Propulsion System Analysis_BhatU

8/3/2019 Propulsion System Analysis_BhatU

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ABSTRACT

This report details the analysis and design for a turbojet based gas turbine

engine to drive an electrical generator of an Auxiliary Power Unit (APU)

which can be used for a long-haul transonic aircraft. The report is composed ofthree parts. Part 1 details the analysis and design for the programming code, to

be used in MATLAB, in order to automate the calculation required for the

project. Part 2 details the analysis of the outputs using the program developed

in Part 1. This analysis is used for the justification of the selected valued of the

different design parameters required for the engine. Part 3 details the report

about the material, and component selection for the project.


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EXECUTIVE SUMMARY

This report details the analysis and design for a turbojet based gas turbine

engine to drive an electrical generator of an Auxiliary Power Unit (APU)

which can be used for a long-haul transonic aircraft. The report is composed ofthree parts.

Part 1 details the analysis and design for the programming code, to be used in

MATLAB, in order to automate the calculation required for the project.

Part 2 details the analysis of the outputs using the program developed in Part

1. This analysis is used for the justification of the selected valued of the

different design parameters required for the engine.

Part 3 details the report about the material, and component selection for the

project.


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LIST OF CONTENTS

ABSTRACT 1

EXECUTIVE SUMMARY 2

LIST OF CONTENTS 3

1 INTRODUCTION

2 PART 1:

3 PART 2:

4 PART 3:

5 CONCLUSION

REFERENCES


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1.INTRODUCTION

1.1BackgroundThe aim of this report is to give an optimal design for a turbojet based turboshaft engine which can be used to generate electricity as an APU in an aircraft.

The design requires selection of values for the different parameters, materials

and components for the engine. The different requirements for the engine

design are described in the next section.

The APU supplies electricity to the aircraft only when sufficient electricity is

not produced by the main engines. The APU does not produce any trust. This

means that the APU is only utilized intermittently. This is an important factor

in the design as it sets some limits over the cost and size for the project.

Compressor Turbine

CombustionChamber

Air

ElectricalGenerator

Fuel

ElectricalPower

T2, P2

T3, P3 T4, P4

T5, P5

ExhaustGas

Wg

APU

Fig 1.1 APU

A gas turbine (in this case a turbojet / turbo shaft) engine follows a Brayton

(Joule) cycle. Fig 1.1 shows the basic structure of the APU. Some of the

parameters have already been set which are shown in the next section.

1.2 Requirements


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The design needs to follow some constraints. These have already been given

and are fixed. These parameters form the base around which whole design will

be developed. These given constraints are as follows:

Inlet dry air properties

R 287 J/kg/KCp 1500J/kg/K

1.4

Compressor properties

Compressor inlet pressure 1 bar

Compressor inlet temperature 15o

C

Compressor isentropic efficiency 93%

Bleed flow from exit of compressor 1.5%

Combustion chamber properties

Combustion Pressure Loss 1%

Combustion Fuel to Air Ratio 0.025:1Turbine properties

Turbine exit pressure 1.2

Turbine isentropic efficiency 95%

Generator properties

Electrical generator efficiency 95%

Aircraft Engine power demand (Wg) 400Kw

Some of the other design requirements that have been taken into consideration

for the design of the engine based on the role of the APU as stated in section

1.1 are:

The engine needs to be cheap. The size of the engine should be as small The weightof the engine should be small The engine should have low maintenance and running cost

The other requirements at which the design has to aim at are like material that

is available, types of components that are available etc. Most of these depend

on the optimal values of compressor pressure ratio and the turbine inlet

temperature.

1.3 Design methodology and structure

The whole design process has been has been divided in to three parts.

As there are lots of calculations to be done in order to show the range of

possible values that can be chosen a MATLAB script file needs to be develop.

This will enable to automate this calculation. Part 1 details the design and

coding needed for this purpose. The output of Part 1 is in the form of different

plots.

Part 2 looks into the analysis of the plots produced from Part 1 in order to

make proper selection for the optimal design parameters.


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Part 3 looks into the selection of different components and materials which for

the optimal design.


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2.PART 1

2.1 Overview

This part deals with the analysis and design for the coding to be used in for theautomation of the thermodynamic calculation which would facilitate the next

stage of APU design. The coding is done using MATLAB. The program will

give the out put of the design in forms of different plots which are used in Part

2 of this report.

This section is divided into four parts Requirement Specification, Analysis,

Design, MATLAB code and Code testing.

2.2 Requirement specification

The requirements of this program are to use the predefined parameters andconstraints given for the engine design in different thermodynamic

calculations and produce the results in form of plots which can be analyzed at

a later stage.

