Jet Engine parts and it's working and function

Brayton Cycle
• The Brayton cycle is also known as the constant-pressure cycle.
•
• Point I in the drawing indicates the condition of the air in front of the
engine before it is affected by the inlet duct of the engine.
• After the air enters the inlet duct, it is diffused and the static pressure
increases. This is indicated by point 2, which represents the air condition
at the entrance to the compressor.
• Through the compressor, the air volume is decreased and the pressure is
increased substantially, as shown by the curve from point 2 to point 3. At
point 3, fuel is injected and burned, causing a rapid increase in volume
and temperature. Because of the design of the combustion chamber, the
pressure drops slightly as the velocity of the hot gas mixture increases to
the rear.
• At point 4, the heated gases enter the turbine where energy is
extracted, causing decreases in both pressure and temperature.
• The curve from point 5 to point 6 represents the condition in the
exhaust nozzle as the gases flow out to ambient pressure.
• The difference between the positions of point I and point 6 indicates the
expansion of air caused by the addition of heat from the burning fuel

• The temperature diagram for the Brayton
cycle is shown in Figure.
• Note that the temperature increases
because of compression.
• During combustion, the pressure drops
slightly and the temperature increases at a
rapid rate as a result of the burning of fuel.
• During expansion of the gases through the
turbine, the temperature and pressure of
the gases are reduced to the point where
the gases enter the atmosphere.
• At this time, the temperature is still
considerably above the temperature of the
ambient atmosphere, but it decreases
rapidly as the gases leave the jet nozzle and
go into the atmosphere behind the engine

• A gas turbine engine consists of:
1. An air inlet,
2. Compressor section,
3. Combustion section,
4. Turbine section,
5. Exhaust section,
6. Accessory section, and
7. The systems necessary for starting, lubrication, fuel
supply, and auxiliary purposes, such as anti-icing,
cooling, and pressurization.
• The types of gas turbine engines used to propel and power aircraft are
the turbofan, turboprop, turboshaft and turbojet.
• High speed jet is more efficient at high flight speeds.
• High atmospheric humidity has little or no effect on the operation of a
jet engine.
• Thrust horse power is calculated using net thrust.
• The equation of thrust is derived from Newton's second law of motion.
• The principle of jet reaction is an internal phenomenon.
• The four common types of jet reaction engines are the turbo jet, ram
jet, pulse jet, rocket jet.
• The engine pressure ratio (EPR) is the total pressure ratio across a jet
engine, measured as the ratio of the total pressure at the exit of the
propelling nozzle divided by the total pressure at the entry to the
compressor. Jet engines use either EPR or compressor/fan RPM as an
indicator of thrust.
• The amount of force or thrust produced depends on the amount of
mass of air moved through the engine and the extent to which this air
can be accelerated.
• At the higher engine speeds, thrust increases rapidly with small
increases in RPM.

Turbofan Engine
• The turbofan engine was developed to turn a large fan or set of
fans at the front of the engine and produce about 80 percent of
the thrust from the engine.
• This engine are quieter and has better fuel consumption in the
speed range (.8 Mach).
• Turbofan engines have more than one shaft in the engine; many
are two-shaft engines. This means that there is a compressor and
a turbine that drives it and another compressor and turbine that
drives it.
• These two shafted engines use two spools (a spool is a
compressor and a shaft and turbines that driven that compressor).
• In a two-spool engine, there is a high-pressure spool and a low
pressure spool.
• The low-pressure spool generally contains the fan(s) and the
turbine stages it takes to drive them. The high-pressure spool is
the high-pressure compressor, shaft, and turbines.
• This spool makes up the core of the engine, and this is where the
combustion section is located. The high-pressure spool is also
referred to as the gas generator because it contains the
combustion section.

• Turbofan engines can be low bypass or high bypass.
• The amount of air that is bypassed around the core of the engine determines the bypass ratio.
• The air generally driven by the fan does not pass through the internal working core of the engine.
• The amount of air flow in lb/sec from the fan bypass to the core flow of the engine is the bypass ratio.
• Some low-bypass turbofan engines are used in speed ranges above .8 Mach (military aircraft). These
engines use augmenters or afterburners to increase thrust.
• By adding more fuel nozzles and a flame holder in the exhaust system extra fuel can be sprayed and
burned which can give large increases in thrust for short amounts of time.
• The effect of turbofan design is to increase power/ weight ratio and improve thrust specific fuel
consumption.

Turboshaft Engine
• The turboshaft engine is a gas turbine engine made to transfer
horsepower to a shaft that turns a helicopter transmission or is an
onboard auxiliary power unit (APU).
• An APU is used on turbine-powered aircraft to provide electrical
power and bleed air on the ground and a backup generator in
ﬂight.
• Turboshaft engines can come in many different styles, shapes, and
horsepower ranges.

