where it is amplified sufficiently to power
the thrust indicator. The thrust indicator is
calibrated to read directly in pounds of engine thrust.
Engine Test Connector Panel Assembly
The engine test connector panel assembly
(connector panel) is provided as a terminal board for
interconnecting the engine with the test equipment.
The engine test connector panel system contains
interconnecting the test engine, control board, and
the remaining test equipment. High- and lowpressure quick-disconnect fittings (self-sealing) for
pressure lines and capped electrical fittings for
electrical lines are provided.
Control Board Instrument Assembly
The control board instrument assembly (control
board) provides for central control of the engine and
the test equipment. The necessary indicating and
Figure 10-19.-Control board.
When this switch is raised to the ON position, its
mating green pilot light is illuminated and engine
master relay is energized. Power is then available
at the 28-volt engine control switch, having
arrived by way of the closed contacts of a
normally open relay.
NOTE: All ON and OFF switches installed
on the control board are positioned so that
their respective ON positions in relation to
the lower edge of the applicable panel are
established with the toggle raised.
Control point No. 1 consists of a CO2 pressure
switch and mating green pilot light. The switch
is a manual reset switch, remotely located at
the connector panel in the test cell, and is
programmed to seek the OFF position whenever
the test cell of the CO2 fire-extinguishing system
is energized manually or automatically. If the
pressure switch contacts are broken, they must be
manually reset to the ON position before the
interlock system circuit can be completed.
Control point No. 2 consists of a front door
key-operated switch and two mating pilot lights.
The key that operates the switch is the key locking
the front door of the test cell; this key cannot be
removed from the door lock mechanism until the
door is locked.
Before the interlock system can be energized
by the switch, the key from the front door must
be removed and brought to the control room for
use with the switch. The two mating pilot lights
indicate the key switch position—green for
CLOSED and red for OPEN.
Control point No. 3 consists of a front door
bypass switch and mating red pilot light. This
switch is provided to permit removal of the front
door key from the switch at control point No. 2.
The key cannot be removed without placing the
switch in the OPEN position, so that the key may
be used to open the front door when it is necessary
for authorized personnel to attend to the engine
or test equipment during periods of engine
both manually and automatically, with the controls and instrumentation installed. For example,
if exhaust stack temperatures in excess of a preset
temperature are experienced during afterburner
operation, the afterburner is automatically cut off
by a latching relay. When this occurs, the afterburner cannot be relit until safe temperature
conditions prevail and the relay is reset. Provisions
have been made for manually overriding the
exhaust stack temperature control system should
it fail or otherwise prove inadequate.
The principle of automatic safety control is
a built-in characteristic of test equipment. An
interlock system has been provided for automatic
engine shutdown or prevention of engine start
when the following conditions prevail:
1. Inadequate water pressure to maintain
exhaust gas cooling, except when bypassed by the
2. Thrust bed not locked and interlocked in
height position.
3. Front door key not in control panel lock,
or door unlocked without utilization of front door
bypass switch and rear door key not inserted in
control panel lock.
4. Primary air supply fire curtain not open.
5. CO 2 fire-extinguishing system energized,
either manually or automatically.
6. Electric power failure.
7. System-wide loss of compressed air supply.
The exhaust augmenter system consists of two
stages. Both stages are adjustable to meet specific
engine exhaust conditions.
Exhaust Augmenter Assembly and Exhaust
Gas Cooling System Assembly
The first-stage exhaust augmenter is located
in the forward test cell wall of the spray room.
The horizontal center line of the first-stage and
the second-stage exhaust augmenter is approximately 5 feet above the test cell floor. The firststage augmenter receives the exhaust discharge of
the test engine and by venturi action tends to draw
such exhaust to the second-stage augmenter in the
spray room. The first-stage augmenter spray ring,
which is controlled manually from the control
board, is included in the first stage.
The exhaust augmenter assembly (augmenter)
and the exhaust gas cooling system assembly
(exhaust cooling system) are provided to ensure
proper disposition of engine exhaust by controlling exhaust gas, flow characteristics, and
temperature. These factors can be monitored,
The size of the first-stage augmenter orifice
and the distance from the plane of the engine
exhaust exit nozzle to the mount of the insert
should be correlated for each engine model to
obtain true engine performance and proper
control of secondary airflow. Improper sizing and
positioning of the first-stage augmenter may result
in the following:
1. Excessive turbulence in the region of the
exit nozzle.
2. Excessive buildup of exhaust backpressure.
3. An increase in exhaust gas temperature.
4. Loss of thrust.
5. Insufficient or excessive secondary airflow.
(Excessive secondary airflow produces ram drag
across the engine; and with operation of highthrust engines, an excessive test cell depression
The second-stage exhaust augmenter is located
in the spray room. A movable collector section
is provided along with a flat orifice plate. Movement of the collector section and the indexing of
the orifice plate afford quantity adjustment of the
airflow from the secondary air intake stack
through the second-stage augmenter. The lateral
diffusing vanes are water cooled and afford
additional exhaust gas cooling. Discharged
exhaust gases, mixed with spray water and air,
pass from the second-stage augmenter into upward diffusing vanes. These vanes are located at
the base of the exhaust stack and upward and out
the exhaust stack.
Fuel System Monitoring Assembly
lubricating oil is supplied from the storage tank,
by way of the oil outlet to the engine lubricating
system and back to the tank, through the oil inlet.
The amount of oil in the tank can be readily
determined by observing the oil level in the
sight gauge. The storage tank is located at an
extension of the connector panel in the test cell.
The auxiliary lubricating oil-cooling system is
provided for use as required by specific engines.
Refer to applicable engine test instructions. It
consists of an oil temperature regulator valve, a
heat exchanger, and suitable plumbing. Water,
by way of the oil temperature regulator valve,
passes through the cooling elements of the
lubricating oil heat exchanger, which is used to
cool the engine oil circulating through the
exchanger. The temperature of the oil returned
to the engine is sensed by the oil temperature
regulator valve by a sensing bulb installed in the
heat exchanger oil line. The oil temperature
regulator reacts to the bulb signal by positioning
a poppet valve. This increases or decreases water
flow through the heat exchanger, thus maintaining
the oil outlet temperature between proper limits.
The oil temperature regulator is provided with a
handwheel so that the temperature range maybe
adjusted to maintain outlet oil temperature within
the allowable range. The water used for cooling
the engine lubricating oil flows from the main
water supply system to the load bank water tank
(containing heating elements), and then to the oil
temperature regulator valve.
The fuel system monitoring assembly (fuel
system) consists of the devices required for fuel
filtration, flow and specific gravity measurement,
and flow control.
