1 INTRODUCTION
1.1
Background
Moran and Shapiro [1] describes the basic operation of a gas turbine engine. The engine
functions using the Brayton cycle, consisting of a: compressor, combustor and turbine. In
particular and aircraft engine can have a diffuser placed before the compressor and a nozzle
after the turbine. Fundamentally air enters the front of the engine through the diffuser, which
causes a pressure rise resulting from a drop in velocity. This is referred to as the ‘ram effect’ [1].
Then the compressor draws the air into the combustor where the pressure rises further. A fuel
mixture (usually aviation kerosene, calorific value of 46.36 MJ/kg [2]) is then combusted which
increases the temperature and hence the temperature of the gas inside the combustor. This gas
is then allowed to expand through a turbine (generally a reaction axial type). The turbine is used
to drive the compressor and auxiliary equipment. The gas exiting the turbine is still at a relatively
high temperature and pressure which then escapes to the atmosphere through the nozzle at a
high velocity. (The nozzle is used to increase the exit velocity). This discharge of gas at a high
velocity relative to the engine generates thrust due to the momentum transfer.
1.2
1.2.1
Theory
The gas turbine
Gas turbines are suited for aircraft propulsion because of their superior power to weight
ratio. The operation of the gas turbine may be classified as a thermodynamic Brayton
cycle. This type of cycle may operate on an open or closed basis. With the closed cycle,
the gas exiting the turbine (nozzle) is passed through a heat exchanger, where it is
cooled before it reenters compressor, while for open cycle, the gas exiting the nozzle is
discharged out the back of the engine and is dissipated to the surrounding atmosphere.
[1b]
The most popular aircraft propulsion engine used is the turbojet engine. This type of
engine consists of three main sections, the diffuser, the gas generator and the nozzle.
1
The diffuser is placed before the compressor and which causes the incoming air to
reduce its speed. The gas generator consists of a compressor, combustor and turbine.
The incoming air is compressed to a higher pressure before it enters the combustion
chamber. The air is then mixed with fuel and combustion occurs. The expanding gas
then expands through the turbine and the nozzle where the high velocity of the gas
creates the propulsive force which drives the aircraft forward. [1b]
1.2.2
The Brayton cycle
The process in which thrust is created is called the Brayton cycle for gas turbines. The
following subdivision is a more methodological look into how the Brayton cycle works..
The following diagram represents each component in the open Brayton cycle for the gas
turbine.
Figure 1.1: The schematic diagram of the Brayton cycle [2b].
All gas turbines make use of this open Brayton thermodynamic cycle. This cycle
assumes that air is an ideal gas and that there are standard free stream conditions. The
air enters the cycle at (1) and is thereafter compressed. The compression is isentropic
which results in an increase in temperature of the gas concerned. This means that there
is work done on the air. The temperature of the air then increases at a constant
pressure. There is, therefore, an increase in entropy. Once the temperature stops
2
increasing the air then does work on the system. This results in a decrease in
temperature and pressure. The flow decreases in an isentropic manner back to the
initial free stream nature. The air then leaves the system and returns entirely back its
free stream conditions. This completes an open Brayton cycle.
The open Brayton thermodynamic cycle is based on ideal conditions. But in reality there
are slight differences based on the certain factors. In a gas turbine engine, the
components are based on the processes which occur in an open Brayton cycle. The
main sections of a gas turbine engine are the diffuser, the gas generator and the nozzle
sections. The main components include the compressor, the combustors and the
turbine. In real actual conditions the performance of the engine would differ to ideal
situations
1.2.3
General definition and formulae
Thrust: This is the propulsive force which is formed due to the overall change in the
velocity of the gases relative to the turbine. [b1]
Efficiencies: Efficiency is normally referred to as the ratio of power output to power
input. As the effect of irreversibilities in the turbine and compressor become present, the
work developed by the turbine decreases and the work input into compressor increases.