PULSE JET ENGINE
• Pulse jet is a basic type of jet engine which comprises of a flapping air entry section.
• The pulse jet engine is somewhat more complex than the ram jet since it has a grill of shutters
located at the inlet of the engine. However, the operation of the pulse jet is easier to
understand. The shutters, which are kept open with springs, allow air to enter the combustion
chamber. As the air is packed into the combustion chamber, it is mixed with fuel and ignited. The
shutters are forced closed by the pressure of the exhaust gases when combustion takes place.
• Consequently, the exhaust gases can only move down the tailpipe and out the exhaust. Then the
springs force the shutters to reopen, allowing more fresh air to enter, and the cycle repeats. The
length of the tailpipe of the pulse jet regulates the frequency of the engine. Fuel flow is
continuous, but flame propagation is intermittent, since the pulse jet operates in a step-by-step
cycle. This is the only form of jet propulsion that operates by intermittent power surges, utilizing
explosive rather than progressive or continuous combustion. However, in most pulse jet
engines, the cycles per second are rather high and the net effect is practically continuous thrust.
Pulse jet engines provide thrust for some guided missiles.

Turbo-prop Engines
• The turboprop engine is a gas turbine engine that turns a propeller
through a speed reduction gear box. This type of engine is most efficient
in the 300 to 400 mph speed range and can use shorter runways.
• Approximately 80 to 85 percent of the energy developed by the gas
turbine engine is used to drive the propeller. The rest of the available
energy exits the exhaust as thrust.
• A turboprop power plant propeller accounts for 75 to 85 percent of the
total thrust output.
• By adding the horsepower developed by the engine shaft and the
horsepower in the exiting thrust, the answer is equivalent shaft
horsepower
• At low altitude and below sonic speed turboprop is more efficient.
• In normal cruising speed ranges, the propulsive efficiency of a turboprop
engine decreases as speed increases.
• In a turboprop aircraft, the propeller generates thrust by imparting small
acceleration to large amount of air.
• In a turboprop engine only a small amount of jet thrust is available from
the exhaust system.
• In a turboprop engine reverse thrust action is obtained by changing the
pitch of the propeller blade.
• On cold days Short take off run is required.
• When the jet velocity increases,
the thrust produced by a
turbojet Increases.
• Maximum kinetic energy is
wasted in a turbojet.
• Turbojet engines are very
efficient at high speed & high
altitude flying because of high
propulsive efficiency.
• Turbojet is lighter in terms of
specific weight per kg thrust.
Turbojet

Ramjet Engine
• The simplest type of air-breathing reaction engine. Air entering the
front of the engine at a high velocity has fuel sprayed into it and
ignited. A barrier formed by the incoming air forces the expanding
gases to leave through the nozzle at the rear. The energy added by the
burning fuel accelerates the air and produces a forward thrust.
Ramjet engines are used in some military unmanned aircraft that are
initially boosted to a speed high enough for the engine to function.

Efficiencies
The efficiency of any engine can be described as the
output divided by the input. One of the main
measures of turbine engine efficiency is the amount of
thrust produced or generated, divided by the fuel
consumption. This is called thrust specific fuel
consumption, or tsfc. The tsfc is the amount of fuel
required to produce 1 lb [0.004 45 kN] of thrust and
can be calculated as follows:
Tsfc= wf/fn
where wf= fuel flow, lb/h [kg/h]
Fn = net thrust, lb [kg]
This leads to the conclusion that the more thrust
obtained per pound of fuel, the more efficient the
engine is.
Specific fuel consumption is made up of a number of
other efficiencies. The two major factors affecting the
tsfc are propulsive efficiency and cycle efficiency.
• Propulsive efficiency is the amount of thrust
developed by the jet nozzle compared with the
energy supplied to it in a usable form.
• In other words, the propulsive efficiency is the
percentage of the total energy made available
by the engine which is effective in propelling
the engine.
• Propulsive efficiency can also be expressed as:
Work completed/ Work completed+ work
wasted in the exhaust
• Propulsive efficiency is defined as internal
engine efficiency.
• The propulsive efficiency of a by pass engine
is more than that of a turbojet because of
lower velocity of jet efflux.
• In normal cruising range the propulsive
efficiency of a turbojet decreases as speed
increases.
Propulsive Efficiency

Cycle Efficiency
Cycle efficiency is the amount of energy put into a usable form in
comparison with the total amount of energy available in the fuel. It
involves combustion efficiency, thermal efficiency, mechanical
efficiency, compressor efficiency, etc. It is, in effect, the overall
efficiency of the engine components starting with the compressor and
going through the combustion chamber and turbine. The job of these
components is to get the energy in the fuel into a form which the jet
nozzle can turn into thrust.