The fuel system consists of two 10,000-gallon
underground fuel tanks, two fuel pumps, two
motor-driven fuel line valves, the fuel system
monitoring assembly, and various controls and
pilot lights at the control board. The fuel system
is interlocked to the basic interlock system, to the
CO 2 fire-extinguishing system, and to the exhaust
gas cooling system.
Compressed Air Component Assembly
The compressed air component assembly
(compressed air system) provides pneumatic
power, air system hardware, and pneumatically
operated controls and actuating cylinders
necessary for remote control of numerous devices
throughout the test facility.
The air compressor (a two-stage, air-cooled,
electric motor-driven system) is located in the
pump room, and is controlled by a switch located
on the master control center in the control
room. Two plug-in outlets supply unregulated
compressed air for test cell utility purposes as
required. Air filtration and partial dehydration
are accomplished by two float-type air filters for
two main branch supply lines. The pressure to
each branch line is controlled by two manually
set air pressure regulators. Beyond the pressure
regulators, the air system branches off to the
various control components for other major
Engine Oil Assembly
The engine oil reservoir and engine auxiliary
lubricating oil-cooling system component
assembly (lubricating system) provide a 20-gallon
engine oil reservoir and an auxiliary means of
cooling engine oil to a controlled temperature.
The engine oil reservoir system is made up of
a 20-gallon tank equipped with suitable fittings,
a sight gauge, and breather vent. The engine
systems, such as the fuel system and the exhaust
gas cooling system.
Intercommunication System
The intercommunication system consists of an
eight-station intercom master in the control room,
a suitable amplifying system, and two remote
stations equipped with trumpets and microphones. The six spare station switches at the
master control are not used as installed. The
master unit is equipped with a volume control,
a push-to-talk button, and push-to-talk lock
button. Remote stations are equipped with pushto-talk buttons only. Station No. 1 is located in
the test cell; station No. 2 is located at the test
cell observation port.
Electrical Power Systems
Three primary sources of electric power are
provided for the test facility. They are as follows:
1. Main service to facility: 120/208-volt ac,
three-phase, 60-Hz, four-wire.
2. Motor-driven generator test set: 28-volt dc.
3. Motor-alternator test set: 115-volt ac, onephase, 400-Hz.
Electric power, developed by any test engine
equipped with an engine-driven alternator, is not
used for test purposes. If power is required by the
applicable engine test instructions, a water-cooled
load bank provides for application of a token
5-kva alternator load, consisting of immersion
heater elements.
Engine Starting and Ignition System
The starting of a jet engine or any other type
of gas turbine engine requires that the engine be
rotated at a speed that will provide sufficient air
for the required fuel-air ratio. Provisions are made
for energizing the ignition system to fire the spark
(igniter) plugs at the proper time, and for the
engine to be accelerated until the power developed
by the turbine is adequate for self-sustained
rotation. Initial rotation of the engine during
starting may be accomplished by use of either an
Figure 10-20.-Portable universal engine runup test system.
electrically operated starter motor or a compressed, air-operated, air turbine starter motor.
The electric starter motor requires a source of dc
voltage. The air turbine motor requires a source
of compressed air.
CO2 System
The CO 2 fire system for the test facility
consists of a 2-ton CO2 storage tank, which
incorporates a refrigeration system to maintain
the liquid carbon dioxide at the proper storage
pressure; a handwheel-operated shutoff valve on
the dip tube of the storage tank; pressure-operated
control valves for quickly releasing the carbon
dioxide; a piping system terminating in the CO2
discharge nozzles, strategically located in the
protected areas; and various controls, relays,
thermostats, alarm gongs, spurt and flood push
buttons, and pressure-operated switches. For
complete fire control coverage, the CO2 system
is electrically linked to the interlock system and
pressure linked to the main fuel line valve of the
engine fuel supply.
T h e C O2 storage tank is designed for
maximum working pressure and is equipped with
a complete Freon (F-12) refrigeration system,
which is automatically controlled by an internally
mounted pressure switch. The refrigeration system
maintains a nominal storage tank pressure. The
storage tank is protected by a safety assembly
consisting of a dual switching valve, which, when
in the normal position, places in service one
auxiliary automatic refrigeration valve, one safety
valve, and one high-pressure relief disc. The safety
devices provide complete protection against
abnormally high tank pressures. An abnormally
high pressure usually results from power or
compressor failure continuing over a period of
several hours.
There are several different models of universal
test cells in use. Figures 10-20, 10-21, 10-22, and
10-23 show some of the more commonly used test
Figure 10-21.-Turboprop engine test system.
Figure 10-22 .-Portable engine runup test system.
Figure 10-23.-Engine test and runup system.
cells that provide the aircraft maintenance
activities with a portable and universal system for
the operational and functional testing of jet aircraft engines.
These systems perform the basic functions
of checking all the engine performance characteristics against the engine manufacturer’s
operational parameters, as approved by NAVAIRSYSCOM. The test cells display engine
temperatures, vibrations, fuel metering, fuel flow
pressures, thrust, lube oil temperatures and
pressure, compressor pressure, hydraulic oil
pressure, anti-ice pressure, turbine rpm, and
position indications such as nozzle and stator vane
and throttle.
(not normally found in a test cell) needed to
satisfactorily conduct functional and performance
The console provides junction facilities to
connect the cell power to the engine, a system for
remote control measurement of throttle position,
a transmitter and receiver to indicate inlet guide
vane position, and a dc electronic indicating
system for measuring nozzle position. A thermocouple type of anti-icing temperature indicator,
a starter circuit, and switches and cables necessary
for operation of the engine and console under test
conditions are included in some consoles.
Just as starting procedures vary with the
various types of engines, the controls and
instrumentation vary with different test cells.
Checking the engine for proper operation consists
primarily of reading engine instruments and then
comparing the deserved values with those given
by the manufacturer for specific engine conditions, atmospheric pressures, and temperatures.
NOTE: These test systems may be used at
any site location that has been provided
with adequate tie-downs (either concrete
embedded or buried expansion anchors).
Some engines require special testing consoles.
The console provides the electrical circuits
Figure 10-24.—Engine runup test log sheet.
The engine test log records the data obtained
during the engine test run. See figure 10-24. The
log provides a record of the engine test for future
reference and acts as documentary proof that the
engine was subjected to the prescribed test
procedures. The data must be complete, accurate,
neat, and legible. Upon receipt of the engine for
testing, the operator will enter the name of the
testing activity, the engine model, serial number
of the engine, and the date of the engine test.
During the test, record all unusual occurrences
in the remarks section of the test log. Record all
starts, shutdowns, times for accelerations, and
times for adjustments and settings. During starts,
record the time of day the engine was started,
maximum turbine inlet temperature encountered,
and the duration of that temperature. Record the
time for acceleration and the stabilization time.