The following formulas are used to calculate the efficiency of the compressor and
turbine. [2b]
comp 
h2  h1
'
h2  h1
(1)
Where  comp is the compressor efficiency.
All h correspond to enthalpy at that stage of the Brayton cycle [kJ/kg]
3
turbine 
T4  T3
T4  T3
'
(2)
Where turbine is the turbine efficiency.
Specific fuel consumption: This is the fuel flow through unit thrust. It can be
determined using flow divided by specific thrust. It may be calculated using the following
formula. [2b]
.
(m)(  )(3600)
Specific _ fuel 
thrust
(3)
.
Where m is the fuel flow rate in m3/s.
 is the density of air in kg/m3
1.3
Motivation
Gas turbine engines are particularly suited to aircraft applications because of their relatively high
power to weight ratio [1]. From a commercial aspect decreasing the specific fuel consumption
and maximizing efficiencies can increase profitability. Hence the parameters
1.4
Objectives
1. Determine the thrust, specific fuel consumption and performance parameters of a SR-30
gas turbine.
2. Determine and analyze the efficiencies of the different components within the engine.
3. Contrast the conditions of real and ideal Brayton Cycle.
4
2 APPARATUS
2.1 Test Rig: MiniLab Test Cell
The MiniLab unit is an integrated test cell consisting of the SR-30 turbo-jet engine (and all
necessary components including fuel and oil pumps, motors, operator instrumentation, sensors,
transducers and data acquisition hardware [1]. The entire rig (comprising the SR-30 engine,
control panel and data acquisition system) is schematically represented in Figure 1 below.
Figure 1.2: Schematic of Entire Test Rig
The mini lab is an important item of equipment which was ideal for investigating the
performance of the gas turbine because it was suited for;-
5
•
Purpose Built Gas Turbine Engine Designed and Manufactured to Aerospace Standards
•
All Key Engine Stations Fully Instrumented for Temperature and Pressure Measurement
•
Fully Instrumented Operator Control Panel
•
DigiDAQTM Data Acquisition System Utilizing USB Technology
•
User Configurable Real Time Computer Data Display
2.2 Apparatus: SR-30 Turbo-Jet Engine
The SR-30 Turbo-Jet engine used was specifically designed by Turbine Technologies for use in
the MiniLab unit [2]. The engine features a centrifugal flow compressor, reverse flow annular
combustor and an axial flow turbine stage. The SR-30 follows the Brayton Open cycle as
described in Section 1.2 of the theory.
Figure 1.3: Schematic of SR-30 Turbo-Jet Engine
6
Details of the SR-30 turbo jet engine are:-
•
Dimensions = 36cm x 28cm x 20cm
•
Weight = 15kg
•
Intake Suppressor = 25 cm diameter x 81 cm
•
Exhaust Ducting = 15cm diameter x 91 cm
•
Exhaust Expansion Chamber = 25 cm diameter x 25 cm
Instrumentation: Data Acquisition System
The program used to record the data measured by the different temperature and pressure
sensors was the DigiDAQ Data Acquisition system. Readings was recorded at three second
intervals.
A High Speed Data Acquisition System which was connected to a Windows® XP Computer for
On-Screen Data Display via a single cable DigiDAQ™ USB to PC Connection.
Sensors (Preinstalled and Calibrated) were used to measure various sources of information
directly from the SR Turbine (see Figure 2) such as;-
•
Compressor Inlet Temperature and Pressure (T1/P1)
•
Compressor Exit Temperature and Pressure (T02/P02)
•
Turbine Stage Inlet Temperature and Pressure (T03/P3)
•
Turbine Stage Exit Temperature and Pressure (T04/P04)
•
Thrust Nozzle Exit Temperature and Pressure (T05/P05)
•
Fuel Flow
•
Thrust
•
Engine Rotational Speed (RPM)
7
Digital and Analog outputs as provided on the operator control panel was used to show the
measurements of the following:-
•
Digital Turbine Inlet Temperature (TIT) ±1oC
•
Digital Exhaust Gas Temperature (EGT) ±1oC
•
Digital Engine Rotational Speed (RPM) ±50rpm
•
Analog Oil Pressure ±1psi
•
Analog Engine Pressure ±1psi
•
Analog Air Start Pressure ±1psi
4.
