Summer Internship at Cranfield University-Report

TITLE

GAS TURBINE PERFORMANCE SIMULATION

STUDENT INTERNSHIP PROJECT UNDERTAKEN

AT

CRANFIELD UNIVERSITY

BY

N.SATHYANARAYANAN

IV YEAR –MECHANICAL

INDIAN INSTITUTE OF INFORMATION TECHNOLOGY

DESIGN AND MANUFACURING – KANCHEEPURAM

REPORT SUBMITTED IN JULY 2013

CONTENTS

1. Acknowledgement

2. Introduction to Gas Turbines

3. A short note on Gas Turbine Performance Simulation and TURBOMATCH

4. Problem statement of the project

5. Engine Discussions

6. Analysis and Conclusions

7. Bibliography

ACKNOWLEDGEMENT

At the outset I would like to acknowledge with thanks to the Department of

Power and Propulsion, Cranfield University for permitting me to undergo a

summer research internship in their department. I was offered an interesting

problem to work on. Thanks are due to my supervisor Dr.Theoklis Nikolaidis,

Lecturer, Department of Power and Propulsion, for explaining the technicalities

and providing guidance at various stages in completing this internship study. I

would also like to express my gratitude to Prof.Pericles Pilidis, Head of

Department, for accepting my Visiting Student application, and to Ms. Faye

Winstanlet and Ms.Clarie Bellis, for helping me with the application process.

The internship has provided me with an insight into the component

characteristics and performance characteristics of Gas Turbines.

Thanks are also due to Dr.R.Gnanamoorthy, Director and Dr.K.Selvajyothi,

Assistant Professor of our institute IIIT DM, Kancheepuram for encouraging us

to undergo summer internship and also for providing with necessary letters of

introduction in securing this internship

N.SATHYANARAYANAN

INTRODUCTION TO GAS TURBINES

A gas turbine is a type of an internal combustion engine, operating on Brayton

cycle. With its high power-weight ratio the gas turbine engines have

dominated the area of aerospace propulsion for quite some time now. It is

becoming an increasingly popular prime mover in the power, process and oil

industries.

Some components of a gas turbine engine vary based on the application, even

though there are three basic components that are common to a gas turbine

engine. A compressor, combustor and a turbine. The compressor does the

function of increasing the pressure of the incoming air thus facilitating

combustion, and aiding power extraction in the turbine. From a compressor,

the high pressure air enters the combustor, where fuel is injected to increase

the energy and temperature of the gas. The hot, high pressure, high energy air

then passes through the turbine where the enthalpy of the gas is converted

into the rotational energy of the turbine. A typical ideal Brayton cycle is shown

below.

Image courtesy: Wikipedia

Since the project involved working on turbofan aero-engines, a bit more

detailed discussion on aero engines will be done in the forthcoming section.

An aero-engine has an intake, through which the air from the atmosphere

passes into the engine, and is followed by compressor(s), combustor(s) and

turbine(s) in the core. The aero-engine has a propelling nozzle in the core

following the turbine(s), that converts the pressure energy of the gases to

kinetic energy, thus accelerating the fluid. In case of turbofan engines there is

bypass duct through which the air sucked by the frontal fan is bypassed. It is

followed by a propelling nozzle. The thrust required by the aircraft if provided

by the propelling nozzle(s), according to Newton’s third law.

A descriptive picture and theory of thrust generation is given in the extract

below.

Image courtesy: Gas turbine Theory, Cohens, Rogers, Saravanamuttoo

A turbofan engine has a fan after the intake, which is a low pressure

compressor and sucks in huge amounts of air so as to provide the required

thrust. High mass flow rate facilitates reduction of exit velocity which also

reduces the noise generated by the aero-engine. The net thrust generated by

the turbofan engine equals the thrust generated by the bypass channel and

that generated by the core.