At the end of the test, record the engine coastdown time. Coast-down time is defined as the time
elapsed from the moment fuel is cut off to the
time the engine comes to a complete stop. Coastdown time has no absolute value. A record maintained for engines will show what the expected
average coast-down time should be. Any engine
with an abnormally short coast-down time should
be viewed with suspicion and investigated for
compressor rub or other malfunctions. The
operator is required to sign all the test run logs,
and is held responsible for the accuracy and
completeness of all the test data.
Test schedules will vary with each different
model of engine and manufacturer. Always
refer to the appropriate engine manual when
performing engine test runs.
ANODIZE—To coat a metal by chemical or
electrolytic processes.
ABORT—To cut short or break off an action,
operation, or procedure with an aircraft, guided
missile, or the like, especially because of equipment failure; for example, to abort a mission.
ANTI-ICING—To prevent ice formation on
an aircraft’s surface or engines.
ACCELERATION—A change in the velocity
of a body, or the rate of such change with respect
to speed or direction.
APU—Auxiliary power unit.
ARTICULATED—A movable joint between
segmented parts. An articulated rotor system is
one whose individual or collective main rotor
blades are free to flap, feather, and drag
individually or collectively.
ACCESSORY—A part, subassembly, or
assembly designed for use in conjunction with or
to supplement another assembly or unit. For
example, the fuel control is an accessory for a
turbojet engine.
ADDITIVE—A chemical that is added in
minor proportion to a parent substance (fuel or
oil) to create, enhance, or suppress a certain
property or properties in the parent material.
ATE—Automatic test equipment.
ATMOSPHERE—The body of air surrounding the earth. The atmospheric pressure at sea level
is 14.7 psi.
AERODYNAMICS—The science that deals
with the motion of air and other gaseous fluids
and the forces acting on bodies in motion relative
to such fluids.
ATOMIZE—To reduce a fluid (fuel) to a fine
spray or mist.
AUTOROTATION—The process of producing lift with airfoils that rotate freely and are not
engine driven. When a helicopter autorotates, the
air flows upward through the rotor system, rather
than downward, as is the case when the rotors are
engine driven.
AFTERFIRE—The term used when a tail-pipe
fire develops after engine shutdown.
ALLOY—A substance composed of two or
more metals.
AMBIENT—Surrounding; adjacent to; next
to. For example, ambient conditions are physical
conditions of the immediate area such as ambient
temperature, ambient humidity, ambient pressure,
AVGAS—Aviation gasoline for reciprocating
ANGLE OF ATTACK—The angle at which
a body, such as an airfoil or fuselage, meets a flow
of air. The acute angle between the chord of an
airfoil and the direction of the relative wind.
AXIS—An imaginary line that passes through
a body about which the body rotates or may be
assumed to rotate. For example, the horizontal
axis, the lateral axis, and the longitudinal axis
about which an aircraft rotates.
ANODIC—The term used for the terminal
point where electricity passes from the electrolyte
to an external point; for example, the negative
terminal on a cell or battery.
BAROMETRIC—The term used for instruments and devices that measure atmospheric
pressure and respond in some way to the changes
in pressure.
flowing through a tube reaches a constriction, or
narrowing of the tube, the velocity of fluid flowing through the constriction increases and the
pressure decreases.
BITE—Built-in test equipment.
BTU—British thermal unit. A unit of heat
commonly used in heat engineering. It is the
amount of heat necessary to raise the temperature
of 1 pound of water 1 degree Fahrenheit.
CAUTION—An operating procedure, practice, or condition, etc., that may result in damage
or destruction to equipment if not carefully
observed or followed.
CELSIUS—The temperature scale using the
freezing point of water as zero and the boiling
point as 100, with 100 equal divisions between
called degrees. A reading is usually written in the
abbreviated form; for example, 75°C. This scale
was formerly known as the centigrade scale, but
was renamed Celsius in recognition of Anders
Celsius, the Swedish astronomer who devised the
CONVERGING—To tend to meet at a point
or line; incline toward each other.
COORDINATOR—An electromechanical
unit that mounts to the back of the fuel control
in turboprop engines. It coordinates the propeller,
electrical temperature datum control, and the fuel
control through signals received from the power
COWLING—A removable cover or housing
placed over or around an aircraft component or
section, especially an engine.
CSE—Common support equipment.
CURE—The changing of matrix properties
(hardening) by chemical reaction, usually accomplished with heat and vacuum pressure.
CYCLIC PITCH—The mechanical means to
change the pitch of the main rotor blades during
the rotor system cycle of rotation to control the
tilt of the rotor disk.
DE-ICING—The breaking off or melting of
ice from aircraft surfaces or fuel induction
CENTRIFUGAL—Moving or directed outward away from the center or axis.
splitting or separating of a laminated composite
material between laminate plies.
means of simultaneously increasing or decreasing the pitch of all main rotor blades.
DENSITY—The weight per unit volume of a
(CIT or T2)—The temperature of the air entering
the gas turbine compressor as measured at the
front frame. One of the parameters used to
calculate engine power output (torque) and
scheduling combustion fuel flow and variable
stator vane angle.
DIFFUSER—A specially designed duct,
chamber, or section equipped with guide vanes,
which decrease velocity and increase pressure.
DILUTENT—A substance or medium to
reduce strength or make thinner.
CONCENTRIC—Having a common center.
DISILICATES—Compounds containing two
atoms of silicon in conjunction with other
CONDENSATION—The act of reducing a
gas or vapor to a liquid or solid form.
CONING—The upward flexing of the rotor
blades resulting from the vectorial combined
effects of centrifugal force and lift.
DISSIPATE—To scatter or spread.
DIVERGENT—To extend in different directions from a common point.
CONTAMINATION—Any foreign material
or substance whose presence (in the fluid) is
capable of adversely affecting the performance
or reliability of a system.
DYNAMICALLY—As in balanced by
rotating at high speed.
EAMP—Engine Analytical Maintenance
HELICAL—Pertaining to or having the form
of a spiral.
ELECTROCHEMICAL—Chemical changes
produced by electricity and the production of
electricity through chemical reactions.
HERO—Hazard of electromagnetic radiation
to ordnance.
HORSEPOWER—A unit of power equal to
the power necessary to raise 33,000 pounds 1 foot
in 1 minute.
ELECTROLYTE—A conducting medium in
which the flow of current is accompanied by the
movement of ions.
HOT START—The term used when exhaust
gas temperatures rise above specified limits during
the engine starting cycle.
ELECTROPLATE—To coat with a metal
through the use of electricity.
