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GAS TURBINE



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Acknowledgment

"It is not possible to prepare this project report without the assistance &

encouragement of other people. This one is certainly no exception."

On the very outset of this report, I would like to extend my sincere & heartfelt

obligation towards all the personages who have helped me in this work. Without

their active guidance, help, cooperation & encouragement, I would not have made

headway in the project. I am ineffably indebted to (supervisor name) for

conscientious guidance and encouragement to accomplish this assignment. I extend

my gratitude to (COLLEGE NAME) for giving me this opportunity. At last but not

least gratitude goes to all of my friends who directly or indirectly helped me to

complete this project report. Any omission in this brief acknowledgement does not

mean lack of gratitude.



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Declaration

This report was written by (student Name) a student in the (Department Name) at

(University Name). It has not been altered or corrected as a result of assessment and

it may contain errors and omissions. The views expressed in it together with any

recommendations are of the student.



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Table of Contents

Acknowledgment ................................................................................................................ 1

Declaration .......................................................................................................................... 2

List of figures ...................................................................................................................... 6

Chapter 1 ............................................................................................................................. 9

(Introduction) ...................................................................................................................... 9

1.1 Gas turbine definition ......................................................................................... 10

1.2 Gas turbine types ................................................................................................ 11

The gas turbine can be classified into different types, which listed below. .............. 11

1.2.1Turbojet ............................................................................................................. 11

1.2.2 Turbofan ........................................................................................................... 12

1.2.3 Turboprop ........................................................................................................ 13

1.2.4 Turbo-shaft ....................................................................................................... 13

1.2.5 Industrial gas turbines for power generation ................................................... 14

1.3. Parts of Gas Turbine .............................................................................................. 14

1.3.1 Inlet section ............................................................................................................. 15

1.3.2 Compressor ............................................................................................................. 16

1.3. 3 Combustion system ............................................................................................... 16

1.3.4 Turbine.................................................................................................................... 17

1.3.5 Exhaust system ....................................................................................................... 17

1.3.6 Exhaust diffuser ..................................................................................................... 18

1.4 Open-Cycle and Closed Cycle .............................................................................. 18

1.5 Gas Turbine Design Procedure ............................................................................... 19

Chapter Two...................................................................................................................... 21



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(Related studies)................................................................................................................ 21

2.1 Introduction ............................................................................................................. 22

2.2 literature review ...................................................................................................... 22

2.2.1 Engine cycle ..................................................................................................... 23

2.2.2 Engine modification ......................................................................................... 25

2.2.3 Gas Turbine Emission ...................................................................................... 29

2.2.4 Combined Cycle............................................................................................... 30

2.2.5 Gas Turbine Troubleshooting Loads and Failure Modes .................................... 32

2.2.6 Gas Turbine Maintenance ................................................................................ 34

2.2.7 Gas Turbine Materials and Blades ................................................................... 35

Chapter Three.................................................................................................................... 39

(Theory of Gas Turbines) .................................................................................................. 39

3.1 Introduction ............................................................................................................. 40

3.2 Operation Principle of the Main Components of the Gas Turbine ......................... 40

3.2.1 Compressor ...................................................................................................... 40

3.2.2 Combustor ........................................................................................................ 44

3.2.3 Turbine ............................................................................................................. 48

3.2.4 Nozzle and Diffuser ......................................................................................... 49

3.3 Gas Turbine and Brayton cycle............................................................................... 51

3.4 payload calculations ............................................................................................... 51

3.5 engine noise ........................................................................................................... 51

CHAPTER FOUR ............................................................................................................. 56

(Engine specifications Selection) ...................................................................................... 56

4.1Engine selection requirements ................................................................................. 56

4.1.1 Engine Requirement......................................................................................... 57



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4.1.2 Performance parameters................................................................................... 57

4.1.3 Select engine component and materials. .......................................................... 58

4.1.4 Gas turbine limitations ..................................................................................... 63

4.1.4 Performance theoretical analysis ..................................................................... 64

Chapter Five ...................................................................................................................... 65

(Theoretical cycle analysis) .............................................................................................. 65

5.1 Introduction ............................................................................................................. 66

5.2 The gas turbine cycle governing equations ............................................................. 67

5.3Turbojet engine performance calculations ............................................................... 70

Flow velocity and mass flow rate .......................................................................... 70

Compressor ............................................................................................................ 70

Combustion chamber ............................................................................................. 71

Turbine ................................................................................................................... 71

Nozzle .................................................................................................................... 72

Thermal efficiency: ................................................................................................ 72

Thrust calculations ................................................................................................. 72

4.3.1 Discussion of calculations................................................................................ 73

4.4 Discussion of results ............................................................................................... 75

4.4 Performance analysis using Matlab.................................................................... 81

Conclusion ........................................................................................................................ 85

References ......................................................................................................................... 87

Appendix ........................................................................................................................... 91



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List of figures

Figure 1 energy transformation ......................................................................................... 10

Figure 2 the turbojet engine. Available at ......................................................................... 11

Figure 3 Turbofan gas turbine........................................................................................... 12

Figure 4Turboprop engine ................................................................................................ 13

Figure 5 Turbo-shaft layout, Available at ........................................................................ 14

Figure 6 Gas Turbine Parts. Available at ......................................................................... 15

Figure 7 Compressor Parts. (ACATERPILLAR Company, 2010) .................................. 16