Turbofan engines can be classified into 2 spool or 3 spool engines. A spool can

be said to consist of a turbine which runs either a single compressor or

multiple compressors. Multi-spool engines facilitate running of the different

turbines and their corresponding compressors to run at different rotating

speeds, which is very important to avoid problems like stalling and surging in

compressors.

A cutaway diagram of a Rolls Royce Trent-900 is shown below: Trent-900 is a

high bypass three spool turbofan aero-engine.

Image Courtesy: Google

GAS TURBINE PERFORMANCE SIMULATION AND TURBOMATCH:

With development in computational power and numerical tools in the recent

past, a lot of emphasis has been laid on numerical simulations for getting the

required Design and Performance results required, since it involves lesser time

and money.

For an aeroengine, the performance of an engine could be talked about in

terms of Specific thrust, Thrust produced by the engine and the specific fuel

consumption, given a particular running condition.

Performance prediction is something that could not be done by manual

calculations and there arises a need for a computational tool to perform the

simulation that is required.

TURBOMATCH, an in-house code developed in the School of Mechanical

Engineering, Cranfield University is one such code that predicts the

performance of a gas turbine engine under a given condition.

COMPRESSOR AND TURBINE MAPS:

Compressor maps and turbine maps are a graphical representation of a

behaviour of compressors and turbines under different rotational speeds, mass

flow rates, entry total temperatures and entry total pressures. These maps are

crucial in predicting the performance characteristics of a gas turbine engine.

A typical compressor map has the pressure ratios and the efficiencies plotted

for a range of values of non-dimensional speeds and non-dimensional mass

flow rates. A typical turbine has non-dimensional mass flow rate plotted

against the ratio of total pressure ratio at the entry to that at the exit for

various non-dimensional speeds. In this project report, the compressor map

has corrected mass flow in (Kg/s) on the X-axis and pressure ratio on the Y-axis.

Corrected Flow is the mass flow that would pass through a device (e.g.

compressor, bypass duct, etc.) if the inlet pressure and temperature

corresponded to ambient conditions at Sea Level, on a Standard Day

Typical representations of a compressor map and a turbine map are given

below.

COMPRESSOR MAP

Image courtesy: Gas turbine Theory, Cohens, Rogers, Saravanamuttoo

TURBINE MAP

Image courtesy: Gas Turbine Theory, Cohens, Rogers, Saravanamuttoo

WORKING OF TURBOMATCH:

Turbomatch is an iterative FORTRAN code that runs a series of calculations

based on off-design matching process. Turbomatch has a set of inbuilt

compressor and turbine maps which are scaled according to fit the given

design conditions of a particular gas turbine engine. The scaled compressor

map is then used as a tool to predict the off-design performance

characteristics of the engine.

The modelling of the engine is done using a .dat file consisting of various

codewords called “BRICKS”, that represent the various components of an

engine like Intake, Compressor, Gas Duct, Turbine, Nozzle, Heat exchanger etc.

The data that is required by the code to do the performance calculations are

fed in the form of brick data and Station Vectors.

The numerical results are then extracted in excel and the comparison of value

is done using spreadsheets and graphs.

The compressor map is then plotted by using the scaled compressor map data

values and the running lines for various compressors at different conditions are

plotted and analysed.

PROBLEM STATEMENT:

3 turbofan high bypass aero-engines namely,

i.) CFM-56-7B27

ii.) Rolls Royce Trent 1000

iii.) Pratt and Whitney 4084, were chosen for the project.

The project involved simulating the off-design conditions for the above-

mentioned aero-engines on Turbomatch version 1 and the in-development

Turbomatch Version 2.

The project involved:

1.) Plotting the running lines and graphs for SFC, Specific thrust and Net thrust

vs TET, for different flight mach numbers and altitudes.

2.) Studying the variation of takeoff thrust, SFC and specific thrust with

variation in ambient temperature, for a constant TET.

