Gas Turbine Engine

Alexandra Slabutu DT011/1 AVTE 1102 – Karl Casey 05 December 2013

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Bachelor of Engineering Technology

(B.Eng.Tech) Aviation Technology

Aircraft Engines

The Gas Turbine Turbojet Engine -The flow of air through a gas turbine turbojet engine-


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Declaration

This is an original work. All references and assistance are acknowledged.

Date: 5 Dec 2013

Candidate Name: Alexandra Slabutu

If an assignment or project or part of an assignment or

project has been plagiarised from any source, this will result

in a fail for that assignment or project. (DIT, 2011)


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Contents

Introduction ___________________________________________________4

Intake _____________________________________________________4

Compressor _____________________________________________________4

Combustion_____________________________________________________7

Turbine ________________________________________________________8

Exhaust ________________________________________________________9

Conclusion _____________________________________________________ 9

Reference _____________________________________________________ 10


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Introduction

Air is the fluid used in Gas Turbine engines. Modern Gas Turbine engines have

developed many techniques to gain maximum efficiency and work from the air as it

flows through the engine.

The flow of air through a gas turbine turbojet engine will be described in this

assignment, as well as the constructional arrangement of the engine and the effect

such has on the flow of the air through the engine.

The description of the main components of a gas turbine engine such as the inlet,

compressor, combustion, turbine and exhaust will be discussed in this assignment as

upon their functionality, design and construction, the efficiency and the airflow

through the engine depends.

Intake

The main purpose of the inlet is to drive the air into the compressor at a lower speed,

approximately between 0.4 – 0.5 Mach despite of the aircrafts speed, with very little

loss of energy. Otherwise inefficient compressor performance will be resulted.

Various types of intakes are found on today’s aircraft generation such as Pitot, divided

entrance, variable geometry and external/internal compression. The most commonly

used intake in commercial aviation, at which aircraft subsonic speeds are

accomplished, is the Pitot type inlet. Its duct design is divergent in shape. Another

divergent type intake is the Divided Entrance, found on some single engine aircrafts.

However for supersonic speed aircrafts the duct shape is designed as convergent-

divergent intakes as their function is to restrict the amount of supersonic air flow

entering the compressor as it is not necessary and also to overcome the inlet shock

waves.

Compressor

(Compressor Section)


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The compressor function is to encourage the mass of airflow throughout the engine

and also compressing the air, converting it into high pressure and temperature as it

passes through each stage. A stage consists of a stator and a rotor.

There are three main types of compressors used in gas turbine engines: centrifugal

compressors, axial flow compressors and a combination of both. However the use of

axial flow compressors is more used as it promotes pressure ratios greater than 30:1,

thus making it more efficient in the airline industry. This type of compressor is further

subdivided into the following: Single-spool axial flow compressor; Dual-spool axial

flow compressor; Twin spool low/high bypass compressor; Triple spool high bypass

compressor. Each of this compressor types are manufactured and used in accordance

to the engine purpose and requirements.

According to the COMPRESSOR SECTION ARTICLE the axial-flow compressor

consists of a stator and a rotor which can be found in a drum or a disc type.


The construction of these compressors can be made of various distinct materials as it

depends on the operating environment, temperature and the load exerted on them.

Rings that are flanged to fit against each other in order to hold together the whole

assembly through bolts is known as the drum- type rotor. This type of design is

tolerable for low speed compressors where centrifugal stresses are also low.

According to (Aircav, n.d.) the construction of the disc type rotor is made up of: stub

shafts, discs, blades, ducts, air vortex spoilers, spacers, tie bolts, torque cones. The

rotor blades are mostly manufactured of stainless steel forgings, however in some

cases they might be made of titanium in the colder (forward) area of the compressor,

but not rearwards as this material may have the tendency of igniting if friction occurs.

The various methods of attaching the blades into the disc rim are dovetail, fir-tree and

bulb root type design. Those are then secured into place with pins, keys, screws or

locking wires. However the blades fitting must not be too tight into the disc as they

are forced to seat together by the centrifugal force produced during the engine

operation.


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To obtain high efficiency is imperative that the area between the casing and the rotor

blade is free of any debris. Therefore different techniques have been used by

manufacturers to prevent that from happening, according to the article Compressor

Section there are: ‘knife edge blades design that wear away to form their own

clearances as they expand from the heat generated by air compression…coat the inner

surface of the compressor case with a soft material (Teflon) that can be worn away

without damaging the blade …’.