Figure 8 Combustion System. (ACATERPILLAR Company, 2010)............................... 17

Figure 9 Open cycle diagram. (Roco et al., 1997 ............................................................. 18

Figure 10 Closed cycle diagram. (Roco et al., 1997) ........................................................ 19

Figure 11 Processes in P, V diagram. (Soares, 2007) ...................................................... 24

Figure 12 Process in h, S diagram (Soares, 2007) ............................................................ 24

Figure 13 Schematic diagram for heat exchanger. (Langston& Opdyke, 1997) .............. 26

Figure 14 The enthalpy-entropy diagram for intercooling. (Soares, 2007) ...................... 27

Figure 15 Schematic diagram for intercooling processes. (Langston& Opdyke, 1997) ... 27

Figure 16 The enthalpy-entropy diagram for reheating. (Soares, 2007)........................... 28

Figure 17 Schematic diagram for reheating processes. (Langston& Opdyke, 1997). ...... 29

Figure 18: Schematic diagram of the three processes that introduced to gas turbine.

(Brooks, 2000) .................................................................................................................. 31

Figure 19: T-S diagram modification on Brayton cycle. (Brooks, 2000) ......................... 32

Figure 20: Overall definition. (Sorokes et.al , 2006) ....................................................... 34

Figure 21 Temperature and pressure profile in gas turbine. (Carlos, 2007) ..................... 35

Figure 22 Creep rupture lives of alloy A, B and C plotted against the Solution Index

value. ................................................................................................................................. 37

Figure 23 : compressor classification (Mattingly and Ohain, 1996) ................................ 41

Figure 24 operating cost of the compressor (Mattingly and Ohain, 1996) ....................... 43



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Figure 25: Can type (Lefebvre, 1999). .............................................................................. 44

Figure 26 Cannular type (Lefebvre, 1999). ...................................................................... 45

Figure 27 Annular type (Lefebvre, 1999). ........................................................................ 45

Figure 28 burner component (Lefebvre, 1999). ................................................................ 47

Figure 29 Nozzle (Jiang, 1997) ......................................................................................... 50

Figure 30: Diffuser (Jiang, 1997)...................................................................................... 50

Figure 31: Open cycle (Tanaka et al., 2007) ..................................................................... 52

Figure 32: Closed cycle (Tanaka et.al, 2007) ................................................................... 52

Figure 33 T-S diagram (Tanaka et.al, 2007) ..................................................................... 53

Figure 34 major engine components ................................................................................. 58

Figure 35 The axial flow compressor ............................................................................... 59

Figure 36 combustion chamber ......................................................................................... 60

Figure 37 A twin turbine and shaft arrangement. ............................................................. 62

Figure 38 exhaust system .................................................................................................. 63

Figure 39 the cycle block diagram (DUNN, 2005). .......................................................... 66

Figure 40 The schematic diagram for a simple gas turbine (DUNN, 2005). .................... 67

Figure 41 T-s and P-v diagrams LANE, 2001) ................................................................. 68

Figure 42 The schematic diagram for a Atmospheric Temperature Vs. Altitude ............. 75

Figure 43 Atmospheric Pressure Vs. Altitude .................................................................. 76

Figure 44: Combustion chamber inlet Temperature Vs. Altitude ..................................... 76

Figure 45: Combustion chamber inlet Pressure Vs. Altitude ........................................... 77

Figure 46: Heat addition Vs. Altitude ............................................................................... 78

Figure 47: Heat rejection Vs. Altitude .............................................................................. 78

Figure 48: Power in Vs. Altitude ...................................................................................... 79

Figure 49: Power in Vs. Altitude ...................................................................................... 80

Figure 50 Matlab code ...................................................................................................... 81

Figure 51 the turbine work Vs. the compressor ................................................................ 82

Figure 52 the Wt Vs. ( T2 & T4) ...................................................................................... 83

Figure 53 the network plot Vs.( wt and Wc ) ................................................................... 83

Figure 54 T4 Vs. T2 .......................................................................................................... 84

http://c/Users/bradfordkw6/Desktop/%D9%85%D8%AD%D9%85%D8%AF%20%D8%B7%D9%8A%D8%A7%D8%B1%D9%87.docx%23_Toc354470410








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Abstract

Gas turbine engines are an internal combustion engine that provides mechanical power

through using gas as the working fluid. The definition, types, components, operation as

well as performance analysis for this engine was being the main aims that were discussed

in this project. To simply understand the basic concepts associated with this engine, the

project was divided into five chapters, with each of them discussed a specific topic.

The selection for the engine type to be used depends on the type of application at which it

used. In addition to that, there are other requirements that should be taken into account in

selecting the engine which are: the selection of the engine parameters, components,

materials and performance analysis. The performance analyses were developed to select

the engine. It was evaluated mathematically and then by using MATLAB software. The

analysis of the cycle was obtained at different altitude where the power required to

operate the engine is reduced with increasing the height. Also the performance of the

engine was analyzed related to the efficiency of the cycle where the values of the

efficiency were taken from 0.6-1 using MATLAB.

The selected engine was a turbojet engine. The principle of work for the turbojet engine

is to extract the fuel chemical energy and convert it to mechanical work through using the

gaseous energy of gas to drive both the engine and the propeller, which is then propelling

the airplane. The operation of this engine can be summarized as “the gas flows

continuously and entered a compressor at where it will be compressed and then heated in

the combustion chamber, and lastly, the heated gas flows through the turbine which

converts the gas energy into mechanical work.