5. [1] turbine technologies ltd- mini-lab
6. [2] turbine technologies ltd- sr-30 turbo jet engine
8
2.1
3 PROCEDURES
3.1
Method
1. Record the values of ambient temperature and pressure.
2. Start the turbine and bring it to idle speed (between 45 000 and 50 000 RPM).
3. Allow two minutes for the readings to stabilise and record the values for Turbine Intake
Temperature (T.I.T.), Exhaust Gas Temperature (E.G.T.) and revolutions per minute
(RPM).
4. Increase the turbine speed by approximately 7000 RPM.
5. Wait for two minutes and then record the values for TIT, EGT and RPM.
6. Repeat steps 4 and 5 until the first value over 80 000 RPM is reached.
7. Now decrease the speed to idle, waiting for 2 minutes at approximately 7000 RPM
intervals, recording the values for TIT, EGT and RPM at the end of the 2 minutes.
8. Repeat steps 4 to 6.
9. Decrease the speed to idle and wait 2 minutes.
10. Record the values of TIT, EGT and RPM.
11. Switch off the fuel supply to the turbine to stop it.
9
3.2
Precautions
1.
2.
3.
4.
5.
Move the throttle gently and constantly.
Do not “jockey” the throttle (move it up and down) to try get to a specific speed.
Do not stand behind the exhaust as you may get burned.
Do not stand or walk in front of the intake as this will disrupt the airflow into the turbine.
The maximum allowable Turbine Intake Temperature is 900°C, do not allow it to exceed
this.
6. The maximum allowable Exhaust Gas Temperature is 720°C, shut down the turbine
immediately if this value is exceeded.
10
4 OBSERVATIONS
The running plots obtained from the data acquisition system fitted to the engine are now
presented.
Firstly the turbine inlet and outlet temperatures are shown in
Figure 4-1.
800.00
900.00
800.00
700.00
Turb Inlet Temp
Turb Exit Temp
700.00
600.00
Secondary Axis
600.00
500.00
500.00
Temperature [deg C] 400.00
Temperature [deg C]
400.00
300.00
300.00
200.00
200.00
100.00
100.00
0.00
0.00
14:16:48 14:24:00 14:31:12 14:38:24 14:45:36 14:52:48 15:00:00 15:07:12 15:14:24 15:21:36
Figure 4-1 Running plot of turbine inlet and exit temperatures
Similarly the compressor temperatures are shown in .
Finally the engine speed is shown in .
11
RPM
90000.00
80000.00
70000.00
60000.00
50000.00
RPM
40000.00
30000.00
20000.00
10000.00
0.00
14:16:48
14:24:00
14:31:12
14:38:24
14:45:36
14:52:48
15:00:00
15:07:12
15:14:24
15:21:36
Figure 4-2 Running plot of engine speed
12
5 DATA PROCESSING
As can be seen in the running plots, there are distinct regions in the results where the values
form plateaus. The intervening steps are transient conditions where the system was not in
equilibrium. Hence only the values that represent dynamic equilibrium conditions were used.
These values were averaged over the time interval to reduce the error associated with system
noise.
The averaged results are shown in Table 5-1.