3.) Finding the errors between numerical values and compare running lines

predicted by the two versions of Turbomatch, for different off design

conditions.

ENGINE DISCUSSIONS:

CFM56-7B-27:

CFM56-7B-27 is a high bypass 2 spool turbo-fan aero-engine powering Boeing-

777. The engine has a single stage fan (low pressure compressor) and the

intermediate 3-stage compressor(booster) on the same spool, which is driven

by a 4-stage low pressure Turbine. It has a 9-stage high pressure compressor

on the second spool which is driven by a 1-stage high pressure turbine.

Technical specifications at takeoff(Rated to ISA conditions):

Mass flow rate into the engine=354 Kg/s

Overall pressure ratio=32.8

Turbine entry temperature=1600 K

Bypass ratio=5.1

Net thrust=121 KN

PW 4084:

PW 4084 is a high bypass 2 spool turbo-fan aero-engine powering Boeing-737.

The engine has a single stage fan (low pressure compressor) and the

intermediate 5-stage compressor(booster) on the same spool, which is driven

by a 5-stage low pressure Turbine. It has a 15-stage(5 stages variable) high

pressure compressor on the second spool which is driven by a 2-stage high

pressure turbine.

Technical specifications at takeoff(Flat-Rated to +15 K deviation ISA

conditions):

Mass flow rate into the engine=1157 Kg/s



Bypass ratio=6.41

Net thrust=373.8 KN

TRENT-1000:

Trent-1000 is a high bypass 3 spool turbo-fan aero-engine powering Boeing-

787. The engine has a single stage fan (low pressure compressor) , which is

driven by a 6-stage low pressure Turbine, an 8-stage intermediate pressure

compressor on the second spool which is driven by a 1-stage intermediate

pressure turbine, and a 6-stage high pressure turbine driven by a single-stage

high pressure turbine.

Technical specifications at takeoff (Rated to +15K deviation from ISA

conditions):

Mass flow rate into the engine=1199.30 Kg/s



Bypass ratio=10.38

Net thrust=309 KN

ANALYSIS AND CONCLUSIONS:

CFM56-7B27:

RUNNING LINES:

The running lines of the compressors got for CFM56 are given below: (TM-2)

Running lines for the fan

Running lines for the booster compressor

Running lines for the high pressure compressor

COMPARISION OF RUNNING LINES OBTAINED ON TM-1 AND TM-2:

Running line comparison for fan

Running line comparison for Intermediate compressor:

Running line comparision for high pressure compressor

INFERENCES:

1.) It could be said that the operating region of the CFM-56 fan lies below

the surge line for altitudes between 0 and 8000 metres and between

speeds of M=0 and M=0.8.

2.) The surge margin of the fan is very low for non dimensional running

speeds beyond 1.06. So care needs to be taken when the engine is

operated in the region.

3.) The running lines for for different altitudes and a constant number are

almost constant for the fan.

4.) The intermediate compressor severely limits the running speed of the

spool, though.

5.) Since the square roots of total temperature at the entry of the fan and

intermediate compressor are comparable, we could say the same about

their non dimensional speeds at a particular running condition.

6.) The running lines of the intermediate compressor clearly imply that the

spool is free to operate at any speed for Mach numbers between 0 and

0.4 .

7.) Beyond a Mach number of 0.4, care needs to be taken to ensure that the

CN of the fan is lies between 0.85 and 1.15 to ensure smooth running of

both the fan and intermediate compressor.

8.) High pressure compressor could be operated safely for any Mach

number between 0 and 0.8 and at any altitude between 0 and 8000m,

without surging problem, eventhough the thrust required at a particular

condition might restrict the operating condition. Running lines lie in a

very narrow region implying good running stability.

COMPARISON BETWEEN RUNNING LINES PREDICTED BY TM-1 AND TM-2:

1.) The interpolation feature in Turbomatch-2 helps predict the running

lines beyond the surge line.