Air Entrance
• The air entrance is designed to conduct incoming air to the
compressor with a minimum energy loss resulting from drag or ram
pressure loss; that is, the ﬂow of air into the compressor should be
free of turbulence to achieve maximum operating efficiency.
• Proper inlet design contributes materially to aircraft performance by
increasing the ratio of compressor discharge pressure to duct inlet
pressure. This is also referred to as the compressor pressure ratio.
• This ratio is the outlet pressure divided by the inlet pressure. The
amount of air passing through the engine is dependent upon three
factors:
1. The compressor speed (rpm)
2. The forward speed of the aircraft
3. The density of the ambient (surrounding) air
• Duct pressure efficiency ratio is defined as ability of the duct to
convert kinetic energy of the air stream into static pressure energy
without loss in total pressure.

• Turbine inlet type is dictated by the type of gas turbine
engine.
• A high-bypass turbofan engine inlet is completely different
from a turboprop or turboshaft inlet.
• Large gas turbine powered aircraft almost always have a
turbofan engine. The inlet on this type of engine is bolted to
the front (A flange) of the engine. These engines are
mounted on the wings, or nacelles, on the aft fuselage, and a
few are in the vertical fin.
• Since on most modern turbofan engines the huge fan is the
first part of the aircraft the incoming air comes into contact
with, icing protection must be provided.
• This prevents chucks of ice from forming on the leading edge
of the inlet, breaking loose, and damaging the fan.
• Warm air is bled from the engine’s compressor and is ducted
through the inlet to prevent ice from forming. If inlet guide
vanes are used to straighten the air flow, then they also have
anti-icing air flowing through them. The inlet also contains
some sound-reducing materials that absorb the fan noise and
make the engine quieter.
• Turboprops and turboshafts can use an inlet screen to help filter out
ice or debris from entering the engine. A deflector vane and a
heated inlet lip are used to prevent ice or large chunks from
entering the engine.
• On military aircraft, the divided entrance permits the use of very
short ducts with a resultant small pressure drop through skin
friction.
• Military aircraft can fly at speeds above Mach 1, but the airflow
through the engine must always stay below Mach 1. Supersonic air
flow in the engine would destroy the engine.
• By using convergent and divergent shaped ducts, the air flow is
controlled and dropped to subsonic speeds before entering the
engine.
• Supersonic inlets are used to slow the incoming engine air to less
than Mach 1 before it enters the engine.
• Helicopters generally use a bell mouth inlet duct. These are
generally used for engine rating purpose.
• The air inlet duct is generally rated in Ram recovery point.
• Buzz is an airflow instability which takes place in air inlet duct at low
air flow velocity.
• A subsonic business jet will have an inlet duct of divergent design.

Compressor Section
• The compressor section of the gas turbine
engine has many functions.
• Its primary function is to supply air in sufficient
quantity to satisfy the requirements of the
combustion burners.
• The compressor must increase the pressure of
the mass of air received from the air inlet duct,
and then, discharge it to the burners in the
quantity and at the pressures required.
• The compressor stall is more severe at Low
altitude and high rpm.
• A secondary function of the compressor is to supply
bleed air for various purposes in the engine and aircraft.
• The bleed-air is taken from any of the various pressure
stages of the compressor.
• Air is often bled from the final or highest pressure stage
since, at this point, pressure and air temperature are at a
maximum.
• Bleed air is utilized in a wide variety of ways. Some of
the current applications of bleed air are:
1. Cabin pressurization, heating, and cooling;
2. Deicing and anti-icing equipment;
3. Pneumatic starting of engines; and
4. Auxiliary drive units (ADU)

Compressor Types
• The two principal types of compressors currently being used in gas
turbine aircraft engines are centrifugal flow and axial flow.
• The centrifugal-flow compressor achieves its purpose by picking up the
entering air and accelerating it outwardly by centrifugal action.
• The axial-flow compressor compresses air while the air continues in its
original direction of flow, thus avoiding the energy loss caused by
turns.
• A stage in a compressor is considered to be a rise in pressure.
• Variable area entry guide vanes provide surge free compressor
operation.

Centrifugal-Flow Compressors
• The centrifugal-ﬂow compressor consists of an
impeller (rotor), a diffuser (stator), and a
compressor manifold.
• Centrifugal compressors have a high pressure rise
per stage that can be around 8:1.
• Generally centrifugal compressors are limited to
two stages due to efficiency concerns.
• The two main functional elements are the impeller
and the diffuser. Although the diffuser is a separate
unit and is placed inside and bolted to the manifold,
the entire assembly (diffuser and manifold) is often
referred to as the diffuser.
• The impeller is usually made from forged
aluminum alloy, heat treated, machined, and
smoothed for minimum ﬂow restriction and
turbulence.