From the theoretical analysis, this engine has an efficiency of 66 %, this means that it

produced a high power due to the direct correlation between output power and engine

efficiency. As a result, when less fuel is consumed through the engine and when the

output power is more, the engine will be more efficient.



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Chapter 1

(Introduction)



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1.1 Gas turbine definition

Gas turbine is a mechanical machine transformed energy from chemical to thermal and

then to mechanical as shown in Figure.1. In gas turbines is a constant flow of the liquid

agent. At first this fluid is compressed in the compressor then heated in the combustion

chamber. Finally, it passes through the turbine to convert the stored energy in gas to

mechanical work (Kulikov & Thompson, 2004).

Figure 1 energy transformation

Gas turbines have great roles in the electricity generation during the past years. Because

they have low investment cost, the gas turbines can be implemented to supply the power

for the transmission lines, and also they have considerable roles during the emergency

statues. Different researches have done focuses on the estimation of the performance of

gas turbines; these studies are divided into two main classes. The first one involves

studies have focused on the modeling of gas turbine cycle’s parts, the second type of

researches have focused on the most important parameters effecting on the design

performance (Lane. D, 2010).

chemical

energy

thermal

energy

mechanical

energy



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1.2 Gas turbine types

The gas turbine can be classified into different types, which listed below.

1.2.1Turbojet

Turbojet is an earliest and simplest kind of gas turbine; this type is usually used within

high speed aircraft Figure.2 shows the main parts of turbojet engine

Figure 2 the turbojet engine. Available at

http://wings.avkids.com/Book/Propulsion/advanced/types-01.html (accessed in 24 Jan 2012)

These turbines have different advantages such as- high jet velocity and small frontal area,

also the turbojet extracts the energy from the gas stream in order to drive the compressor

(Liu, F., 2001).

http://wings.avkids.com/Book/Propulsion/advanced/types-01.html

http://wings.avkids.com/Book/Propulsion/advanced/types-01.html



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1.2.2 Turbofan

This type is the widest type used for the aircraft propulsion Figure.3 shows the

main parts of this type. The air enters the engine and then is compressed and passedto the combustion chamber, part of air in this type is compressed to the lesser

extent, and bypasses the combustion section as shown in Figure .3, this method will

help to cool the thrust also the bypass air will rejoin the hot gases downstream of the

turbofan. In these cases, the overall velocity of the jet will be reduced in order get

lower noise levels, better efficiency and improved ( SFC) Specific Fuel

Consumption.( Roux,E., 2007).

Figure 3 Turbofan gas turbine.

Available at : http://www.techberth.com/high-bypass-ratio-turbofan-engine-components/(accessed at 24

Jan 2012)

http://www.techberth.com/high-bypass-ratio-turbofan-engine-components/

http://www.techberth.com/high-bypass-ratio-turbofan-engine-components/



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Figure 5 Turbo-shaft layout, Available at

http://www.britannica.com/EBchecked/media/19425/Turboshaft-engine-driving-a-helicopter-rotor-as-propulsor

accessed in 25 Jan 2012)

1.2.5 Industrial gas turbines for power generation

Industrial gas turbine has differ construction than other types where the size is various

from small to enormous size. The efficiency of this type can achieve 60% when remains

energy from the gas turbine is used again by a heat recovery to power in steam turbine

during the combined cycle configuration. This type is used generally to produce

electricity the following section will show the parts of this type in details:

1.3. Parts of Gas Turbine

Gas turbine consist of four main parts (compressor, injector, turbine and exhaust system)

figure 6 show them and other component.





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1.3.2 Compressor

Compressors are responsible for getting all the need from gas turbine. In addition to the

amount of compressed air must be fixed and high. Compression ratio ranging from about

9.5:1.

Compressor consists of several stages at each stage gradually increase the pressure for the

previous phase. Each stage contains of a multi of rotors which connected to a rotating

disk, which is followed by pro vanes connected to a fixed ring. And it also includes two

guides’ vanes the inlet and outlet. However vanes, is installed at the inlet and outlet

compressor. As shown in Figure 7

Figure 7 Compressor Parts. (ACATERPILLAR Company, 2010)

1.3. 3 Combustion system

It also known as Burner which provides heat to hot and pressurized gas, the main part in

the combustion system is Injector. The large amounts of fuel and air are difficult to

control. Which must accomplished with minimal loss pressure and the maximum heat

release. As shown in figure 8. (ACATERPILLAR Company, 2010)



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Figure 8 Combustion System. (ACATERPILLAR Company, 2010)

:

1.3.4 Turbine

Covert the high pressure and temperature to mechanical energy using a set of stationary

vanes connected by a set of rotating blades that worked under the reverse of thecompressor. The velocity of gas turbine increased related to the temperature and pressure

which increases the power can be extracted from the turbine.

1.3.5 Exhaust system

After the gas passes the turbine, then discharged to the exhaust the thermal energy is

extracted from the gas to be mechanical energy , the remains energy in gases can be used

again to enhance the performance of the gas turbine where these gases can be used in

regenerative unit. (Kulikov& Thompson, 2004)



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Figure 10 Closed cycle diagram. (Roco et al., 1997)

1.5 Gas Turbine Design Procedure

Different researches have done focuses on the estimation of the performance of gas

turbines; these studies are divided into two main classes. The first one involves studies

have focused on the modeling of gas turbine cycle’s parts, the second type of researches

have focused on the most important parameters effecting on the design performance.