2.) Turbomatch-2 predicted fairly smooth running lines while Turbomatch-1

showed wavy fluctuations in running lines whenever the running line

approached the surge line. The differences are clearly visible in the

results predicted for the intermediate compressor.

SFC,THRUST, SPECIFIC THRUST VS TET PLOTS AND INFERENCES:

The figures below represent the plots predicted.

Thrust vs TET for Altitude=2000m

1.) As expected for a given TET, thrust decreases with increasing Mach

number owing it to the increased mass flow into the engine, and thus a

reduction in the overall pressure ratio of then engine, failure of ram

pressure increase to compensate for the decrease in pressure ratio and

also to higher momentum drag.

2.) As expected, increase in TET increases the net thrust produced, for a

given mach number.

Specific thrust vs TET for Altitude=2000m

1.) As expected the specific thrust increased with an increase in turbine

inlet temperature for a particular mach number.

2.) For a given turbine inlet temperature, increase in Mach number reduces

the specific thurst owing it to increase in momentum drag.

SFC vs TET for Altitude=2000m

1.) With the increase in turbine inlet temperature there is an increase in the

rotational speed of the spool, thus increasing the overall pressure ratio

and the overall thrust. So a drop in SFC with increase in TET is clearly

noticeable.

2.) Though SFC falls initially, the SFC starts increasing from a particular point

on the curve. This could be attributed to the choking of the nozzle that

occurs with an increase in mass flow rate due to increase in rotational

speed.

3.) After the point of choking, there is no increase in thrust due to

momentum thrust, eventhough there is an increase in pressure thrust.

But the increase in overall thrust does not compensate proportionally

for the increase in pressure thrust, thus increasing the SFC.

Similar curves obtained for flight altitude=4000m, 6000m and 8000m

show a similar trend.

VARIATION OF THRUST,SFC AND SPECIFIC THRUST WITH AMBIENT

TEMPERATURE FOR CONSTANT TURBINE INLET TEMPERATURE:

Thrust vs Ambient temperature(TET=1600 K)

SFC vs Ambient Temperature(TET=1600 K)

Specific thrust vs Ambient Temperature(TET=1600 K)

INFERENCES:

1.) For a given TET, thrust increases almost linearly, with decrease in

ambient temperature.

2.) SFC and specific thrust increase with decrease in ambient temperature.

There is an increase in fuel consumption, because lower ambient

temperature thermodynamically causes lesser temperature at the

entrance of combustor, thus requiring more fuel to be burnt to reach

the required TET.

3.) The graphs clearly imply that achieving a certain amount of thrust

consumes lesser fuel with reduction in ambient temperature and

requires a lesser TET, since a lower rotational speed would balance out

the reduction in temperature and give the required value of CN, to

generate the required amount of thrust.

NUMERICAL ERROR ANALYSIS:

The figure below shows the percentage of numerical error between the

values predicted by Turbomatch 1 and Turbomatch 2.

Numerical errors in thrust values predicted for Ambient temperature Off-

design conditions

TM1 value TM2 value %error

Atttribute: Temperature deviation

from ISA condtions(K)

168450 163260.0000 3.0810329 -50.0

164630 159470.0000 3.1343012 -45.0

160880 155510.0000 3.3378916 -40.0

157220 151560.0000 3.6000509 -35.0

153590 147490.0000 3.9716127 -30.0

149930 143190.0000 4.4954312 -25.0

146010 138900.0000 4.8695295 -20.0

142520 134700.0000 5.4869492 -15.0

138510 130630.0000 5.6891199 -10.0

134510 126660.0000 5.8359973 -5.0

130460 122480.0000 6.1168174 0.0

126470 118350.0000 6.420495 5.0

122760 114200.0000 6.9729554 10.0

119150 110260.0000 7.4611834 15.0

114470 105600.0000 7.7487551 20.0

110120 101180.0000 8.1184163 25.0

106180 97072.0000 8.5778866 30.0

102050 93284.0000 8.5899069 35.0

98008 89427.0000 8.7554077 40.0

94160 85562.0000 9.1312659 45.0

90149 81765.0000 9.3001586 50.0

As expected there were differences between the values predicted by

Turbomatch-1 and Turbomatch-2. The error varied between 3 to 9 percent,

with Turbomatch-1 predicting higher values than those predicted by

Turbomatch-2.