The given constrains have been set into different variables. The MATLAB

code for this is saved in file psacon.m (See section 1.1 for the different given

constraints. )

2.3 Analysis

In order to give an over view of the program on the next page is the Program

Outline.


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Program Outline

INPUT

Given constraints Range for Pressure ratio Range for Turbine exit

temperature

PROCESS

Thermodynamic calculations for:o Temperatures T3, T4, T5o Mass flow rate of airo Mass flow rate of fuelo SFC

FILES

No filesOUTPUT

Plots of different calculated values


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The different equations to be that are used in the program are:

2

3

P

Prcomp

21

3TrT compis

2

23

3T

TTT

comp

is

23PP comp

combustorPLPP 134

5

4

P

Prturb

1

4

5

turb

is

r

TT

turbisTTTT 5445

generator

gnet WW

2354

5

TTTTC

Wm

p

net

54mm

FAR

mm

1

4

3

bleedfmm /32

FARmmfuel 3

gWmsfc /


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2.3 Design

The flow chart below shows the design and working of the program:

Start

Calculate T3is &T3

Run psacon.mto load the

constraints intomemory

Calculate P3 & P4

Calculate Turbinepressure ratio

Calculate T5is &T5

Calculate massflow rates

Calculate SFC

Plot thecalculatedvalues

End

Load the chosenrange of

Compressorpressure ratio & T4

Fig 2.1 Program flow chart

All the calculation are done basis of the equations shown in section 2.3.


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2.4 Program code

The following section shows the program code for the configuration file

psacon.m

% PSACON%% (c) Ubaier Ahmad Bhat%

% PSACon or Propulsion System Analysis Constraints is part of Propulsion% System Analysis program.

clear% --------------------------------------------------------------------% Given Constraints

% Inlet Air propertiesR = 287; % J/kg/KCp = 1005; % J/kg/Kgamma = 1.4;

% Station 2p2 = 1; % Compressor inlet pressure in bars

p2 = p2 * 10^5; % in Pat2 = 15; % Comperssor inlet temp. in C

t2 = t2 + 273.15; % in Kn_comp = 0.93; % Compressor isentropic efficiency

% Station 3bfr_comp = 0.015; % Bleed flow rate at comp. exit

p3 = []; % Compressor exit pressure , NOT KNOWNt3 = []; % Compressor exit temp. , NOT KNOWN

% Combustion chamber (cc)pl_cc = 0.01; % Pressure loss at ccfar_cc = 0.025/1; % Fuel to Air ratio

% Station 4n_turb = 0.95; % Turbine isentropic efficiencyp4 = []; % Turbine inlet pressure , NOT KNOWNt4 = []; % Turbine exit pressure , NOT KNOWN

% Station 5p5 = 1.2; % Turbine exit pressure in bars

p5 = p5 * 10^5; % in Pat5 = []; % Turbine exit temp. , NOT KNOWN

% Electric generatorn_gen = 0.95; % Efficiency of generatorWg = 400 ; % Power demand in kW

Wg = Wg * 10^3; % in W% ---------------------------------------------------------------------


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Following is the code for the main program psa.m.

% PSA%% Propulsion System Analysis% (c) Copyright 2006, 2007 Ubaier Ahmad Bhat%% This program is part of Propulsion System Analysis coursework for module% 216SYS Aerospace Technology 1.

% Load the constraints from the configuration file

psacon

% Chosen valuespr_comp = 2:1:50; % Pressure ration at compressort4 = 1800:10:2000; % Turbine entry temperature

% Calculation of t3

t3is = t2 * pr_comp .^((gamma - 1)/gamma);t3 = ((t3is - t2)/n_comp) + t2;

% Calculation of p3 and p4p3 = p2 .* pr_comp;p4 = p3 - (p3 * pl_cc);

% Calculation of pr_turbpr_turb = p4 / p5;

% Calculation of t5

t5is = zeros(numel(t4),numel(pr_turb));t5 = zeros(size(t5is));for i = 1:numel(t4)

for j = 1:numel(pr_turb)t5is(i,j) = t4(i) / (pr_turb(j) ^((gamma - 1)/gamma));t5(i,j) = t4(i) - ((t4(i) - t5is(i,j))* n_turb);

endend

% Calculation for total output needed

Wnet = Wg / n_gen;

% Calculation for mass flow rates

t3mint2 = t3 - t2;

t4mint5 = zeros(size(t5));for i = 1:numel(t4)

for j = 1:numel(t5(1,:))t4mint5(i,j) = t4(i) - t5(i,j) ;

endend

m5 = zeros(size(t4mint5));for i = 1:numel(t4)

for j = 1:numel(t3)