The most important factors will be studied in this project are: the cycle efficiency

(turbine, compressor), the power output, and the temperature and pressure at each state.

Advancements in materials used for manufacturing gas turbine which operate on high

temperatures and pressures have contributed during the enhancement of the temperature

capabilities and efficiencies of the gas turbines.. Finite element analysis involves of a

computer model of a material of any mechanical design that is subjected under stress in

order to get specific results.



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To design a gas turbine there are many steps are mention below:

1- Calculate the needed power for gas turbine.

2-Take the geographical factors in deigns account.

3- Select the operation cycle of gas turbine.

4- Make all calculation to find the variables of all point in the operation cycle.

5- Select the optimum part according to design calculation. (Strand, 2006).



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Chapter Two

(Related studies)



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2.1 Introduction

This chapter will discuss a literature review for the previous works related to the gas

turbine. the literature includes a discussion for the ideal operation cycle of the gas

turbine (Brayton), discuss its calculation like (performing, efficiency and pressure ratio)

and all possible modification to enhance it efficiency which can be applied on the

Brayton cycle such as (heat exchanger, inter cooling, reheat and combination between

them).

2.2 literature review

According to (TSA, 2004) research about design and performance requirement for a

“Gas-Turbine Engine from an Automobile Turbocharger”, and as a result defining the gas

turbine as systems produce a positive work transfer by using air and fuel so it considered

as thermodynamic systems converting the chemical energy in fuel to mechanical energy.

The gas turbine activates on an unlock cycle consisting of a combustor, a compressor,

and a turbine collective in series. Atmospheric air goes through the compressor to be

compressed by “a negative shaft work transfer” to be combined and burned in the

combustion chamber with fuel. Both the specific volume of the air and the temperature

increases in the combustor. Then the hot air is fed into to be expanded in the turbine. The

air expansion generates “a positive shaft work transfer”, finally, exhausting the expanded

air to the atmosphere. A positive shaft work transfer is created since the negative shaft

work transfer created by the turbine need to operate the compressor is lower than the

positive work transfer.

(Roberts, 1990) was worked on one type of gas turbine, which is the turbofan engine. He

defined it as an “engine where the first stage compressor rotor is larger in diameter than

the rest of the engine” (Airplane…, n.d, pg.3). In the turbofan engine the air passesthrough the fan that is situated near the inner diameter the same air also passes through

the compressor stages situated in the core of the engine and then it is furthermore

processed and compressed through the engine cycle (Airplane…, n.d.pg.5). He said, to

design Turbo fan engine there is a need of wave rotor which is stated to be the partial

admission device that causes the one gas to compress another through the use of wave



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propagation. The development of these devices for Turbo fan engines were designed in

the early 1950s as stated by Roberts in his Thesis. Further on, there were few

developments which were stated to be attempted after the 1960’s. It is also stated that the

Wave rotor components have promising applications in small engines that range from

600 to 1000 in pounds thrust range. Such kind of engines are stated to have applications

which are used in cruise missile, piloted vehicles, helicopters and the small thrust

turbofans (Roberts, 1990).

2.2.1 Engine cycle

Soares, 2007) was studied the ideal cycle for gas turbine engine which is called aBrayton. This theoretical cycle for simple gas turbine consists of two isentropic and two

constant pressure processes. The compressor consists of four main parts:

Inlet duct

Compressor

Fuel injector

Exhaust diffuser.

In every inlet and outlet of the compressor parts Joule-Brayton define some variables

such as temperature (T), pressure (P), entropy (S), enthalpy (h) and volume (V) can be

used to study the compressor, as shown in figure 11 and 12.



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Figure 11 Processes in P, V diagram. (Soares, 2007)

Figure 12 Process in h, S diagram (Soares, 2007)

From the relationships have been drawn in figures (11 and 12), it was concluding that:

- During step 2------- 3 the input power to compressor given as:

)



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Reheating

The main idea of reheating is similar to intercooling but it applied in the turbine at point

4 which lead to increase the output work as shown in figure below.

Figure 16 The enthalpy-entropy diagram for reheating. (Soares, 2007)

The output work given as.

W output= W4−4.5 +W4.5−5 = C p (T4− T4.5) + C p (T4.5 − T5).

Figure 17 below show the schematic diagram for reheating processes.



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As a result of existing HRSG’s in the com pound cycle which makes it ideal for medium

load application. The HRSG’s same as SCR in the NOx control.

The emission percentage in the compound cycle depends on the hourly limitation and

Ton per Year (TPY) produced.

2.2.4 Combined Cycle

The thermal efficiency of a gas turbine can be increased by decreasing the input power or

increasing the output work or both.to reduce the input work the required ratio pressure

must be smallest as possible and the compression done in multi stage with intercooling.

To enhance the output work the required ratio pressure must be larger as possible, the

expansion process done in multi stage with reheating. (Brooks, 2000)

Two previous processes must do without raising the maximum temperature in the cycle

that leads to increase the number of stages of the cycle.

During the combustion process in the gas turbine a heat exchanger introduced between

the gases leave the turbine (high temperature) and the gases leaves the compressor (low

temperature) to increase the thermal efficiency of the cycle. Figure 18 shows the

schematic diagram of the three processes that introduced to gas turbine.