• The impeller is fabricated from a single forging.
The impeller, whose function is to pick up and
accelerate the air outwardly to the diffuser, may be
either of two types—single entry or double entry.
• The principal differences between the two types
of impellers are size and ducting arrangement. The
double-entry type has a smaller diameter, but is
usually operated at a higher rotational speed to
assure sufficient airﬂow.
• The single-entry impeller, permits convenient
ducting directly to the impeller eye (inducer vanes)
as opposed to the more complicated ducting
necessary to reach the rear side of the double-
entry type.
• Although slightly more efficient in receiving air, the
single entry impeller must be large in diameter to
deliver the same quantity of air as the double-
entry type. This, of course, increases the overall
diameter of the engine

• Included in the ducting for double-entry compressor engines is the plenum
chamber. This chamber is necessary for a double-entry compressor because the
air must enter the engine at almost right angles to the engine axis.
• Therefore, in order to give a positive flow, the air must surround the engine
compressor at a positive pressure before entering the compressor.
• Included in some installations as necessary parts of the plenum chamber are
the auxiliary air-intake doors (blow-in doors).
• These blow-in doors admit air to the engine compartment during ground
operation, when air requirements for the engine are in excess of the airflow
through the inlet ducts. The doors are held closed by spring action when the
engine is not operating.
• During operation, however, the doors open automatically whenever engine
compartment pressure drops below atmospheric pressure.
• During takeoff and flight, ram air pressure in the engine compartment aids the
springs in holding the doors closed.

The diffuser is an annular chamber provided with a number of vanes forming a series of divergent passages into the
manifold. The diffuser vanes direct the ﬂow of air from the impeller to the manifold at an angle designed to retain
the maximum amount of energy imparted by the impeller. They also deliver the air to the manifold at a velocity and
pressure satisfactory for use in the combustion chambers. Refer to Figure and note the arrow indicating the path of
airﬂow through the diffuser, then through the manifold.

The compressor manifold shown in Figure diverts the ﬂow of air from the diffuser, which is an integral part of the
manifold, into the combustion chambers. The manifold has one outlet port for each chamber so that the air is
evenly divided. A compressor outlet elbow is bolted to each of the outlet ports. These air outlets are constructed
in the form of ducts and are known by a variety of names, such as air outlet ducts, outlet elbows, or combustion
chamber inlet ducts. Regardless of the terminology used, these outlet ducts perform a very important part of the
diffusion process; that is, they change the radial direction of the airﬂow to an axial direction, in which the
diffusion process is completed after the turn. To help the elbows perform this function in an efficient manner,
turning vanes (cascade vanes) are sometimes fitted inside the elbows. These vanes reduce air pressure losses by
presenting a smooth, turning surface.

Axial-Flow Compressor
• The axial-flow compressor has two main elements: a rotor and a
stator.
• The rotor has blades fixed on a spindle. These blades impel air
rearward in the same manner as a propeller because 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.
• From inlet to exit, the air ﬂows along an axial path and is
compressed at a ratio of approximately 1.25:1 per stage.
• The action of the rotor increases the compression of the air at
each stage and accelerates it rearward through several stages.
• With this increased velocity, energy is transferred from the
compressor to the air in the form of velocity energy.
• The stator blades act as diffusers at each stage, partially
converting high velocity to pressure. Each consecutive pair of rotor
and stator blades constitutes a pressure stage. The number of
rows of blades (stages) is determined by the amount of air and
total pressure rise required. Compressor pressure ratio increases
with the number of compression stages. Most engines utilize up to
16 stages and more.

The stator has rows of vanes, which are in turn attached inside an enclosing case. The stator vanes, which are stationary,
project radially toward the rotor axis and fit closely on either side of each stage of the rotor blades. In some cases, the
compressor case, into which the stator vanes are fitted, is horizontally divided into halves. Either the upper or lower half
may be removed for inspection or maintenance of rotor and stator blades.
The function of the stator vanes is to receive air from the air inlet duct or from each preceding stage and increase the
pressure of the air and deliver it to the next stage at the correct velocity and pressure. They also control the direction of air
to each rotor stage to obtain the maximum possible compressor blade efficiency. The first stage rotor blades can be
preceded by an inlet guide vane assembly that can be fixed or variable.
The guide vanes direct the airflow into the first stage rotor blades at the proper angle and impart a swirling motion to the
air entering the compressor. This pre swirl, in the direction of engine rotation, improves the aerodynamic characteristics of
the compressor by reducing drag on the first stage rotor blades. The inlet guide vanes are curved steel vanes usually
welded to steel inner and outer shrouds.
At the discharge end of the compressor, the stator vanes are constructed to straighten the airflow to eliminate turbulence.
These vanes are called straightening vanes or the outlet vane assembly. The casings of axial-flow compressors not only
support the stator vanes and provide the outer wall of the axial path the air follows, but they also provide the means for
extracting compressor air for various purposes. The stator vanes are usually made of steel with corrosion- and erosion
resistant qualities. Quite frequently, they are shrouded(enclosed) by a band of suitable material to simplify the fastening
problem. The vanes are welded into the shrouds, and the outer shroud is secured to the compressor housing inner wall by
radial retaining screws.