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m5(i,j) = Wnet /( Cp * (t4mint5(i,j) - t3mint2(j)));end

end

m4 = m5;m3 = m4 / (1 + far_cc);

m2 = m3 / 0.985;

m_fuel = m3 * far_cc;

% SFC calculationsfc_eng = m_fuel / Wg;

% Graphical outputsfigure

surf(pr_comp,t4,Wg * ones(numel(t4),numel(pr_comp)))hold onsurf(pr_comp,t4,Wnet * ones(numel(t4),numel(pr_comp)))

xlabel('Compressor pressure ratio')ylabel('Turbine inlet entry temperature')zlabel('Power out')hold off

figuresurf(pr_comp,t4,m2)hold onxlabel('Compressor pressure ratio')ylabel('Turbine inlet entry temperature')zlabel('Mass flow rate')hold off

figureplot(pr_comp,m2)hold onxlabel('Compressor pressure ratio')ylabel('Mass flow rate of air')hold off

figuresurf(pr_comp,t4,m_fuel)hold onxlabel('Compressor pressure ratio')ylabel('Turbine inlet entry temperature')zlabel('Mass flow rate')hold off

figureplot(pr_comp,m_fuel)hold onxlabel('Compressor pressure ratio')ylabel('Mass flow rate of fuel')hold off

figuresurf(pr_comp,t4,sfc_eng)hold on

xlabel('Compressor pressure ratio')ylabel('Turbine inlet entry temperature')


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zlabel('SFC')hold off


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2.5 Testing

The screen shot shows the results after the code was run. The code runs with

out any errors.

Fig 2. Screen shot after the code run


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3.PART 2

3.1 Overview

This part looks at the selection process for the different parameters namely

compressor pressure ratio, turbine temperature and thus mass flow rate of theair and fuel.

3.2 Selection of parameters

3.2.1 Range

In order to select a particular value for Pressure ratio and turbine inlet

temperature a reasonable range has to be taken into consideration. The range

that has been used for this purpose is:

Range for pressure ratio: 2 to 50Range for Turbine inlet temperature: 1800K to 2000K

This range has been selected on the bases of the possible values that can be

achieved using the latest technology.

3.2.2 Engine Output Required

Since the electric generator is not 100% efficient the output needed form the

engine is more than the output of the generator, which is set of 400kW. The

output thus needed is 421.5kW which has been calculated on the bases of the

given 95% efficiency of the engine.

05

1015

2025

3035

4045

50

1800

1850

1900

1950

2000

4

4.05

4.1

4.15

4.2

4.25

x 105

Powerout

Compressor pressure ratioTurbine inlet entry temperature

Fig 3.1 Pressure ratio range, Turbine inlet temperature range and Outputs

3.2.3 Analysis of the plots


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Using the MATLAB script developed in Part 1 plots comparing the follow

parameters have been produced

Compressor pressure ratio. These values have been chosen form therange selected in section 3.2.1. This parameter not only is important for

the calculations of other parameters but is also directly proportional tothe length and thus the weight of the engine.

Turbine inlet temperature. This parameter is also been chosen for therange defined in section 3.2.1. The turbine inlet temperature is very

important in the performance of a gas turbine engine.

Mass follow rate of air through the engine This is directlyproportional to the size of the engine. Higher the value bigger should

be the inlet diameter.

Mass flow rate of fuel This gives an idea of how much fuel will beused to produced the 400 kW of power that is required. His valued

helps to determine the running cost of the engine. Smaller the value

less will be the running cost.

SFC This defines the fuel consumption per unit of power generated.Its significance is same as the mass flow rate of the fuel.

The following are the different plots produced.

0 10

2030

4050

1800

1850

1900

1950

20000.5

1

1.5

2

2.5

3

Compressor pressure ratiourbine inlet entry temperature

Massflow

rate

Fig 3.2

Surface plot

For Comp.

Pressure

ratio, Turb.Inlet

temperature

and mass

flow rate


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0 5 10 15 20 25 30 35 40 45 50

0.8

1

1.2

1.4

1.6

1.8

2

2.2

2.4

2.6

Compressor pressure ratio

Massflow

rateofair

Fig 3.3 Plot

showing

compressor

pressure

ratio and

mass flowrate at

different

turbine inlet

temperature

s

010

2030

4050

1800

1850

1900

1950

20000.01

0.02

0.03

0.04

0.05

0.06

0.07

Compressor pressure ratioTurbine inlet entry temperature

Massflow

rate

Fig 3.4

Surface plot

For Comp.

Pressure

ratio, Turb.