ENERGY—The ability or capacity to do
HYDRAULICS—The branch of mechanics
that deals with the action or use of liquids forced
through tubes and orifices under pressure to
operate various mechanisms.
ENGINE TRIMMING—The term used for
fine tuning a newly installed engine to the airframe, or whenever engine fuel system major
components have been removed and/or replaced.
HYDROUS—Containing water or its elements
in some form.
FAIRING—A part or structure that has a
smooth, streamlined outline, used to cover a
nonstreamlined object.
IMPEL—To propel or impart motion.
IMPINGEMENT—The act of striking
against. An impingement starting system is where
starter air is directed to strike against the turbine
blades, causing the engine to rotate.
FAHRENHEIT—The temperature scale that
registers the freezing point of water at 32 degrees
and the boiling point at 212 degrees.
INERTIA—The tendency of a body at rest to
remain at rest, and a body in motion to continue
to move at a constant speed along a straight line,
unless the body is acted upon in either case by an
unbalanced force.
FLAMMABLE—Describes any combustible
material that can be easily ignited and that will
burn rapidly.
with flammable being the preferred term.
FLASH POINT—The lowest temperature at
which a sufficient amount of vapor is given off
by a liquid to form an ignitable mixture with air.
INHIBITOR—A substance that decreases the
rate of or stops completely a chemical reaction.
FOD—Foreign object damage. Damage
resulting from foreign objects (nuts, bolts, rocks,
etc.) entering the engine inlet.
JETTISON—To throw or dump overboard.
For example, to drop or eject fuel, tanks, or gear
from an aircraft to lighten the load for emergency
FORCE—The action of one body on another
tending to change the state of motion of the body
acted upon. Force is usually expressed in pounds.
JOAP—Joint Oil Analysis Program.
FRICTION—Surface resistance to relative
KINETIC—Energy in motion.
LABYRINTH SEALS—Oil seals with a
nonrubbing surface. Designed to allow minimal
amounts of leakage, which is controlled by air
pressure on one side.
GRAPHITE—A soft, native black to darkgray carbon used for pencil leads and lubricants.
LAMINATE—A combination of two or more
single plies of laminae bonded together to form
a structure.
NACELLE—A streamlined structure,
housing, or compartment on an aircraft; for
example, a housing for the engine.
LAMPS—Light Airborne Multipurpose
System (Helicopter).
NAVOSH—Navy Occupational Safety and
NOAP—Navy Oil Analysis Program.
LATERAL—The width dimension; for
example, the lateral axis of an aircraft runs
lengthwise from wing tip to wing tip.
NOTE—An operating procedure, practice, or
condition, etc., that is essential to emphasize.
LONGITUDINAL—The lengthwise dimension; for example, the longitudinal axis of an aircraft runs lengthwise from the nose to the tail.
minimum requirements for rating and rate, as set
forth by the Chief of Naval Personnel.
MACH—A number indicating the ratio of the
speed of an object to the speed of sound in the
medium through which the object is moving.
OXIDATION—In general, the process where
oxygen is added to a compound. The oxidation
process in petroleum may lead to gum or resin
MASS—The amount of fundamental matter
of which an object is composed.
OXIDE—A compound containing oxygen and
one or more elements.
MATTER-The substance or substances of
which all physical objects consist or are composed.
PLENUM CHAMBER—A chamber in which
the gas or air pressure is greater than atmospheric
MAY AND NEED NOT—Terms indicating
that the application of a procedure is optional.
used for oil, fuel, and hydraulic pumps that
deliver more fluid to the system than is actually
MICA—Any member of a group of minerals,
hydrous disilicates of aluminum with other bases,
that separate readily into thin, tough, often
transparent laminates.
PPB—Power plants bulletin.
MICRON—A millionth of a meter or about
0.000039 inch.
PPC—Power plants change.
PRESSURE DIFFERENTIAL—The difference or ratio between input and output
MIL-C—Military compounds.
MIL-H—Military hydraulic fluids.
PRESSURE—The amount of force
distributed over each unit of area. Pressure is
expressed in pounds per square inch (psi).
MIL-L—Military lubricants.
PROPAGATION—The act of transmitting
and controlling heat in a flame.
MILSTD—Military standard.
MOLECULES—The smallest physical units
o f an element or compound.
PROPORTIONAL—Having the same or
constant ratio or relation.
PSE—Peculiar support equipment.
N—Abbreviation for speed or rpm. When
used with other letters, it indicates the speed of
that section. Examples include Ng for compressor
speed and Nf for power turbine or fan speed.
PSI—Position sensitive indicator. Pounds per
square inch.
SLIPSTREAM—The stream of air driven
backward by a rotating propeller.
PYLON—A structure or strut that supports
an engine pod, external tank, etc., on an aircraft.
SLUGGING—In airplanes, loss of liquid fuel
from tank vents because of the pulling action of
escaping vapors.
QEC—Quick engine change.
QECA–Quick engine change assembly.
SOAPSTONE—A hand talc with a soapy or
greasy feel.
QECK–Quick engine change kit.
RAM AIR—Air forced into an air intake or
duct by the motion of the intake or duct through
the air.
SPECIFIC GRAVITY—The ratio of the
weight of a given volume of a substance to the
weight of an equal volume of some standard
substance, such as water.
RAM PRESSURE—Air pressure in the inlet
caused by a jet aircraft’s forward motion.
RATIO—The relation between two similar
items in respect to the number of times the first
contains the second.
SPECTROMETER—An optical device used
to determine the amounts and types of contaminants in oil sample droplets under the Navy
Oil Analysis Program.
RESERVOIR—A receptacle or chamber for
holding a liquid or fluid.
SPEED—The distance a body in motion
travels per unit of time.
STATOR—A system of stationary airfoils.
ROTOR—A system of rotating airfoils.
SUMP—A chamber into which a fluid drains
for pickup and recirculation.
RPM—Revolutions per minute.
SCAVENGE OIL—The term used for oil that
has been through the engine lubrication system
and is on its way back to the tank.
T—Abbreviation for temperature. Combined
with a number it will indicate temperature at a
specific engine station. For example, T 2 indicates
compressor inlet temperature.
SCUPPER—Any drain for fluid runoff.
TEFLON®—Trademark used in making a
tough, nonsticking coating for gaskets, backup
rings, bearings, electrical insulators, etc.
SE—Support equipment. All of the equipment
on the ground needed to support aircraft in a state
of readiness for flight.
SEGTE—Support equipment gas turbine
TENSION—A force or pressure that exerts a
pull or resistance.
SELECTOR VALVE—A valve used to
control the flow of fluid to a particular
mechanism, as in a hydraulic system.
THERMOCOUPLE—A device for measuring
temperature, consisting of two dissimilar metallic
conductors joined at their ends.