• blade attachment to the rotor disk rims varies, but they are
commonly ftted into disks by either bulb-type or fir-tree
methods. The blades are then locked into place by differing
methods.
• Compressor blade tips are reduced in thickness by cutouts,
referred to as blade profiles. These profiles prevent serious
damage to the blade or housing should the blades contact the
compressor housing.
• This condition can occur if rotor blades become excessively
loose or if rotor support is reduced by a malfunctioning
bearing.
• Even though blade profiles greatly reduce such possibilities,
occasionally a blade may break under stress of rubbing and
cause considerable damage to compressor blades and stator
vane assemblies.
• The blades vary in length from entry to discharge because the
annular working space (drum to casing) is reduced
progressively toward the rear by the decrease in the casing
diameter. This feature provides for a fairly constant velocity
through the compressor, which helps to keep the ﬂow of air
constant.

• The combination of the compressor stages and
turbine stages on a common shaft is an engine
referred to as an engine spool. The common shaft is
provided by joining the turbine and compressor
shafts by a suitable method. The engine’s spool is
supported by bearings, which are seated in suitable
bearing housings.
• There are two configurations of the axial
compressor currently in use: the single rotor/spool
and the dual rotor/spool, sometimes referred to as
solid spool and split spool (two spool, dual spool).
• Some high-volume turboprop and turbojet engines
are equipped with two-spool or split compressors.
When these engines are operated at high altitude,
the low-pressure rotor will increase in speed as the
compressor load decreases in the lower density air.
• One version of the solid-spool (one spool) compressor
uses variable inlet guide vanes. Also, the fIrst few rows
of stator vanes are variable.
• The main difference between variable inlet guide vane
(VIGV) and a variable stator vane (VSV) is their position
with regard to the rotor blades.
• VIGV are in front of the rotor blades, and VSV are
behind the rotor blades. The angles of the inlet guide
vanes and the first several stages of the stator vanes are
can be variable.
• During operation, air enters the front of the engine and
is directed into the compressor at the proper angle by
the variable inlet guide and directed by the VSV. The air
is compressed and forced into the combustion section.
• A fuel nozzle that extends into each combustion liner
atomizes the fuel for combustion. These variables are
controlled in direct relation to the amount of power the
engine is required to produce by the power lever
position

Advantages of centrifugal flow compressor
• High pressure rise per stage,
• Efficiency over wide rotational speed range,
• Simplicity of manufacture and low cost,
• Low weight, and
• Low starting power requirements.
Disadvantages of centrifugal flow
compressor
• Its large frontal area for a given airflow and
• Losses in turns between stages.
Advantages of axial flow compressor
• High peak efficiencies;
• Small frontal area for given airflow;
• Straight-through flow, allowing high ram
efficiency; and
• Increased pressure rise by increasing number
of stages,with negligible losses.
Disadvantages of axial-flow compressor
• Good efficiencies over only narrow rotational speed
range,
• Difficulty of manufacture and high cost,
• Relatively high weight, and
• High startig power requirements (partially overcome by
split compressors).

Diffuser
• The diffuser is the divergent section of the engine after the compressor and
before the combustion section. It has the all-important function of reducing
high-velocity compressor discharge air to increased pressure at a slower
velocity. This prepares the air for entry into the flame burning area of the
combustion section at a lower velocity so that the flame of combustion can
burn continuously. If the air passed through the flame area at a high velocity, it
could extinguish the flame.
• Kinetic energy contained by the air in a centrifugal compressor is converted in
to pressure energy by the diffuser.
• In a turbine engine the purpose of the diffuser section is to increase pressure
and reduce velocity.

Combustion Section
• The combustion section houses the combustion process, which raises the
temperature of the air passing through the engine. This process releases energy
contained in the air/ fuel mixture.
• The major part of this energy is required at the turbine or turbine stages to drive the
compressor. About ²⁄³ of the energy is used to drive the gas generator compressor.
• The remaining energy passes through the remaining turbine stages that absorb more
of the energy to drive the fan, output shaft, or propeller. Only the pure turbojet
allows the air to create all the thrust or propulsion by exiting the rear of the engine in
the form of a high-velocity jet.
• In a gas turbine engine combustion chamber contributes maximum forward thrust.
• In the combustion chamber of a gas turbine engine local deceleration of air is
required to provide a low velocity zone in which the flame can burn.