Table 5-1 Average results from data acquisition system
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
RPM
Thrust Comp Inlet Comp Exit Temp Turb Inlet Temp Turb Exit Temp Comp Inlet Press Comp Exit Press Turb Inlet Press Turb Exit Press Noss Exit Press Fuel Flow
°C
°C
°C
°C
RPM
N
Pa (abs)
Pa (abs)
Pa (abs)
Pa (abs)
Pa (abs)
m^3/h
63023.64 25.70
17.95
147.71
539.78
535.49
85062.70
167091.89
165650.21
93947.22
87813.43 1.35E-02
69593.10 34.87
17.58
172.69
581.69
546.86
85570.92
190620.05
189264.83
96893.23
89845.35 1.54E-02
76406.06 47.07
17.41
200.00
624.36
569.24
86233.18
221035.86
219740.44
101073.31
92690.66 1.70E-02
82051.72 60.30
17.07
220.83
672.97
598.14
86788.25
250606.55
249393.98
104784.43
96317.85 1.82E-02
78213.33 53.05
16.91
203.93
645.80
586.00
86404.91
229598.20
228632.90
102194.69
93585.13 1.73E-02
73325.93 45.50
16.90
187.00
619.93
572.26
85868.61
206737.44
205588.28
99290.36
91090.42 1.62E-02
67379.17 37.43
16.51
165.75
564.92
544.21
85429.21
182241.04
180775.85
96205.81
89229.22 1.46E-02
60904.55 31.04
16.20
139.36
522.41
524.18
84993.44
159545.00
158166.00
93625.67
87834.18 1.28E-02
54994.59 25.96
16.08
119.65
516.92
515.73
84700.15
142096.92
141109.26
91802.56
87039.98 1.13E-02
50165.12 22.31
15.70
106.19
507.93
514.35
84503.03
130318.22
129474.78
90363.30
86380.07 9.94E-03
57624.44 26.28
15.24
124.29
518.47
513.09
84829.40
149295.97
148483.89
92818.25
87524.58 1.19E-02
63938.10 30.95
15.04
146.93
523.12
521.07
85178.81
169329.33
168541.33
95019.29
88469.86 1.35E-02
70084.09 37.11
14.48
169.86
570.43
538.95
85632.32
192578.32
191967.17
97649.84
90058.44 1.51E-02
75813.89 44.84
14.16
192.94
612.56
562.06
86176.84
218209.96
217558.76
101178.83
92406.85 1.64E-02
81881.82 56.71
13.83
216.85
661.82
591.15
86823.85
249723.24
249159.11
105139.27
96094.28 1.76E-02
48912.50 16.44
14.45
107.83
502.42
517.29
84447.25
127175.85
126859.83
89952.62
86208.34 9.39E-03
5.1
Sample Calculation
The pressures obtained from the data acquisition system where given in PSIG (pounds per
square inch, gauge), and hence were converted to SI units.
13
5.2
Processed Results
The results for the turbine are shown in Table 5-2.
Table 5-2 Turbine results
1.00
2.00
3.00
4.00
5.00
6.00
7.00
8.00
9.00
10.00
11.00
12.00
13.00
14.00
15.00
16.00
RPM
RPM
63023.64
69593.10
76406.06
82051.72
78213.33
73325.93
67379.17
60904.55
54994.59
50165.12
57624.44
63938.10
70084.09
75813.89
81881.82
48912.50
Thrust
Turb Inlet Temp Turb Exit Temp Theoretical Turb Exit Temp
Turb Inlet Press Turb Exit Press
N
°C
°C
°C
Pa (abs)
Pa (abs)
280341.74
539.78
535.49
427.75
165650.21
93947.22
309564.04
581.69
546.86
444.42
189264.83
96893.23
339869.44
624.36
569.24
459.44
219740.44
101073.31
364982.48
672.97
598.14
481.05
249393.98
104784.43
347908.55
645.80
586.00
471.34
228632.90
102194.69
326168.38
619.93
572.26
465.17
205588.28
99290.36
299716.01
564.92
544.21
437.50
180775.85
96205.81
270915.60
522.41
524.18
420.49
158166.00
93625.67
244626.96
516.92
515.73
432.93
141109.26
91802.56
223144.47
507.93
514.35
437.83
129474.78
90363.30
256325.05
518.47
513.09
426.96
148483.89
92818.25
284409.44
523.12
521.07
412.32
168541.33
95019.29
311748.05
570.43
538.95
433.78
191967.17
97649.84
337235.34
612.56
562.06
451.89
217558.76
101178.83
364226.70
661.82
591.15
473.01
249159.11
105139.27
217572.58
502.42
517.29
435.75
126859.83
89952.62
Isentropic
Compressor
Effeciency
%
43.31
43.66
44.40
45.50
45.15
44.03
42.52
42.05
40.45
38.28
42.12
43.00
43.69
44.18
44.97
34.80
Similarly for the compressor, the results are shown in Table 5-3.