SERVICING—The refilling of an aircraft
with consumables such as fuel, oil, and compressed gases to predetermined levels, pressures,
quantities, or weights.
THERMODYNAMICS—The science concerned with the relations between heat and
mechanical energy or work, and the conversion
of one to the other.
SHALL—Term indicating that application of
a procedure is mandatory.
THRUST—The forward-direction pushing or
pulling force developed by an aircraft engine or
rocket engine.
SHOULD—Term indicating that application
of a procedure is recommended.
TORQUE—A turning or twisting force.
TOXIC—Harmful, destructive, deadly;
VISCOSITY—The internal resistance of a
liquid that tends to prevent it from flowing.
TRACKING—To ensure that helicopter
blades rotate in the same horizontal plane.
VOLATILITY—The ability of a fluid to
TURBULENCE—The flow of a fluid (liquid
or gaseous) within an object such that the velocity
at any fixed point in the fluid varies irregularly.
WARNING—An operating procedure, practice, or condition, etc., that may result in injury
or death if not carefully observed or followed.
allows the angle of a wing, intake ramp, or
exhaust duct to adjust to the best condition for
the most efficient use.
WILL—A term used to indicate futurity,
never to indicate any degree of requirement for
application of a procedure.
VELOCITY—The rate of motion in a
particular direction.
WINDMILLING—Rotating a gas turbine
engine without energizing the ignition system.
Aero bomb hoists, 3-17
Aircraft intermediate maintenance department,
10-2 to 10-19
accessories drive section and mating
gear inspections, 10-19
cleaning, 10-2 to 10-4
combustion section repairs, 10-10
to 10-13
compressor section repairs, 10-5 to 10-10
exhaust section inspection, 10-18 to 10-19
general engine repair and
inspections, 10-5
main engine bearings inspection, 10-19
markings, 10-4
turbine section repairs, 10-13 to 10-18
Airframe fuel system, 4-10 to 4-20
external fuel tank system
description, 4-12 to 4-20
fuel tank construction, 4-10 to 4-12
Athrodyd (ramjet), 1-2 to 1-3
Auxiliary power units, 6-20 to 6-21
Aviation support equipment, 3-1 to 3-19
identification of support equipment, 3-1
to 3-3
nonpowered support equipment, 3-13 to
engine trailers and workstands,
3-15 to 3-17
maintenance platforms, 3-13 to
special-purpose support equipment,
3-17 to 3-19
powered support equipment, 3-3 to 3-13
gas turbine compressors (GTCs),
3-10 to 3-12
hydraulic power supplies/hydraulic
test stand, 3-12 to 3-13
mobile air-conditioners, 3-6 to 3-10
mobile electric power plants
(MEPPs), 3-3 to 3-6
mobile motor-generator sets
(MMGs), 3-6
SE training and licensing program, 3-19
types of support equipment, 3-3
Axial-flow compressors, 1-14 to 1-18
Bell crank support mounting pad, 7-18 to 7-19
Bladder-type fuel cells, 4-12
Bleed-air systems (airframe), 6-16 to 6-20
Bleed-air systems (engine), 6-13 to 6-16
Borescope inspection, 9-17 to 9-22
Centrifugal-flow compressors, 1-13 to 1-14
Clamps, 2-16 to 2-17
Combination wrenches, 2-5 to 2-6
Drive shaft, main, 7-17 to 7-18
Drop tanks, 4-12 to 4-14
Electrical test equipment, 9-13
Engine and airframe related systems, 6-1 to 6-21
basic electricity, 6-6
basic hydraulics, 6-1 to 6-2
hydraulic fluids, 6-2 to 6-5
contamination, 6-2 to 6-3
contamination control, 6-3
origin contaminants, 6-2 to
Engine and airframe related systems—Continued
hydraulic fluids—Continued
fuel as a hydraulic fluid, 6-3 to 6-5
combination inlet guide vane
and bleed valve system,
variable-area exhaust
nozzles, 6-5
variable inlet guide vanes
and stators, 6-5
hydraulic system maintenance, 6-3
ignition, starting, bleed, air and auxiliary
power unit systems, 6-6 to 6-21
aircraft ignition systems, 6-6 to
high-energy, capacitordischarge ac ignition
system, 6-9 to 6-10
high-energy, capacitordischarge dc ignition
system, 6-7 to 6-9
ignition operation, 6-10
ignition system maintenance,
6-10 to 6-11
auxiliary power units, 6-20 to 6-21
bleed-air systems (airframe), 6-16
to 6-20
airframe deicing and antiicing systems, 6-18 to
airframes fuel systems, 6-20
bleed-air systems (engine), 6-13 to
anti-icing system, 6-15 to
across-bleed air engine
starting system, 6-15
oil and seal pressurization
system, 6-15
starting systems, 6-11 to 6-13
air turbine starter, 6-11
electrical starters, 6-12
hydraulic starters, 6-13
turbine impingement starter,
6-11 to 6-12
Engine control quadrant, 7-13
Engine designation systems, 1-8 to 1-9
Engine-driven pumps, 4-28 to 4-31
Engine fuel system, 4-20 to 4-31
engine-driven pumps, 4-28 to 4-31
filters, 4-27 to 4-28
fuel control (JFC 25-3) operation, 4-20 to
fuel selectors, 4-26 to 4-27
fuel valves, 4-23 to 4-26
Engine oil system description, 5-16 to 5-17
External fuel tank system description, 4-12 to
Filters, 4-27 to 4-28
Flight control system, 7-6
Fluid line identification, 9-11 to 9-13
Fuel control (JFC 25-3) operation, 4-20 to 4-23
Fuel system maintenance, 4-31 to 4-41
defueling, depuddling, purging, 4-33 to
fuel cell removal and installation, 4-35 to
fuel leak analysis, 4-32
fuel system part maintenance, 4-37 to 4-40
location of leaks, 4-32 to 4-33
rigging and adjusting, 4-40 to 4-41
Fuel tank construction, 4-10 to 4-12
Fuel valves, 4-23 to 4-26
Fuels, 4-1 to 4-9
characteristics, 4-2 to 4-4
fuel contamination, 4-4 to 4-9
types of fuel, 4-4
Gas-turbine engine, 1-4