• The primary function of the combustion
section is, of course, to burn the fuel/air
mixture, thereby adding heat energy to the
air. To do this efficiently, the combustion
chamber must:
Provide the means for proper mixing of the
fuel and air to assure good combustion
Burn this mixture efficiently,
Cool the hot combustion products to a
temperature that the turbine inlet guide
vanes/blades can withstand under
operating conditions, and
Deliver the hot gases to the turbine section
• The location of the combustion section is directly
between the compressor and the turbine sections.
The combustion chambers are always arranged
coaxially with the compressor and turbine
regardless of type, since the chambers must be in a
through-ﬂow position to function efficiently. All
combustion chambers contain the same basic
elements:
1. Casing
2. Perforated inner liner
3. Fuel injection system
4. Some means for initial ignition
5. Fuel drainage system to drain off unburned fuel
after engine shutdown
• There are currently three basic types of combustion
chambers:
1. Can type
2. Can-annular type
3. Annular type

Can type combustion chamber
• The can-type combustion chamber
is typical of the type used on
turboshaft and APUs.
• Each of the can-type combustion
chambers consists of an outer case
or housing, within which there is a
perforated stainless steel (highly
heat resistant) combustion
chamber liner or inner liner.
• The outer case is removed to
facilitate liner replacement.

• Older engines with several combustion cans
had each can with interconnector (ﬂame
propagation) tube, which was a necessary part
of the can-type combustion chambers.
• Since each can is a separate burner operating
independently of the other cans, there must
be some way to spread combustion during the
initial starting operation.
• This is accomplished by interconnecting all the
chambers. As the ﬂame is started by the spark
igniter plugs in two of the lower chambers, it
passes through the tubes and ignites the
combustible mixture n the adjacent chamber,
and continues until all the chambers are
burning.

• The liners of the can-type combustors have perforations of
various sizes and shapes, each hole having a specific purpose
and effect on ﬂame propagation within the liner.
• The air entering the combustion chamber is divided by the
proper holes, louvers, and slots into two 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, cooling the combustion gases
from about 3,500 °F to near 1,500 °F.
• To aid in atomization of the fuel, holes are provided around
the fuel nozzle in the dome or inlet end of the can-type
combustor 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 also tends to control the ﬂame pattern by
keeping it centered in the liner, thereby preventing burning of
the liner walls.
• Some provision is always made in the
combustion chamber case for installation of a
fuel nozzle.
• The fuel nozzle delivers the fuel into the liner
in a finely atomized spray. The more the spray
is atomized, the more rapid and efficient the
burning process is.
• Two types of fuel nozzle currently being used
in the various types of combustion chambers
are the simplex nozzle and the duplex nozzle.
• There are usually two igniters mounted on the
boss provided on each of the chamber
housings. The igniters must be long enough to
protrude from the housing into the
combustion chamber.

• The forward face of each chamber presents six
apertures, which align with the six fuel nozzles
of the corresponding fuel nozzle cluster.
• These nozzles are the dual-orifice (duplex)
type requiring the use of a flow-divider
(pressurizing valve)
• Around each nozzle are pre swirl vanes for
imparting a swirling motion to the fuel spray,
which results in better atomization of the fuel,
better burning, and efficiency.
• The swirl vanes function to provide two
effects imperative to proper flame
propagation:
1 High flame speed—better mixing of air and
fuel, ensuring spontaneous burning.
2 Low air velocity axially—swirling eliminates
overly rapid flame movement axially.

• 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
engine.
• The flow of air through the holes and louvers of the
can annular chambers, is almost identical with the
flow through other types of burners.
• Special baffling is used to swirl the combustion
airflow and to give it turbulence.
• Figure shows the flow of combustion air, metal
cooling air, and the diluent or gas cooling air. The air
flow direction is indicated by the arrows.

• The basic components of an annular
combustion chamber are a housing and a
liner, as in the can type.
• The liner consists of an undivided circular
shroud extending all the way around the
outside of the turbine shaft housing.
• The chamber may be constructed of heat-
resistant materials, which are sometimes
coated with thermal barrier materials, such
as ceramic materials.
• Modern turbine engines usually have an
annular combustion chamber. The annular
combustion chamber also uses louvers and
holes to prevent the ﬂame from contacting
the side of the combustion chamber.