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Figure 18: Schematic diagram of the three processes that introduced to gas turbine. (Brooks, 2000)

Figure 19 show the modification occurs in the T-S diagram on the gas turbine.



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Figure 19: T-S diagram modification on Brayton cycle. (Brooks, 2000)

2.2.5 Gas Turbine Troubleshooting Loads and Failure Modes

The failure of gas turbine occurs due to many reasons like: loss of performance,

excessive noise, vibration, poor engine control, structural failure, excessive exhaust

emission and over load.

The structural failure of gas turbine can be defined as any change in the shape, size, or

mechanical properties lead to make the part unsatisfied to do it work.

On the other hand, can be define as any physical or chimerical change lead to make the

part unsatisfied to do it work. (Stelling & MNIMH, 2010)

Vibration is one of the most important issue must be consider due to it effect, as known

compressor and turbine blades are subjected to High Cycle Fatigue (HCF) its connected



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Actually for every 25000 to 50000 operation work hour gas turbine needs an overhaul

(dimensional inspection, testing and product upgrade) depending on load and type of

maintenance. (Energy Nexus Group, 2002)

2.2.7 Gas Turbine Materials and Blades

As a result of little operation hour of the gas turbine due to high pressure and temperature

which a result of the material used, unacceptable reliability and thermodynamic

efficiency, damaging to equipment and injures to people, all these reason make the

material used as an important factor in the design process of gas turbine.(Carlos, 2007)

Compression, combustion, generation power and exhaust processes occurs in gas turbine

the most dangerous point located between the combustion chamber outlet and the turbine

inlet, it is the most sensible and challenging design point. Look at figure 21.

Figure 21 Temperature and pressure profile in gas turbine. (Carlos, 2007)



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3.1 Introduction

This chapter will discuss the operation principle of the main parts of the gas turbine

system (compressor, combustor, turbine and nozzle and diffuser), and the possible failure

modes for each component in the gas turbine system. Then discussion will be carried for

how to use the gas turbine in distillation process and what is the possibility and

affectivity of the distillation in the gas turbine power plant. Moreover, the efficiency of

the gas turbine will be discussed according to the Brayton cycle; finally the main

differences between the steam turbine and the gas turbine will be mentioned.

3.2 Operation Principle of the Main Components of the Gas Turbine

Gas turbine is one of the engineering applications which have a wide range uses in the

industrial applications; the gas turbine consists of four main components which are:

1- Compressor

2- Combustor

3- Turbine

4- Nozzle and Diffuser (Mattingly and Ohain, 1996)

3.2.1 Compressor

Compressor in different applications is used to improve the overall efficiency of the

system. The compressor in general consists of many parts, which are (Mattingly and

Ohain, 1996):

Intake Air Filters: this part prevent any dust from entering into the compressor

Cooler stage: in this stage the temperature of the inlet air will be reduced to

decrease the compression ratio and enhance the compressor efficiency.

Impeller: it makes the compression in the stage of the compressor.



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COMPRESSORS TYPES

Figure 23 shows the main classification of the compressor based on the principle of

operation and type of forces.

Figure 23 : compressor classification (Mattingly and Ohain, 1996)

The actual volumetric flow rate of the compressor can be calculated form the following

equation:

Where



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Figure 24 operating cost of the compressor (Mattingly and Ohain, 1996)

There are many problems occur in the gas turbines which are:

1- Fouling: is one of the most important problems which can be presented as a result from

the particle in the intake air (smoke, carbon, sea salt and the oil mists) or from the

annulus surface, these particles have a range from 2 to 10 μm which can cause a

change in the airfoil shape, so it is important to make sure there are a good filter in the

first stage in the compressor to avoid this problem and must be optimum for particle

remove, cost, service life and the pressure losses (Mattingly and Ohain, 1996).

2- Hot corrosion: this can be defined as a chemical reaction between a specific material

and the compressor components; this reaction can make a big change in the flow path,

crack or damage in the impeller. The oxidization process become more active under a

high temperature, so the compressor component must be protected from the oxidization

process by coating the component (Mattingly and Ohain, 1996).



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2- Cannular type which shown in figure 26

Figure 26 Cannular type (Lefebvre, 1999).

3- Annular type which shown in figure 27

Figure 27 Annular type (Lefebvre, 1999).



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This process is used in the application which needs a high pressure ratio but in this

method the specification of the fluid must be taken in the considerations and additional

requirements to reach the wanted pressure.

c- Pre-evaporating

In this method the fuel evaporates before enter the combustion zone (Lefebvre, 1999).

d- Vaporizing

The operation principle of this method is same as the operational principle of the air blast

but in this method the surrounding heat is used to improve the evaporation process.

7- Igniter

It can be defined as a device which gives the spark (to start the combustion) in the

combustion zone, the electrical spark is usually used in different applications but one of

the other important types is the oxygen injection which makes the combustion process

easier.

The burner and its main component can be shown in figure 28.

Figure 28 burner component (Lefebvre, 1999).



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The process of the closed cycle type in the Brayton can be drawn in the Temperature-

Entropy relationship as shown in figure 33.

Figure 33 T-S diagram (Tanaka et.al, 2007)

From figure 33, it can be shown that the compressor and turbine work at isentropic

process, heat addition and rejection work at constant pressure.



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3.5 Aircraft Noisy

Aircraft noise represents a type of pollution called noise pollution, which is generated

mainly due to two reasons; the first is the intake, exhaust and other major engine

components and the aerodynamic drag of air flow around the aircraft body and wings.