Gas turbine test equipment, 9-13 to 9-22
Gaskets, seals, and packings, replacement of,
5-17 to 5-19
Gear-type pumps, 4-28 to 4-29
Glossary, AI-1 to AI-6
GTCs, gas turbine compressors, 3-10 to 3-12
GTC-85, 3-11
NCPP-105, 3-11 to 3-12
Hand tools, common, 2-3 to 2-12
combination wrenches, 2-5 to 2-6
hammers, 2-3
pliers, 2-7 to 2-9
screwdrivers, 2-6 to 2-7
socket sets, 2-3 to 2-5
special tools, 2-9 to 2-12
micrometers, 2-11 to 2-12
torque wrenches, 2-9 to 2-11
Hardware, 2-12 to 2-19
clamps, 2-16 to 2-17
safety methods, 2-17 to 2-19
cotter pins, 21-9
safety wiring, 2-17 to 2-19
threaded fasteners, 2-12 to 2-15
identification of threaded fasteners,
installation of threaded fasteners,
nuts, 2-14
washers, 2-15 to 2-16
Helicopters and turboshaft power plants, 7-1 to
flight control system, 7-4 to 7-6
helicopter flight characteristics, 7-1 to 7-4
factors affecting helicopter flight,
7-2 to 7-4
autorotation, 7-4
blade flapping, 7-3
coning, 7-3
dissymmetry of lift, 7-2 to 7-3
ground effect, 7-4
gyroscopic precession, 7-3 to 7-4
power settling, 7-4
torque, 7-2
factors affecting rotor blade lift, 7-4
density altitude, 7-4
pitch of rotor blades, 7-4
rotor area, 7-4
smoothness of rotor blades,
helicopter theory of lift, 7-1
to 7-2
helicopter power plants, 7-8 to 7-12
T58-GE-10 turboshaft engine, 7-8
to 7-11
gas generator section, 7-8 to 7-11
power turbine section, 7-11
T700-GE-401 turboshaft engine,
7-11 to 7-12
combustor section, 7-12
compressor section, 7-11 to
exhaust section, 7-12
inlet section, 7-11
turbine section, 7-12
Helicopters and turboshaft power plants—
power transmission component
maintenance, 7-18 to 7-21
main transmission inspection, 7-18
to 7-19
barrel nut inspection, 7-18
bell crank support mounting
pad, 7-18 to 7-19
main module mounting feet,
main transmission
installation, 7-19
main transmission removal,
tail drive shaft inspection, 7-19
tail drive shaft repair, 7-19 to 7-21
primary helicopter components, 7-12 to
engine control quadrant, 7-13
intermediate gearbox, 7-16 to 7-17
main drive shaft, 7-17 to 7-18
main transmission (SH-60), 7-16
accessory module, 7-16
input module, 7-16
main module, 7-16
main transmission gearbox (SH-3),
7-14 to 7-15
oil cooler and blower, 7-18
power transmission system, 7-13 to
rotor brake, 7-18
tail drive shaft, 7-18
tail gearbox, 7-17
rotor systems, 7-21 to 7-22
main rotor blades, 7-21 to 7-22
main rotor head, 7-21
tail (rotary) rudder head, 7-22
rotor system maintenance, 7-23 to 7-25
electronic tracking, 7-25
rotor blade tracking, 7-23
strobex tracking, 7-24
types of helicopters, 7-6 to 7-8
High-voltage insulation tester, 9-16 to 9-17
Hydraulic fluids, 6-2 to 6-5
contamination, 6-2 to 6-3
fuel as a hydraulic fluid, 6-3 to 6-5
hydraulic system maintenance, 6-3
Hydraulic power supplies/hydraulic test stand,
3-12 to 3-13
Identification of support equipment, 3-1 to 3-3
Ignition systems, aircraft, 6-6 to 6-11
Intermediate gearbox, 7-16 to 7-17
Jet aircraft engine lubrication systems, 5-1 to
lubricants, 5-1 to 5-4
types of lubricants, 5-1 to 5-2
contamination of lubricating
oils, 5-2 to 5-3
designations of lubricating
oils, 5-2
functions of jet engine oils,
lubricating greases and their
properties, 5-3 to 5-4
lubrication systems, 5-4 to 5-17
engine oil system description, 5-16
to 5-17
engine chip detector, 5-17
oil coolers, 5-17
oil filter, 5-17
oil pressure system, 5-17
oil tank, 5-17
Jet aircraft engine lubrication systems.—
lubrication systems —Continued
oil system components, 5-8 to 5-16
chip detectors, 5-13 to 5-14
filters, 5-12 to 5-13
gauge connections, 5-15
oil coolers, 5-14
oil jets, 5-15
oil pumps, 5-10 to 5-11
oil system seals, 5-15 to 5-16
oil tanks, 5-8 to 5-10
valves, 5-11 to 5-12
vents, 5-15
types of lubrication systems, 5-4 to
dry-sump system, 5-7
wet-sump system, 5-4 to 5-7
maintenance, 5-17 to 5-27
adjustment of oil pressures, 5-19
location of leaks and defects, 5-17
metal particle identification, 5-21
Navy Oil Analysis Program
(NOAP), 5-21 to 5-27
identification of wear metals,
NOAP forms and logbook
entries, 5-24 to 5-27
oil sampling techniques, 5-22
to 5-24
wear metals, 5-22
removal and replacement of
magnetic drain plugs, 5-19 to
removal and replacement of oil
filters, 5-19
replacement of gaskets, seals, and
packings, 5-17 to 5-19
Jet aircraft fuel and fuel systems, 4-1 to 4-41
airframe fuel system, 4-10 to 4-20
external fuel tank system
description, 4-12 to 4-20
drop tanks, 4-12 to 4-14
external fuel tank jettison,
external fuel transfer, 4-15
fuel tank components, 4-15
to 4-20
fuel tank construction, 4-10 to 4-12
bladder-type fuel cells, 4-12
self-sealing fuel cells, 4-10
to 4-12
engine fuel system, 4-20 to 4-31
engine-driven pumps, 4-28 to 4-31
gear-type pumps, 4-28 to
variable-displacement pump,
4-29 to 4-31
filters, 4-27 to 4-28
microfilter, 4-27
plain screen mesh filter,
wafer screen filter, 4-28
fuel control (JFC 25-3) operation,
4-20 to 4-23
fuel selectors, 4-26 to 4-27
fuel valves, 4-23 to 4-26
drain valves, 4-24 to 4-25
flow divider, 4-24
fuel spray nozzles and fuel
manifolds, 4-25 to 4-26
fuel-pressurizing valve, 4-23
to 4-24
fuel system maintenance, 4-31 to 4-41
defueling, depuddling, purging, 4-33
to 4-34
defueling, 4-33 to 4-34
depuddling, 4-34
purging, 4-34
Jet aircraft fuel and fuel systems—Continued
fuel system maintenance—Continued
fuel cell removal and installation,