Turbine Section
• The turbine transforms a portion of the kinetic (velocity) energy of the exhaust gases into
mechanical energy to drive the gas generator compressor and accessories.
• The sole purpose of the gas generator turbine is to absorb approximately 60 to 70 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 (i.e., compressor size and type, number of accessories, and the load
applied by the other turbine stages).
• These turbine stages can be used to drive a low-pressure compressor (fan), propeller, and
shaft. The turbine section of a gas turbine engine is located aft, or downstream, of the
combustion chamber.
• Specifically, it is directly behind the combustion chamber outlet.
• Turbines are exposed to very high temperatures. The ability of a metal to withstand
extreme changes in temperature in short periods of time is known as creep strength

• The turbine assembly consists of two basic
elements: turbine inlet guide vanes and turbine
blades.
• The stator element is known by a variety of
names, of which turbine inlet nozzle vanes,
turbine inlet guide vanes, and nozzle diaphragm
are three of the most commonly used.
• The turbine inlet nozzle vanes are located
directly aft of the combustion chambers and
immediately forward of the turbine wheel. This
is the highest or hottest temperature that
comes in contact with metal components in the
engine. The turbine inlet temperature must be
controlled or damage will occur to the turbine
inlet vanes.

• After the combustion chamber has introduced the heat energy into the
mass airflow and delivered it evenly to the turbine inlet nozzles, the
nozzles must prepare the mass air flow to drive the turbine rotor. The
stationary vanes of the turbine inlet nozzles are contoured and set at
such an angle that they form a number of small nozzles discharging gas
at extremely high speed; thus, the nozzle converts a varying portion of
the heat and pressure energy to velocity energy that can then be
converted to mechanical energy through the turbine blades.
• The second purpose of the turbine inlet 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 it
is still rotating, it is essential to aim the gas in the general direction of
turbine rotation.

• The turbine inlet nozzle assembly consists of an inner shroud and an
outer shroud between which the nozzle vanes are fixed. The number
and size of inlet vanes employed vary with different types and sizes
of engines.
• The vanes of the turbine inlet nozzle may be assembled between the
outer and inner shrouds or rings in a variety of ways.
• Although the actual elements may vary slightly in configuration and
construction features, there is one characteristic peculiar to all
turbine inlet nozzles: the nozzle vanes must be constructed to allow
thermal expansion.
• Otherwise, there would be severe distortion or warping of the metal
components because of rapid temperature changes. The thermal
expansion of turbine nozzles is accomplished by one of several
methods. One method necessitates loose assembly of the
supporting inner and outer vane shrouds.
• Each vane fts into a contoured slot in the shrouds, which conforms
to the airfoil shape of the vane. These slots are slightly larger than
the vanes to give a loose ft. For further support, the inner and outer
shrouds are encased by inner and outer support rings, which provide
increased strength and rigidity. These support rings also facilitate
removal of the nozzle vanes as a unit. Without the rings, the vanes
could fall out as the shrouds were removed

• The rotor element of the turbine section consists
essentially of a shaft and a wheel.
• The turbine wheel is a dynamically balanced unit
consisting of blades attached to a rotating disk. The disk,
in turn, is attached to the main power-transmitting shaft
of the engine.
• The exhaust gases leaving the turbine inlet nozzle vanes
act on the blades of the turbine wheel, causing the
assembly to rotate at a very high rate of speed.
• The high rotational speed imposes severe centrifugal
loads on the turbine wheel, and at the same time the
elevated temperatures result in a lowering of the
strength of the material.
• Consequently, the engine speed and temperature must
be controlled to keep turbine operation within safe
limits.

• The turbine shaft is usually fabricated from alloy steel.
• It must be capable of absorbing the high torque loads that are
exerted on it. The methods of connecting the shaft to the turbine disk
vary.
• In one method, the shaft is welded to the disk, 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 fits a machined
surface on the disk face. Then, the bolts are inserted through holes in
the shaft hub and anchored in tapped holes in the disk. Of the two
connection methods, bolting is more common.

• 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. Many turbine
blades are cast as a single crystal, which
gives the blades better strength and heat
properties.
• Heat barrier coating, such as ceramic
coating, and air ﬂow cooling help keep the
turbine blades and inlet nozzles cooler.
• This allows the exhaust temperature to be
raised, increasing the efficiency of the
engine.
• Figure shows a turbine blade with air
holes for cooling purposes.

• In turbine rotor construction, it occasionally becomes necessary to utilize
turbines of more than one stage. A single turbine wheel often cannot
absorb enough power from the exhaust gases to drive the components
dependent on the turbine for rotative power; thus, it is necessary to add
additional turbine stages.
• A turbine stage consists of a row of stationary vanes or nozzles, followed
by a row of rotating blades. In some models of turboprop engine, as many
as five turbine stages have been utilized successfully.
• It should be remembered that, regardless of the number of wheels
necessary for driving engine components, there is always a turbine nozzle
preceding each wheel.