Confirming to TTS (2011) in the axial flow compressor the stator vanes are mounted

into the compressor casing or into stator vane retaining rings. The blades are then

“often shrouded at their inner ends to minimize the vibrational effect of flow

variations on the longer vanes”. Stator vanes are found as being more variable at the

forward section of the compressor if this has a high number of stages and it can also

be fixed. The casing material is made of aluminium and the vanes from stainless steel

or nickel based alloys.

The airflow as it passes each stage through the axial flow compressor; it is accelerated

by the rotors, thus creating an increase in pressure, temperature and velocity.

However due to the stator divergent spacing between the vanes, its impact on air

results in a reduced velocity but a rise in temperature and velocity therefore this

process multiplies with each followed stage of the compressor. The air annulus is

experiencing a gradual reduction across the axial flow compressor, giving it a

convergent shape, due to the decrease in length of the blades from forward to

rearwards in order to maintain the axial velocity of the air, constant. The main

importance of the compressor is to rise the temperature and pressure but to reduce the

volume of air.

The centrifugal compressor is lesser used nowadays in the airline industry due to its

CRP of up to 10:1 but it’s commonly found on smaller aircrafts. According to the

TTS (2011), the centrifugal compressor is made up of three parts: the impeller (rotor),

the diffuser (stator) and the manifold or casing. The impeller’s purpose is to suck up

the air and accelerate it outward to the diffuser. It can be found as a double or single

entry and the dissimilarities between them being the duct arrangement and the size.


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The diffuser helps the air passing through its divergent vanes thus converting the air

velocity into an increasing pressure and heat energy. The airflow is then directed

through the casing to the combustion section or the following stage of the

compression.

Combination compressors have been designed in order to take the advantages of the

great benefits from both centrifugal and axial flow compressors and eliminate the

disadvantages of these. Military aircrafts, business jets and helicopters are using this

type of application.

Combustion

In the gas turbine turbojet engine the main function of the combustor is to ignite the

mixed compressed air with the fuel which results in a release of heat energy which is

expanded and accelerated at very high velocity, giving an even stream of heated gas at

all times when is demanded by the turbine.

According to the TTS (2011) the combustion chamber system is composed from the

following parts: perforated flame tube, outer air casing, a burner system and an igniter

plug. Various chamber layouts are found, however they function on the exact same

principle. These are: multiple can combustion chamber, turbo-annular combustion

chamber, reverse flow combustion system and annular combustion chamber. The

annular type is most commonly used in the commercial airline industry due to its

smaller size, resulting in saving weight and production costs, consequently rising the

efficiency of the combustor as there is a lesser wall area to cool.

The velocity at which the kerosene fuel is able to burn into the combustion chamber is

about a few m/s. Although the real velocity at which the air flows inside the chamber

system it can reach up to 150 m/s, however it can be slowed down by the diffuser

which is divergent in shape and helps the formation of static pressure which results in

lower velocity to about 24 m/s. Thou this is still a quiet increased velocity for the fuel

to burn, the low axial velocity has been introduced into the chamber in order to

protect the flame being alight.

The combustion chamber system can be divided into zones, thus having a primary

zone where the fuel can reach up a temperature of 2000 Celsius degrees due to the

fuel burned with only a part of the air from the compressor entering the chamber,

which is extremely hot to enter the nozzle guide vanes of the turbine. This zone is also

known as the burning total. And the other zone of the combustor is the dilution zone,

where the surplus of airflow is entering the flame tube through the holes in the wall

and it dilutes with the hot gasses, thus cooling the chamber walls and bringing down

the temperature of the hot gasses to about 1000 Celsius degrees which makes it

possible for the nozzle guide vanes and the other neighbouring material parts to

withstand the heat energy.

Thin sheets of corrosion resistant metal that can withstand high temperature are used

in the manufacturing of combustors. Its walls are designed with a series of

strategically purposed slots or holes in order to allow the airflow to enter the chamber.


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In general the used material for the construction of the chamber is nickel based alloy.

This can also withstand creep failure due to gradients and fatigue due to vibrational

stresses.

Turbine

In the turbojet engine the turbine primary function is to drive the compressor that is

connected to the shaft. The torque process for the turbine is produced by the draw in

of the hot gas stream released from the combustor section and a series of stages,

where one stage consists of a stator nozzle guide vane and a disk with rotor blades.