Similarly, errors were found for net thrust values predicted for various off-

design altitudes and mach numbers. The maximum error between the thrust

values predicted by the two versions was found to be around 5 KN and errors

were found to lie within +/- 7 percentage, unless the value of net thrust was

very low (< 5KN)

PW-4084:

RUNNING LINES:

The running lines of the compressors of the PW-4084 engine are shown

below.(As predicted by Turbomatch-2)


Running lines for booster compressor

Running lines for high pressure compressor



Running lines comparision for intermediate compressor

Running lines comparison for high pressure compressor

INFERENCES:

1.) There were striking similarities between the running lines of the

respective compressors of CFM-56 and PW-4084

2.) This could be owed to the fact that both engines are high bypass 2 spool

engines with similar configuration of the spools.

3.) Like in CFM-56 engine, the intermediate compressor restricts the

operating speed of the spool.

4.) The running lines for a given Mach number and various altitude almost

lie on a same line.

5.) CN of the fan needs to be kept below 1.10 to ensure a safe surge margin,

always.

6.) Even though for M<0.4, the spool can be operated at any CN<1.10, for

mach numbers close to 0.4, CN must be kept above 0.5 to ensure

smooth operation and giving the compressor sufficient surge margin.

7.) Above a speed of M=0.4, care must be taken to ensure that the non-

dimensional speed of the intermediate compressor is above 0.80 and

below 1.10 to ensure that there is fairly sufficient surge margin.

8.) The high pressure compressor is safe to operate at any non-dimensional

speed less than 1.15 for any altitude between 0 to 8000m and for any

mach number between 0 and 0.8, without any surging problem, even

though thrust required might restrict the operating range.


1.) The interpolation feature in Turbomatch-2 helps predict the running

lines beyond the surge line, while the running line predicted just

coincided with the surge line.






The reasons for the trends observed in the graphs are the same as those

mentioned in the inferences section of CFM-56 engine.




Specific thrust vs Ambient temperature(TET=1705 K)

SFC vs Ambient temperature(TET=1705 K)

INFERENCES:

The reasons for the trend observed are the same as those mentioned in CFM-

56 case study.


The figure below shows the percentage of numerical error between the values

predicted by Turbomatch 1 and Turbomatch 2.

Numerical errors in thrust values predicted for rotational speed(PCN) off-

design conditions(Altitude=2000m,M=0.2)

TM1 value TM2 value %error Attribute: Change in

PCN(Alt=2000m,M=0.2)

350890 329190.0000 6.1842743 1.0500

288150 277470.0000 3.7064029 1.0000

236840 233840.0000 1.2666779 0.9500

202390 201140.0000 0.6176194 0.9000

172190 173280.0000 -0.633022 0.8500

145400 173280.0000 -19.17469 0.8000

121620 127760.0000 -5.048512 0.750

100360 104890.0000 -4.51375 0.700

80586 86080.0000 -6.817561 0.650

62412 68379.0000 -9.560661 0.600

48418 54361.0000 -12.27436 0.550

37715 41899.0000 -11.09373 0.500


Turbomatch-1 and Turbomatch-2.


design altitudes and mach numbers.

The maximum error between the thrust values predicted by the two versions

was found to be around 20 KN and errors were found to lie within +/- 8

percentage, unless the value of net thrust was lower than or 50KN where

error was found to be higher more than 20 percent since there was around 15

KN error constantly between the 2 versions of Turbomatch.