14
Table 5-3 Compressor results
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
RPM
Comp Inlet
RPM
?C
63023.64
17.95
69593.10
17.58
76406.06
17.41
82051.72
17.07
78213.33
16.91
73325.93
16.90
67379.17
16.51
60904.55
16.20
54994.59
16.08
50165.12
15.70
57624.44
15.24
63938.10
15.04
70084.09
14.48
75813.89
14.16
81881.82
13.83
48912.50
14.45
Theoretical
Comp Exit
Temp
Comp Exit Temp
Comp Inlet Press Comp Exit Press
?C
?C
Pa (abs)
Pa (abs)
147.71
74.15
85062.70
167091.89
172.69
85.30
85570.92
190620.05
200.00
98.48
86233.18
221035.86
220.83
109.79
86788.25
250606.55
203.93
101.36
86404.91
229598.20
187.00
91.80
85868.61
206737.44
165.75
79.96
85429.21
182241.04
139.36
67.99
84993.44
159545.00
119.65
57.98
84700.15
142096.92
106.19
50.34
84503.03
130318.22
124.29
61.17
84829.40
149295.97
146.93
71.75
85178.81
169329.33
169.86
82.36
85632.32
192578.32
192.94
93.15
86176.84
218209.96
216.85
105.12
86823.85
249723.24
107.83
46.95
84447.25
127175.85
Isentropic
Turbine
Effeciency
%
3.83
25.37
33.42
38.99
34.28
30.80
16.25
-1.74
1.42
-9.16
5.88
1.85
23.04
31.43
37.43
-22.31
Finally the thrust, nozzle pressure and fuel flow rate are shown in Table 5-4.
15
Table 5-4 Thrust, nozzle exit velocity and fuel flow
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
16
RPM
Thrust
Noss Exit Press Fuel Flow
RPM
N
Pa (abs)
m^3/h
63023.64
25.70
87813.43 1.35E-02
69593.10
34.87
89845.35 1.54E-02
76406.06
47.07
92690.66 1.70E-02
82051.72
60.30
96317.85 1.82E-02
78213.33
53.05
93585.13 1.73E-02
73325.93
45.50
91090.42 1.62E-02
67379.17
37.43
89229.22 1.46E-02
60904.55
31.04
87834.18 1.28E-02
54994.59
25.96
87039.98 1.13E-02
50165.12
22.31
86380.07 9.94E-03
57624.44
26.28
87524.58 1.19E-02
63938.10
30.95
88469.86 1.35E-02
70084.09
37.11
90058.44 1.51E-02
75813.89
44.84
92406.85 1.64E-02
81881.82
56.71
96094.28 1.76E-02
48912.50
16.44
86208.34 9.39E-03
16
6 RESULTS
6.1
Turbine
300000.00
106000.00
104000.00
250000.00
102000.00
200000.00
100000.00
Turb Inlet Press
Turb Exit Press
Outlet Pressure [Pa abs]
98000.00
150000.00
Inlet Pressure [Pa abs]
96000.00
100000.00
94000.00
92000.00
50000.00
90000.00
0.00
45000.00
88000.00
50000.00
55000.00
60000.00 65000.00 70000.00
Engine speed [r.p.m]
75000.00
80000.00
85000.00
Figure 6-1 Turbine pressures for various engine speeds
17
800.00
700.00
600.00
500.00
Temperature [deg] 400.00
Turb Inlet Temp
Turb Exit Temp
Theoretical Turb Exit Temp
300.00
200.00
100.00
0.00
40000.00 45000.00 50000.00 55000.00 60000.00 65000.00 70000.00 75000.00 80000.00 85000.00
Engine Speed [r.p.m]
Figure 6-2 Turbine temperatures at different engine speeds