Commonly, during the take-offs the noise is generated by the engine. But the airframe

noise is the most important factor during landing, where the engine operates at low

power. The developments and improvement in the manufacturing of the aircraft are

concentrated on enhancing engine and airframe design with an n important aim of

reducing the noise that affects on the aircraft. (CAIRNS PORT AUTHORITY, 2006).

Researchers reach that dying by a heart attack was more public among people with

increased experience to aircraft noise. Those effects was specifically obvious for people

who were exposed to really high levels of noise, and was dependent on how long those

people had lived in the noisy place ( Matthias Egger) . But this study could support to

determine whether the sound is really the main effect, or if it is something else grouped

with the noise, such as air pollution. It's been a problem that when you study the case of

road traffic noise there are both high levels of noise and high level of air pollution. After

taking consideration for the air pollution and other factors including education and

revenue levels, the group found that both the level and duration of aircraft noise increase

the risk of a fatal heart attack (Alleyne, 2010).

Therefore, in the design the turbojet engine is selected since it has the least noise also has

a good performance.



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CHAPTER FOUR

(Engine specifications Selection)

4.1Engine selection requirements

The gas turbine engine is considered to be a complex machine including operation at

extremes of pressure and temperature and demanding expertise at the highest level of

engineering technology. To select the best engine to use it in any application, some

engine selection requirements should be considered as shown in the following.



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run successfully at speeds value below the design condition. As increasing the

pressure ratio, the integration of variable stator vanes makes sure that the airflow

is moved toward the successive stage of “rotor blades” at an suitable angle. Figure

below demonstrates the the axial flow compressor (Rolls-Royce plc company.

2001)

Figure 35 The axial flow compressor

Compressor materials

To select the best material for compressor design, the material should attain the

most cost effective design. For casing designs it requires to build with a light rigid

construction which allows blade tip permissions to be maintained accurately with

a high efficiency. It is generally made from aluminum at the front of the

compression system and at the final stages of the compression system is made

from nickel based alloys (Rolls-Royce plc company. 2001).

b. Combustion chambers



Page | 62

Figure 37 A twin turbine and shaft arrangement.

Turbine MATERIALS

Basically, turbine parts consist of combustion discharge nozzles, the nozzle guide vanes,

turbine discs and the turbine blades.

The Nozzle guide vanes are made from nickel alloy which has a good heat resistance and

Ceramic coatings are used to improve the heat resisting properties and decrease the

quantity of cooling air needed, and thus enhancing the engine efficiency. As well as the

turbine blades are made from Ceramic (Rolls-Royce plc company. 2001).

d. Exhaust system

The main function of the exhaust system of the gas turbine is to pass the discharge gases

of the turbine to atmosphere in the wanted direction to give the resultant thrust. In turbo-

jet engine the pressure and velocity of the exhaust gases generate thrust. The exhaust

system should be designs accurately due to that it has an considerable impact on the

engine performance. The jet pipe and outlet nozzle areas affect the inlet temperature,



Page | 65

Chapter Five

(Theoretical cycle analysis)



Page | 66

5.1 Introduction

The basic and ideal cycle is called the constant pressure cycle due to that the cooling and

heating processes are conducted at constant pressure. Figures (34 and 35) below

demonstrates the cycle block diagram and the cycle schematic diagram respectively. The

basic gas turbine cycle is called the Brayton cycle according to George Brayton (LANE,

2001).

Figure 39 the cycle block diagram (DUNN, 2005).



Page | 67

Figure 40 The schematic diagram for a simple gas turbine (DUNN, 2005).

5.2 The gas turbine cycle governing equations

There are four ideal processes in the cycle that are (DUNN, 2005):

- Process (1-2) is reversible adiabatic (isentropic) compression (in compressor)

requiring power input.

- Process (2-3) is constant pressure requiring heating addition.

- Process (3-4) is a reversible adiabatic (isentropic) expansion (in turbine)

producing power output.



Page | 68

- Process ( 4-1) is a constant pressure cooling requiring heat rejection.

Figure below demonstrates the T-s and P-V diagrams for the brayton cycle.

Figure 41 T-s and P-v diagrams LANE, 2001)

- Efficiency calculations

The thermal efficiency is calculated through applying “the first law of thermodynamic”

which gives:



Page | 70

5.3Turbojet engine performance calculations

The turbojet engine performance calculations are carried out for private passengers plan

that carrying 30 passenger. The calculation where made when the plane was at 35000 ft

height. At this elevation the initial conditions for temperature and pressure are given

according to (U.S Standard Atmosphere):

- T1= -65.61°C = 218.8K

- P1= 23. 8 KPa

Flow velocity and mass flow rate

Where, R is the gas constant.