4-35 to 4-37
handling procedures, 4-35 to
installation, 4-36 to 4-37
removal, 4-35
testing, 4-37
torquing requirements, 4-37
fuel leak analysis, 4-32
fuel system part maintenance, 4-37
to 4-40
combustion chamber drain
valve, 4-39
fuel lines and fittings, 4-38
fuel nozzles, 4-38
fuel part removal/installation,
4-39 to 4-40
fuel system component
inspection, 4-37
high-pressure filter, 4-38
high-pressure fuel lines, 4-38
pumps, 4-37
ram air turbine, hose, and reel
inspection, 4-38 to 4-39
selector valves, 4-38
location of leaks, 4-32 to 4-33
fuel dye to locate leaks, 4-32
to 4-33
preparation for fuel cell
maintenance, 4-33
rigging and adjusting, 4-40 to 4-41
fuels, 4-1 to 4-9
characteristics, 4-2 to 4-4
additives, impurities, and
their effects, 4-3
combustion products, 4-3
flash point and fire point, 4-3
freeze point, 4-4
handling characteristics, 4-3
heat energy content, 4-3
volatility, 4-2 to 4-3
viscosity, 4-3
fuel contamination, 4-4 to 4-9
measuring contamination, 4-5
sampling procedures, 4-7 to 4-9
types and limits of
contamination, 4-5 to 4-7
types of fuel, 4-4
Jet engine test cells, 10-19 to 10-29
enclosed test facility, 10-20 to 10-25
engine test log sheets, 10-29
portable universal engine runup test
systems, 10-25 to 10-28
Jet engine theory and design, 1-1 to 1-33
basic theory of jet propulsion, 1-1 to 1-4
athodyd (ramjet), 1-2 to 1-3
gas turbine engine, 1-4
pulsejet engine, 1-3 to 1-4
rocket, 1-2
jet engine types and designations systems,
1-7 to 1-12
engine designation systems, 1-8 to
ANA bulletin number 306,
type symbols, 1-9
manufacturer symbols, 1-9
model numbers, 1-9 to 1-12
manufacturer’s symbols, 1-10
MIL-STD-879, 1-10
model indicator, 1-11 to 1-12
type indicator, 1-10
special designations, 1-9
turbojet, 1-7 to 1-8
turbofan, 1-8
turboprops, 1-8
turboshaft, 1-7
Jet engine theory and design—Continued
jet turbine engine major assemblies, 1-12
to 1-33
accessory section, 1-31 to 1-32
afterburner section, 1-32 to 1-33
air entrance section, 1-12 to 1-13
single entrance/divided
entrance, 1-12
subsonic/supersonic ducts,
1-12 to 1-13
combustion section, 1-18 to 1-23
annular or basket type, 1-20
to 1-21
can type, 1-19 to 1-20
can-annular type, 1-21 to
compressor section, 1-13 to 1-18
axial-flow compressors, 1-14
to 1-18
centrifugal-flow compressors,
1-13 to 1-14
exhaust section, 1-27 to 1-31
turbine section, 1-23 to 1-27
physical principles of jet propulsion, 1-4
to 1-7
definition of terms, 1-4 to 1-6
Newton’s laws of motion, 1-6 to
Jetcal analyzer and jet calibration test units, 9-11
Lubricants, 5-1 to 5-4
Lubrication systems, 5-4 to 5-17
engine oil system description, 5-16 to 5-17
oil system components, 5-8 to 5-16
types of lubrication systems, 5-4 to 5-7
Magnetic drain plugs, removal and replacement
of, 5-19 to 5-21
Main transmission gearbox (SH-3), 7-14 to 7-15
Main transmission (SH-60), 7-16
MEPPs, mobile electric power plants, 3-3 to 3-6
NC-2A, 3-4
NC-8A, 3-4
NC-10C, 3-5 to 3-6,
Microfilter, 4-27
Micrometers, 2-11 to 2-12
MMGs, mobile motor-generator sets, 3-6
Mobile air-conditioners, 3-6 to 3-10
A/M32C-17, 3-7 to 3-9
NR-5C, 3-8 to 3-10
NR-10, 3-10
Multimeter, 9-14
Navy Oil Analysis Program (NOAP), 5-21 to 5-27
Newton’s laws of motion, 1-6 to 1-7
Nonpowered support equipment, 3-13 to 3-19
engine trailers and workstands, 3-15 to
3000B trailer, 3-15
3110 workstand, 3-17
4000A and 4000B trailers, 3-16 to
maintenance platforms, 3-13 to 3-15
B-2 workstand, 3-13
B-4A and B-5A platforms, 3-14
other maintenance platforms, 3-15
special-purpose support equipment, 3-17
to 3-19
aero bomb hoists, 3-17
jet engine corrosion control cart,
3-18 to 3-19
Ohmmeter, 9-14 to 9-16
Oil cooler and blower, 7-18
Oil filters, removal and replacement of, 5-19
Oil pressures, adjustment of, 5-19
Oil system components, 5-8 to 5-16
Physical principles of jet propulsion, 1-4 to 1-7
Plain screen mesh filter, 4-28
Pliers, 2-7 to 2-9
Power plant inspection, repair, and testing, 10-1
to 10-29
aircraft intermediate maintenance
department, 10-2 to 10-19
accessories drive section and
mating gear inspections, 10-19
cleaning, 10-2 to 10-4
abrasive blasting, 10-4
decarbonizing, 10-3 to 10-4
decreasing (solvent
cleaning), 10-3
steam cleaning, 10-3
vapor decreasing, 10-3
combustion section repairs, 10-10
to 10-13
combustion chamber inner
and outer ducts, 10-12 to
combustion chamber liners,
combustion chamber
support, 10-10 to 10-12
compressor section repairs, 10-5 to
compressor blade damage
and repair, 10-6 to 10-9
compressor contamination,
compressor failures, 10-6
compressor leaks, 10-5
compressor stator vanes,
10-9 to 10-10
exhaust section inspection, 10-18 to
general engine repair and
inspections, 10-5
main engine bearings inspection,
markings, 10-4
permanent markings, 10-4
temporary markings, 10-4
turbine section repairs, 10-13 to
turbine blade replacement,
turbine blade sulfidation,
turbine blades, 10-15 to
turbine stator vanes, 10-16 to
jet engine test cells, 10-19 to 10-29
enclosed test facility, 10-20 to
CO2 system, 10-25
compressed air component
assembly, 10-23 to 10-24
control board instrument
assembly, 10-21 to 10-22
electrical power systems,
engine oil assembly, 10-23
engine starting and ignition
system, 10-24 to 10-25
engine test connector panel
assembly, 10-21
exhaust augmenter assembly
and exhaust gas cooling
system assembly, 10-22
to 10-23