18
Isentropic Turbine Effeciency
50.00
40.00
30.00
20.00
Isentropic Effeciency [%]
10.00
Isentropic Turbine Effeciency
0.00
45000.00 50000.00 55000.00 60000.00 65000.00 70000.00 75000.00 80000.00 85000.00
-10.00
-20.00
-30.00
Engine Speed [r.p.m]
Figure 6-3 Isentropic turbine efficiency variations with engine speed
6.2
Compressor
19
300000.00
87000.00
250000.00
86500.00
Comp Exit Press
Comp Inlet Press
200000.00
86000.00
Compressor Exit Pressure [Pa
abs]
Compressor Inlet Pressure [Pa
abs]
150000.00
85500.00
100000.00
85000.00
50000.00
84500.00
0.00
45000.00
84000.00
50000.00
55000.00
60000.00
65000.00
70000.00
75000.00
80000.00
85000.00
Engine Speed [r.p.m]
Figure 6-4 Compressor pressures for various engine speeds
20
250.00
200.00
150.00
Temperature [deg C]
Comp Inlet Temp
Comp Exit Temp
100.00
50.00
0.00
45000.00
50000.00
55000.00
60000.00
65000.00
70000.00
75000.00
80000.00
85000.00
Engine speed [r.p.m]
Figure 6-5 Compressor temperatures at various engine speeds
21
Isentropic Compressor Effeciency
50.00
45.00
40.00
35.00
30.00
Isentropic Effeciency [%] 25.00
Isentropic Compressor Effeciency
20.00
15.00
10.00
5.00
0.00
45000.00 50000.00 55000.00 60000.00 65000.00 70000.00 75000.00 80000.00 85000.00
Engine Speed [r.p.m]
Figure 6-6 Isentropic compressor efficiency variations with engine speed
6.3
Performance Parameters
22
Figure 6-7 Thrust vs. engine speed
Figure 6-8: previous experiment and theoretical thrust vs. engine speed [5]
23
Figure 6-9: TSFC VS. ENGINE SPEED
Figure 6-10: previous experiment thrust specific fuel consumption vs. speed [5]
24
Goodenough, J.L.
0611973D
7 DISCUSSIONS AND CONCULSIONS
7.1
Discussion: 0611973d; Goodenough, J. L.
25
7.2
Conclusions: 0611973d; Goodenough, J. L.
26
Mugoni, T.
0501642w
7.3
Discussion: 0501642w; Mugoni, T.
7.4
Conclusions: 0501642w; Mugoni, T.
27
Mclean, B.
7.5
Discussion: 0610942G; Mclean, B.
7.6
Conclusion: 0610942G; Mclean, B.
0610942G
Mclean, B.
0610942G
28
8 REFERENCES
1. M. J. Moran, H. N. Shapiro, Fundamentals of Engineering Thermodynamics, John Wiley
& Sons, Inc. 5th edition pp 414-419.
2. Internet:
www.med.govt.nz/upload/21659/edf012003l.pdf,
author:
Unkown,
date
accessed: 03/07/2008.
3.
[1b]: Fundamentals of flight, Eric, E, Edison, McHILL INTERNATIONAL EDITION, 1967.
4.
[2b]: Fundamentals of thermodynamics, Moran Use previous.
5. A. Pourmovahed, C.M. Jeruzal, K.D. Brinker, Development of a Jet Engine
Experiment for the Energy Systems Laboratory, 2003 ASME International Mechanical
Engineering Congress, Washington DC, November 15-21, 2003
Put in fuel props address for density, etc
29