√

Assume, A= 0.05 = 0.00464515 based on literature of (CARETTO, 2008), then

Compressor

According to the jet engine literature (*) the Compressors can achieve compression ratios

in excess of 40:1, then assume it to be 45

()

http://www.engineeringtoolbox.com/standard-atmosphere-d_604.html




Page | 72

Nozzle

Thermal efficiency:

Thrust calculations

For a turbojet the thrust equation is given by the general thrust equation (F) with the

pressure-area term set to zero (MARZOCCA, 2011),

To determine the exit velocity, the Mach number at exit should be evaluated as shown

below:

√

http://www.grc.nasa.gov/WWW/K-12/airplane/thrsteq.html

http://www.grc.nasa.gov/WWW/K-12/airplane/thrsteq.html



Page | 73

According the conservation of mass principle then:

4.3.1 Discussion of calculations

The calculations made for the turbojet engine performance in the case of a plane with 30

passengers at a height of 35000 ft. To begin the analysis the initial conditions of the

compressor inlet needed to be determined, and it is determined according to the (U.S

Standard Atmosphere), then the speed of sound is calculated to evaluate the amount of air

entered to the compressor, by knowing the pressure ratio for the compressor the outlet

temperature of the air at the compressor exit, which is increase because of the

compression process, finally the power needed for the compressor is calculated which is

not significant compared to the output power from the turbine .

In combustion chamber, the combustion process represent heat addition to complete the

cycle occurs under condition of constant pressure, and the amount of heat added is

evaluated. This heat added represent a cost must be added to the system, as this heat

added decreases the system will be more efficient and preferable by assuring that the

performance has to be.

In the turbine the expansion process occurs under reversible adiabatic (isentropic)

condition to produce the output power the outlet temperature of the air and the pressure

of the air reduced significantly because of power extraction. The output power is not

significant compared to the input.



Page | 74

The Nozzle is a device used to accelerate the fluid before it’s enter the compressor , while

the fluid velocity, it's temperature will be increase but it's pressure will decrease and this

represent a heat rejected in the turbojet system, and this is not a significant value.

Thrust is the force which moves an aircraft through the air; it is used to overcome the

drag of an airplane, and to overcome the weight of a rocket, thrust is generated by the

engines of the aircraft through some kind of propulsion system. The thrust is calculated

using a special equation that depends on the velocities at the inlet and exit of the engine.

Finally, the thermal efficiency is a measure of the performance of the device that uses

thermal energy and calculated by evaluate the net output power by subtract the input

power from the output power, and evaluate the heat added to the system. For this system

the thermal efficiency reaches up to 66 % which is a good percentage that demonstrate

the system performance to use the thermal energy.



Page | 76

.

Figure 43 Atmospheric Pressure Vs. Altitude

These changes effect on the properties of the inlet air for the engine, so these changes

effect on the compressor efficiency.

2. Pressure and temperature variation in combustion chamber inlet

The figure below shows the variation in the inlet temperature for the combustion chamber

with altitude, where the graph demonstrates that as the altitude increases the inlet

temperature for the combustion chamber decrease. Therefore, the power input to the

compressor will also decrease.

Figure 44: Combustion chamber inlet Temperature Vs. Altitude

0

10

20

30

4050

60

70

80

0 10000 20000 30000 40000

A t m .

P r e s s u r e

( K p a )

Altitude (ft)

Atmospheric Pressure Vs. Altitude

P1(kpa)

0

200

400

600

800

1000

0 10000 20000 30000 40000

I n l e t c o m b u s t i o n c h a m b e r t e m p e r a t u r e

Altitude (ft)

Combustion chamber inlet Temperature Vs.

Altitude

T2



Page | 78

Figure 46: Heat addition Vs. Altitude

Figure 47: Heat rejection Vs. Altitude

0

100

200

300

400

500

0 10000 20000 30000 40000

Q i n ( K w

)

Altitude (ft)

Heat addition Vs. Altitude

Q_in

0

50

100

150

200

0 10000 20000 30000 40000

Q o u t ( K w )

Altitude (ft)

Heat rejection Vs. Altitude

Qout



Page | 79

4. Power variation with altitude

The figure below shows the variation in the power input to the compressor with altitude,

where the graph demonstrates that as the altitude increases the input power decreases

linearly. Therefore, this will enhance the performance of the turbo jet since it consumes a

less amount of power, and then the efficiency will also enhanced.

Figure 48: Power in Vs. Altitude

The figure below shows the variation in the power output from the turbine with altitude,

where the graph demonstrates that as the altitude increases the output power decreases in

a non- linear shape. Therefore, this will affect in the efficiency of the engine and as a

result the performance of the turbojet. Therefore, the plane cannot reach to high altitude

because it is clearly that the performance of the plane affected by the height, and it may

represent a danger and risk.

0

2

4

6

8

10

0 10000 20000 30000 40000

P i n ( K w )

Altitude (ft)

Power in Vs. Altitude

Power in



Page | 80

Figure 49: Power in Vs. Altitude

0

50

100

150

200

250

300

350

0 10000 20000 30000 40000

P o u t ( K w )

Altitude (ft)

Power out Vs Altitude

power out



Page | 82

The work of the turbine Vs. the compressor work was plotted as shown in the following

figure as shown when the efficiency of the engine increase the output work will increase .

Now the turbine was plotted as function of the inlet temperature of the combustion

chamber and the inlet of the turbine, where the variation was measured at different

efficiencies as shown in the following Figure.

Figure 51 the turbine work Vs. the compressor



Page | 83

Figure 52 the Wt Vs. ( T2 & T4)

Now again the net power was plotted as function of the compressor work and the turbine

work where the variation on these values also was measured depending the efficiency .