The shaft is also attached to the disk by a bolted flange. The number of shafts found in

a gas turbine engine depends on the number of turbines. The turbine can be found in

different shapes and sizes such as Impulse turbines, reaction turbines,

impulse/reaction turbines and radial inflow turbines.

The impulse turbine converts the gas flow energy to the turbine ring by impact or

impulse. A decrease in pressure and temperature is resulted due to the convergent

nozzle which accelerates the gases when leaving the nozzles. These accelerated gasses

are pointed by the nozzle guide vanes onto the turbine blades. In reference to the TTS

(2011, p. 6-7) ‘the cross sectional flow area of the rotor is constant…The torque

produced will be the sum of the forces on all the blades times the effective disk

radius.’

The nozzle’s objective in the reaction turbine is to direct the flow of gasses at the

correct angle onto the turbine rotor blades. A relative constant pressure, temperature

and velocity are seen by the flow of gasses through the nozzle due to its constant flow

area. The cross sectional flow area, on the rotor is found smaller at the discharge than

at the rotor inlet.

The impulse/reaction turbine is used in the gas turbine engine for aircraft propulsion.

The creation of vortex flow is resulted due to the NGV that forms convergent ducts

which allows the gas flow to whirl. In order to ensure the continuous smooth stream

of gas flow through the exhaust from the compressor the turbine rotor blades twist,

thus having an impulse root and a reaction tip.

The radial inflow turbine type is usually used in the APU’s and supercharges on

piston engines. Its construction shape looks very similar to the centrifugal

compressor.

The nozzle guide vanes are mounted into the casing and have the tendency to enlarge

due to the heat. Its construction material is nickel alloys as this is a very good element

that can withstand high heat. They are also cooled by the bleed air coming from the

compressor. Confirming to TTS (2011) there are few cooling techniques, such as:

internal cooling by impact, film cooling, multi-pass cooling, transpiration cooling and

platform film cooling.

The main issue that the turbine disk faces is the fatigue cracking; however the use of

nickel alloying is used nowadays in the manufacturing of this, which has proved much


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more efficient than the past technology. Due to the high stresses, temperature and the

huge centrifugal forces exerted on the rotor blades, while in operation, they have to

resist corrosion, oxidation, fatigue and thermal shock. The turbine blade also

experiences a phenomenon called creep, which expresses over three stages: primary

creep, secondary creep and tertiary creep (TTS 2011). Usually when the third stage is

reached the blades are replaced. This creep occurs over time due to the stresses and

centrifugal forces onto the blades, making them to grow in length.

The most common used method of attaching the turbine blades to the disk is the fir

tree root method.

Exhaust

The function of the exhaust is to let go the gasses into the atmosphere in a laminar,

vortex free and axially orientated flow in order to produce an optimum thrust. Its

construction consists of the following parts: exhaust casing, exhaust duct and nozzle.

Support struts and the cone are assembled onto the exhaust casing which is fitted rear

the turbine casing. For a smoother gas flow the inner cone protects the rear face of the

turbine disk from the exhaust gasses. The exhaust duct also knows as the tail pipe or

jet pipe according to the position of the engine in the aircraft, it is variable in size.

This duct is very important as it can have other functions such as housing the thrust

reversers and act as a silencer.

However the nozzle shape for subsonic engines is convergent in shape, this together

with the area of the nozzle being the most important factors as they dictate the

efficiency with which thrust is being produced. This means that the convergent duct

converts the pressure and heat energy in the gases into kinetic energy, resulting in a

high velocity of the gasses when they leave the nozzle. Thus indicating that in a

subsonic convergent duct the pressure and temperature decrease while the velocity

increases. While in a supersonic convergent nozzle the effects of those are totally

opposite.

Conclusion

In this assignment the flow of air through a gas turbine turbojet engine has been

analyzed. The importance of its components have been described and discussed in

order to see and determine what actually regulates the air to flow in a smooth

streamline without causing turbulence so the engine work efficiently.

It is concluded that upon the great design of the ducts and passages in the engine as

well as the aerodynamic and energy requirements the engine efficiency depends on.

Any inefficiency occurring will result in turbulence of the air flow which creates

vibrations and therefore failure of any component inside the engine.


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References

Aircav, n.d. Compressor Section. [Online]

Available at: http://www.aircav.com/gencompr.html

[Accessed 3 December 2013].

EASA Part 66, TTS Integrated Training System (2011) Module 15: Licence Category

B1-Gas Turbine Engine

Documents

Gas Turbine Engine