Power plant inspection, repair, and testing—
jet engine test cells—Continued
enclosed test facility—Continued
fuel system monitoring
assembly, 10-23
intercommunication system,
variable height stand
assembly, 10-20 to 10-21
engine test log sheets, 10-29
portable universal engine runup test
systems, 10-25 to 10-28
three-degree gas turbine engine repair
program, 10-1 to 10-2
first-degree repair, 10-1 to 10-2
second-degree repair, 10-2
third-degree repair, 10-2
Power plant troubleshooting, 9-1 to 9-22
fluid line identification, 9-11 to 9-13
gas turbine test equipment, 9-13 to 9-22
borescope inspection, 9-17 to 9-22
borescope use, 9-18 to 9-22
types of borescopes, 9-17 to
electrical test equipment, 9-13
general rules for electrical test
meters, 9-13 to 9-17
high voltage insulation
tester, 9-16 to 9-17
multimeter, 9-14
ohmmeter, 9-14 to 9-16
jetcal analyzer and jet calibration
test units, 9-17
water washing, 9-22
general engine troubleshooting, 9-1 to 9-11
common engine problems, 9-2 to
air system, 9-3
combustion chamber and
turbine section, 9-3
compressor section, 9-3
fuel system, 9-3
lubrication system, 9-3
general safety procedures, 9-2
troubleshooting errors, 9-4 to 9-5
troubleshooting procedures, 9-3 to
use of diagrams, drawings, and
charts in troubleshooting, 9-5 to
Power transmission component maintenance, 7-18
to 7-21
Power transmission system, 7-13 to 7-14
Powered support equipment, 3-3 to 3-13
gas turbine compressors (GTCs), 3-10 to
GTC-85, 3-11
NCPP-105, 3-11 to 3-12
hydraulic power supplies/hydraulic test
stand, 3-12 to 3-13
mobile air-conditioners, 3-6 to 3-10
A/M32C-17, 3-7 to 3-9
NR-5C, 3-8 to 3-10
NR-10, 3-10
mobile electric power plants (MEPPs), 3-3
to 3-6
NC-2A, 3-4
NC-8A, 3-4
NC-10C, 3-5 to 3-6
mobile motor-generator sets (MMGs), 3-6
MMG-lA, 3-6
MMG-2, 3-6
Propellers, 8-9 to 8-26
basic propeller operation, 8-11
basic propeller parts, 8-9 to 8-10
external leakage test, 8-25 to 8-26
forces acting on the propeller, 8-11 to 8-12
internal flow and leakage test, 8-26
propeller balancing and leakage tests, 8-23
to 8-25
propeller maintenance, 8-17 to 8-23
propeller model designation, 8-10
propeller system assemblies, 8-12 to 8-17
Pulsejet engine, 1-3 to 1-4
References, AII-1 to AII-7
Rocket, 1-2
Rotor blade lift, factors affecting, 7-4
Rotor brake, 7-18
Rotor system maintenance, 7-23 to 7-25
Rotor systems, 7-21 to 7-22
Safety methods, hardware, 2-17 to 2-19
cotter pins, 2-19
safety wiring, 2-17 to 2-19
Screwdrivers, 2-6 to 2-7
SE training and licensing program, 3-19
Self-sealing fuel cells, 4-10 to 4-12
Socket sets, 2-3 to 2-5
Starting systems, 6-11 to 6-13
air turbine starter, 6-11
electrical starters, 6-12
hydraulic starters, 6-13
turbine impingement starter, 6-11 to 6-12
Strobex tracking, 7-24
Subsonic/supersonic ducts, 1-12 to 1-13
T58-GE-10 turboshaft engine, 7-8 to 7-11
T700-GE-401 turboshaft engine, 7-11 to 7-12
Tail drive shaft, 7-18
Tail drive shaft inspection, 7-19
Tail gearbox, 7-17
TCP, Tool Control Program, 2-1
Threaded fasteners, 2-12 to 2-15
identification of threaded fasteners, 2-13
installation of threaded fasteners, 2-15
Tools and hardware, 2-1 to 2-19
common hand tools, 2-3 to 2-12
combination wrenches, 2-5 to 2-6
hammers, 2-3
pliers, 2-7 to 2-9
screwdrivers, 2-6 to 2-7
socket sets, 2-3 to 2-5
special tools, 2-9 to 2-12
Tools and hardware—Continued
hardware, 2-12 to 2-19
clamps, 2-16 to 2-17
safety methods, 2-17 to 2-19
threaded fasteners, 2-12 to 2-15
washers, 2-15 to 2-16
Tool Control Program (TCP), 2-1
Turbojet, 1-7 to 1-8
Turboprop engines and propellers, 8-1 to 8-26
propellers, 8-9 to 8-26
basic propeller operation, 8-11
basic propeller parts, 8-9 to 8-10
external leakage test, 8-25 to 8-26
forces acting on the propeller, 8-11
to 8-12
aerodynamic twisting force,
centrifugal force, 8-12
centrifugal twisting force,
propeller vibration, 8-12
thrust bending force, 8-12
torque bending force, 8-12
internal flow and leakage test, 8-26
propeller balancing and leakage
tests, 8-23 to 8-25
external and internal
hydraulic leakage test,
final balance check, 8-24
preliminary balance, 8-24 to
propeller maintenance, 8-17 to 8-23
feathering check, 8-22
fuel governor, pitchlock, and
reverse horsepower
checks, 8-23
NTS check on shutdown,
propeller cleaning, 8-20
propeller removal, 8-20
propeller repair, 8-18 to
propeller servicing, 8-21 to
rigging and adjustment,
unfeathering check, 8-22 to
propeller model designation, 8-10
Turboprop engines and propellers—Continued
propeller system assemblies, 8-12
to 8-17
hub mounting bulkhead
assembly and propeller
assembly, 8-12 to 8-15
propeller control assembly
(integral oil control
assembly), 8-15 to 8-17
spinners and afterbody
assemblies, 8-12
turboprop engines, 8-1 to 8-8
turboprop control systems, 8-6 to
coordinator, 8-8
fuel control, 8-8
power levers, 8-7 to 8-8
turboprop engine assemblies, 8-3 to
power section assembly, 8-3
to 8-4
reduction gear assembly, 8-4
torquemeter assembly, 8-4
turboprop safety systems, 8-4 to 8-6
negative torque signal
(NTS), 8-5
propeller brake, 8-6
safety coupling, 8-5 to 8-6
thrust sensitive signal (TSS),
Torque wrenches, 2-9 to 2-11
Variable-displacement pump, 4-29 to 4-31
Wafer screen filter, 4-28
Washers, 2-15 to 2-16
lock, 2-15
plain, 2-15
special, 2-16
Water washing, 9-22
Wrenches, combination, 2-5 to 2-6
Download PDF
Similar pages