Figure 53 the network plot Vs.( wt and Wc )



Page | 85

Conclusion

Gas turbine is a mechanical machine transformed energy from chemical to thermal and

then to mechanical. In this project the performance of the gas turbine was discussed in

order to find a suitable design for the engine at required specifications. According to this

specification the best and simple engine selected was the turbojet engine. for this engine,

a performance calculations were carried out for private passengers plan that carrying 30

passenger. The calculation where made when the plane was at 35000 ft height. At this

elevation the initial conditions for temperature and pressure are given according to (U.S

Standard Atmosphere). The amount of heat added for the system in the combustion

chamber will decrease with increasing the height. At 33 000 ft the minimum heat additionwhere the amount of heat rejection from the system to the surrounding also will decrease

with increasing the altitude. MATALAB software was used to get the performance of the

proposed engine. According to the performance analysis the engine selection can be

achieved. In addition to the performance analysis, the selection of engine also includes

the selection of best engine parameters, outside condition as well as the engine parts

selection with best materials. All parts of the engine were selected depending the

calculations in chapter five. The results of performance analysis were graphically shown.

From these figures the performance of engine was demonstrated at different altitudes, and

the results are shown as:

- The inlet temperature and pressures to the compressor decreasing with increasing

the altitude. These changes effect on the properties of the inlet air for the engine,

so these changes effect on the compressor efficiency.

- The inlet pressure and temperature to the combustion chamber varies with

altitude, as the altitude increases the inlet pressure and temperature decreases

significantly.

- As the altitude increases the amount of heat added in the combustion chamber

decrease.

- As the altitude increases the input power decreases linearly.







Page | 86

- As the altitude increases the output power decreases in a non- linear shape.

Therefore, this will affect in the efficiency of the engine and as a result the

performance of the turbojet.



Page | 87

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Page | 91

Appendix MATLAB calculations

T2 =

519.4413 616.8736 640.1044 661.5823

495.0589 581.6654 602.3151 621.4065

475.5530 553.4989 572.0835 589.2659

T4 =

1.0e+003 *

1.0570 1.2672 1.5488 1.8219

1.0016 1.1756 1.4299 1.6746

0.9463 1.0840 1.3110 1.5273

T2 =

519.4413 616.8736 640.1044 661.5823

495.0589 581.6654 602.3151 621.4065

475.5530 553.4989 572.0835 589.2659



Page | 93

0.9463 1.0840 1.3110 1.5273

wc =

264.2159 381.5282 409.4991 435.3593

234.8586 339.1362 363.9992 386.9860

211.3727 305.2226 327.5993 348.2874

wt =

1.0e+003 *

0.5158 0.8533 1.1076 1.3718

0.5803 0.9599 1.2460 1.5433

0.6448 1.0666 1.3845 1.7148

wnet =



Page | 94

1.0e+003 *

0.2516 0.4717 0.6981 0.9365

0.3454 0.6208 0.8820 1.1563

0.4334 0.7614 1.0569 1.3665

r_bw =

0.5122 0.4471 0.3697 0.3174

0.4047 0.3533 0.2921 0.2508

0.3278 0.2862 0.2366 0.2031

q_reg =

1.0e+003 *

0.6773 0.8194 1.1450 1.4619

0.6383 0.7484 1.0428 1.3270

0.5931 0.6684 0.9310 1.1820



Page | 95

qin =

1.0e+003 *

1.3954 1.1170 0.7589 0.4118

1.4686 1.2373 0.9140 0.6030

1.5411 1.3567 1.0681 0.7931

T2 =

519.4413 616.8736 640.1044 661.5823

495.0589 581.6654 602.3151 621.4065

475.5530 553.4989 572.0835 589.2659

T4 =

1.0e+003 *

1.0570 1.2672 1.5488 1.8219

1.0016 1.1756 1.4299 1.6746

0.9463 1.0840 1.3110 1.5273



Page | 96

T2 =

519.4413 616.8736 640.1044 661.5823

495.0589 581.6654 602.3151 621.4065

475.5530 553.4989 572.0835 589.2659

T4 =

1.0e+003 *

1.0570 1.2672 1.5488 1.8219

1.0016 1.1756 1.4299 1.6746

0.9463 1.0840 1.3110 1.5273

T2 =

519.4413 616.8736 640.1044 661.5823

495.0589 581.6654 602.3151 621.4065

475.5530 553.4989 572.0835 589.2659



Page | 97

T4 =

1.0e+003 *

1.0570 1.2672 1.5488 1.8219

1.0016 1.1756 1.4299 1.6746

0.9463 1.0840 1.3110 1.5273

wc =

264.2159 381.5282 409.4991 435.3593

234.8586 339.1362 363.9992 386.9860

211.3727 305.2226 327.5993 348.2874

wt =

1.0e+003 *



Page | 98

0.5158 0.8533 1.1076 1.3718

0.5803 0.9599 1.2460 1.5433

0.6448 1.0666 1.3845 1.7148

wnet =

1.0e+003 *

0.2516 0.4717 0.6981 0.9365

0.3454 0.6208 0.8820 1.1563

0.4334 0.7614 1.0569 1.3665

r_bw =

0.5122 0.4471 0.3697 0.3174

0.4047 0.3533 0.2921 0.2508

0.3278 0.2862 0.2366 0.2031



q_reg =

1.0e+003 *

0.6773 0.8194 1.1450 1.4619

0.6383 0.7484 1.0428 1.3270

0.5931 0.6684 0.9310 1.1820

qin =

1.0e+003 *

1.3954 1.1170 0.7589 0.4118

1.4686 1.2373 0.9140 0.6030

1.5411 1.3567 1.0681 0.7931

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