Inlet

temperature

and massflow rate of

fuel


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0 5 10 15 20 25 30 35 40 45 500.015

0.02

0.025

0.03

0.035

0.04

0.045

0.05

0.055

0.06

0.065

Compressor pressure ratio

Massflow

rateoffuel

Fig 3.5 Plot

showing

compressor

pressure

ratio and

mass flowrate of fuel

at different

turbine inlet

temperature

s

010

2030

4050

1800

1850

1900

1950

20000

0.5

1

1.5

2

x 10-7

Compressor pressure ratiourbine inlet entry temperature

SFC

Fig 3.6

Surface plot

For Comp.

Pressure

ratio, Turb.Inlet

temperature

and SFC

From the analysis of these plots the following values where chosen for the

different parameters:

Compressor pressure ratio: The compressor pressure ratio chosen for the

design is 11:1. This is because as can be seen from the plots there is not much

significant difference in mass flow rate of the fuel flow rate after this point.

Another reason for choosing this value is that can be easily achieved using

current compressors without much increase in weight.

Turbine inlet temperature: The selection for this value is based on how

much fuel is consumed at the particular temperature. The temperature for theminimum value of the mass flow rate of fuel is chosen to be 2000K. Since this


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value is too high the material chosen for the turbine must be chosen to which

can withstand this temperature. For this reason the temperature chosen is

1900K.

Mass flow rate of air: Based on the values of turbine inlet temperature and

compressor pressure ratio the value for mass flow of air is 0.7649 kg/s

Mass flow rate of fuel: Based on the values of turbine inlet temperature and

compressor pressure ratio the value for mass flow of fuel is 0.0188 kg/s.


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4.PART 3

4.1 Overview

This part deals with the selection of different components of the engine, like

the compressor type, and the materials for the engine. These selections are tosupport the selections taken in part 2.

4.2 Selection for Compressor

The type of compressor chosen for this engine is an Axial compressor. This is

because it can compress the air to the required ratio of 11:1 using various

stages. The other option for a compressor is an Centrifugal type compressor

but since it can only give a pressure ratio of 4.5:1 maximum this cannot be

used.

Fig 4.1 Axial Compressor

The material chosen for the compressor is steel and nickel based alloys. This is

to keep the manufacturing cost low.

4.3 Selection for Combustion chamber

The type of combustion chamber chosen is an Annular combustion chamber

which because of its small size, less weight and low production cost.

Since the containing walls and internal parts of the combustion chamber mustbe capable of resisting the high gas temperature in the primary zone the walls

should be coated with high heat resistant coatings and by cooling the inner

wall of flame tube.


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Fig 4.2 Annular combustion chamber

4.4 Selection for Turbine

The turbine has to face a very high temperature and therefore the materialsrequired for manufacturing it are very important. The turbine can be made for

nickel alloys. A ceramic coating can be used to enhance the heat resistively of

the blades.


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5.CONCLUSION

5.1 Overview

In this section aims to summarize the over all design for the APU and also

look at the errors or defects in the design or the method used.

5.2 Design Summery

Following are the details of the full design

Inlet dry air properties

R 287 J/kg/K

Cp 1500J/kg/K

1.4

Compressor properties

Compressor pressure ratio 11:1

Compressor inlet pressure 1 bar

Compressor inlet temperature 15o

C

Compressor isentropic efficiency 93%

Bleed flow from exit of compressor 1.5%

Type Axial

Materials used Steel and Nickel

alloys

Combustion chamber propertiesCombustion Pressure Loss 1%

Combustion Fuel to Air Ratio 0.025:1

Type Annular

combustion

chamber

Materials High heat

resistive

coatings, Steel

and Nickel

alloys

Turbine properties

Turbine inlet temperature 1900K

Turbine exit pressure 1.2

Turbine isentropic efficiency 95%

Materials Nickel alloys

and ceramic

coatings

Mass flow

Mass flow rate of air 0.7649 kg/sMass flow rate of fuel 0.0188 kg/s


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Generator properties

Electrical generator efficiency 95%

Aircraft Engine power demand (Wg) 400Kw

5.3 Errors and defects

There have been errors in running the MATLAB script. Caution need to be

taken in selection the range of turbine inlet temperature. If the value is less

than the compressor exit temperature the plots will show negative values. This

how ever can be preventing by putting some extra code.

The analysis of the report does not show many comparisons between different

or more complex range.


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REFERENCES

Books:

1. The Jet Engine, By Rolls-Royce Plc2. Aircraft Engines and Gas Turbines, Second Edition by Jack L.

Kerrenbrock3. Aircraft Engine Design, By Jack D. Mattingly, Willian H.

Heiser & Daniel H. Daley

4. Aircraft Gas Turbine Powerplants, By Sanderson TranningProducts.

Internet

1. NASA Glenn Research Centre,url: http://www.grc.nasa.gov/WWW/K-12/aerores.htm

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Propulsion System Analysis_BhatU