For ambient temperature off-design conditions, the error was found to be very

low and less than (+/-) 1 percent.

TRENT 1000:

RUNNING LINES:

The running lines of the compressors of the Trent-1000 engine are shown

below.(As predicted by Turbomatch-2)


Running lines for intermediate compressor

Running lines for the high pressure compressor



Running line comparison for intermediate compressor

Running lines comparison for high pressure compressor

Inferences:

1.) All the compressors are safe to work at any rotational speed for altitudes

between 0 to 8000 m and Mach numbers between 0 to 0.85, without

surging problems, even though the required thrust at a particular

operating condition might restrict the compressor working range.

2.) The intermediate compressor and the high pressure compressor have a

narrow working region, which implies good functional stability.


1.) There were only subtle differences between running lines predicted by

the 2 versions of turbomatch, as the surge margin of the running lines

was quite sufficient for Turbomatch-1 to give smooth running lines

without any problems.






The reasons for the trends observed in the graphs are the same as those

mentioned in the inferences section of CFM-56 engine.




Specific thrust vs Ambient temperature(TET=1820 K)

SFC vs Ambient temperature(TET=1820 K)

INFERENCES:

The reasons for the trend observed are the same as those mentioned in CFM-

56 case study.


The figure below shows the percentage of numerical error between the values

predicted by Turbomatch 1 and Turbomatch 2.

Numerical errors in thrust values predicted for rotational speed(PCN) off-

design conditions(Altitude=2000m,M=0.2)

TM1 value TM2 value %error

Attribute: Change in PCN at Alt=2000m,M=0.2

202580 205150.0000 -1.268635 1.000

177830 177860.0000 -0.01687 0.950

154070 155970.0000 -1.233206 0.900

132450 135490.0000 -2.295206 0.850

95510 95776.0000 -0.278505 0.750

77830 78814.0000 -1.264294 0.700

61078 63256.0000 -3.565932 0.650

48521 50381.0000 -3.833392 0.600

38414 39112.0000 -1.817046 0.550

29633 30581.0000 -3.199136 0.500

22013 23075.0000 -4.824422 0.450

15437 16335.0000 -5.817192 0.400

10113 10223.0000 -1.087709 0.350

6300.3 5188.5000 17.646779 0.300


Turbomatch-1 and Turbomatch-2.


design altitudes and mach numbers.

The maximum error between the thrust values predicted by the two versions

was found to be around 3 KN and errors were found to lie within (+/-) seven

percentage, unless the value of net thrust was lower than or around 3KN

where error was found to be higher more than 20 percent since there was

around 2 KN error constantly between the 2 versions of Turbomatch.

For ambient temperature off-design conditions, the error was found to be less

than (+/-) 5 percent.

CONCLUSION:

The following things were done successfully using TURBOMATCH versions 1

and 2 for the 3 engines, namely CFM56-7B27, PW-4084 and TRENT-1000

1.) Plotting the running lines and graphs for SFC, Specific thrust and Net thrust

vs TET, for different flight mach numbers and altitudes.

2.) Studying the variation of takeoff thrust, SFC and specific thrust with

variation in ambient temperature, for a constant TET.

3.) Finding the errors between numerical values and compare running lines

predicted by the two versions of Turbomatch, for different off design

conditions.

BIBLIOGRAPHY:

1.) H Cohen, GFC Rogers, HIH Saravanamutto, Gas Turbine Theory

4th edition, Longman group limited

2.) Dr. Vasilios Pachidis, Gas Turbine Performance simulation

3.) www.wikipedia.org

4.) www.google.com

5.) www.cfmaeroengines.com

6.) www.pw.utc.com

7.) www.rolls-royce.com

http://www.wikipedia.org/

http://www.google.com/

http://www.cfmaeroengines.com/

http://www.pw.utc.com/

http://www.rolls-royce.com/

Documents

Summer Internship at Cranfield University-Report