• In the single-stage rotor turbine, the power is
developed by one turbine rotor, and all engine-
driven parts are driven by this single wheel. This
arrangement is used on engines where the need for
low weight and compactness predominates. This is
the simplest version of the pure turbojet engine.
• In multiple spool engines, each spool has its own
set of turbine stages. Each set of turbine stages
turns the compressor attached to it.
• Most turbofan engines have two spools:
• low pressure (fan shaft a few stages of compression
and the turbine to drive it) and high pressure (high
pressure compressor shaft and high pressure
turbine)

• The turbine casing encloses the
turbine wheel and the nozzle
vane assembly, and at the same
time gives either direct or
indirect support to the stator
elements of the turbine section.
It always has ﬂanges provided
front and rear for bolting the
assembly to the combustion
chamber housing and the
exhaust cone assembly,
respectively.

Exhaust Section
• The exhaust section must direct the flow of hot gases rearward
in such a manner as to prevent turbulence and, at the same
time, impart a high final or exit velocity to the gases.
• In performing the various functions, each of the components
affects the flow of gases in different ways.
• The exhaust section is located directly behind the turbine
section and ends when the gases are ejected at the rear in the
form of a high-velocity exhaust gases.
• The components of the exhaust section include the exhaust
cone, tailpipe (if required), and the exhaust nozzle.
• The exhaust cone collects the exhaust gases discharged from the
turbine section and gradually converts them into a solid flow of
gases. In performing this, the velocity of the gases is decreased
slightly and the pressure increased.
• This is due to the diverging passage between the outer duct and
the inner cone; that is, the annular area between the two units
increases rearward. The exhaust cone assembly consists of an
outer shell or duct, an inner cone, three or four radial hollow
struts or fins, and the necessary number of tie rods to aid the
struts in supporting the inner cone from the outer duct.

• The outer shell or duct is usually made of stainless steel and is attached to the rear
ﬂange of the turbine case. This element collects the exhaust gases and delivers
them directly to the exhaust nozzle.
• The duct must be constructed to include such features as a predetermined
number of thermocouple bosses for installing exhaust temperature
thermocouples, and there must also be insertion holes for the supporting tie rods.
• In some cases, tie rods are not used for supporting the inner cone. 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,
respectively.
• The radial struts actually have a two fold function. They not only support the inner
cone in the exhaust duct, but they also perform the important function of
straightening the swirling exhaust gases that would otherwise leave the turbine at
an angle of approximately 45°

• The centrally located inner cone fits rather closely
against the rear face of the turbine disk, preventing
turbulence of the gases as they leave the turbine wheel.
The cone is supported by the radial struts. In some
configurations, a small hole is located in the exit tip of
the cone.
• This hole allows cooling air to be circulated from the aft
end of the cone, where the pressure of the gases is
relatively high, into the interior of the cone and
consequently 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 used for cooling the turbine wheel return to
the main path of ﬂow by passing through the clearance
between the turbine disk and the inner cone.
• The exhaust cone assembly is the terminating
component of the basic engine. The remaining
component (the exhaust nozzle) is usually considered an
airframe component.

• The tailpipe is usually constructed so that it is semiﬂexible. On some
tailpipes, a bellows arrangement is incorporated in its construction,
allowing movement in installation, maintenance, and in thermal
expansion. This eliminates stress and warping which would otherwise
be present.
• The heat radiation from the exhaust cone and tailpipe could damage
the airframe components surrounding these units. For this reason,
some means of insulation had to be devised. There are several
suitable methods of protecting the fuselage structure; two of the
most common are insulation blankets and shrouds.

• The insulation blanket, illustrated in Figures consists of several layers of
aluminum foil, each separated by a layer of fiberglass or some other suitable
material.
• Although these blankets protect the fuselage from heat radiation, they are used
primarily to reduce heat loss from the exhaust system. The reduction of heat loss
improves engine performance.
• There are two types of exhaust nozzle designs: the converging design for
subsonic gas velocities and the converging diverging design for supersonic gas
velocities.
• The exhaust nozzle opening may be of either fixed or variable area. The fixed-
area type is the simpler of the two exhaust nozzles since there are no moving
parts.
• The outlet area of the fixed exhaust nozzle is very critical to engine performance.
If the nozzle area is too large, thrust is wasted; if the area is too small, the engine
could choke or stall.
• A variable-area exhaust nozzle is used when an augmenter or afterburner is
used due to the increased mass of ﬂow when the afterburner is activated. It
must increase its open area when the afterburner is selected. When the
afterburner is off, the exhaust nozzle closes to a smaller area of opening.
• In case of variable area jet nozzle, during maximum power rating the nozzle area
is minimum.

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