Transactions on Engineering and Sciences
Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
Performance Analysis Of A Micro Turbojet
Engine
Sajesh M1 Jyothi Sankar P R2 Vishnu Prakash K3 Nikhil Das K4
1 Assistant
Professor, ME Department NSS College of Engineering, Kerala, India
Professor, ME Department NSS College of Engineering, Kerala, India
3Student, ME Department NSS College of Engineering, Kerala, India
4Student, ME Department NSS College of Engineering, Kerala, India
2 Assistant
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Abstract— The reason behind doing this work was to prepare an experimental procedure of a
micro turbojet engine for the laboratory education purpose. By operating a micro turbojet engine
one can get an idea about the basic processes in a Brayton cycle and it helps one to know the
thermodynamics behind each process. The paper presents both the theoretical and experimental
procedures that could be done on a micro turbojet engine. The experiments were conducted on a
standard test rig with a SR-30 micro turbojet engine on which certain parameters like thrust, TSFC
(Thrust Specific Fuel Consumption) and various component efficiencies could be evaluated. The
main components in the test rig include a diffuser, axial flow compressor, reversible annular flow
combustion chamber and an axial flow turbine. The engine on which the experiment has been
conducted runs at 90000 RPM (Revolutions Per Minute) and produces a thrust of 256N. The
efficiencies obtained from theoretical and experimental calculations are found to have a difference
of 18.35% for compressor and 16% for turbine respectively. There is a difference in the overall
thrust of 56N.
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Keywords: Turbo jet engine, Performance, Brayton Cycle, Efficiency, Thrust, TSFC
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1. INTRODUCTION
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Brayton cycle is the basic thermodynamic cycle behind the operation of a jet engine. During the
ideal cycle analysis one may take certain assumptions like; the combustor pressure loss is negligible,
the specific heat is constant and the both the compressor and turbine work is equal. When the jet
engine is tested at the static condition, the inlet diffuser is normally replaced with a bell shaped inlet
to assure smooth flow into the compressor. Gas turbine cycle for jet propulsion differs from the shaft
power cycle because of the fact that the useful power output of the jet propulsion is produced ,wholly
or partially as a result of expansion of gas in the propelling nozzle. The principle of jet propulsion is
based on the Newton’s laws of motion.
The first law of thermodynamics for an open system determines the overall energy transfer .For
the understanding of the cycle it is better to analyze each component separately. For this purpose it is
better to use an air standard model.
The analysis can be of two types. In CAS (cold air analysis) one should assume the specific heat is
constant throughout the cycle. In WAS (warm air analysis) the value of specific heat will vary and
will be a function of temperature. To get a nice idea about the thermodynamics it is better to assume
each component of the engine as separate control volume and analysis can be done on that.
The paper presents an experimental procedure to evaluate the performance of a SR30 micro turbo
jet engine, for the laboratory education purpose. In order to fulfill this function one has to do both
theoretical analysis and experimental validation. Pourmovahed and Jeruzal[1], describes the
standard testing procedures and results based on experiments conducted in SR-30 micro Turbojet
engine, The same type of turbojet engine test rig is used for performing experiments in the project.
They also describes about the turbojet engine test rig emission analysis by using a proper exhaust gas
analyzer (Horiba MEXA 7100D).The engine had compatibility towards using both kerosene and
diesel as fuel. Liou and Leong [2], explains the testing procedure and results based on the
experiments conducted on a micro turbojet engine MW54. This turbojet engine test rig used two types
of fuel; with the ‘propane’ during its starting and ‘kerosene’ or jet A1 for the normal run. For the data
retrieval a LABVIEW based engine control interface was been incorporated. Witkowski, White and
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Transactions on Engineering and Sciences
Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
Duenas [3], describes the testing procedure of an SR-30 turbojet engine in three different temperature
starting condition. They also describe the calculation of the temperature profile across the combustion
chamber. By using a cutaway type SR-30, they found out the profile dimension of each component
and one dimensional calculation is made accordingly. Perez-Blanco [4], describes about the
experiments conducted on a ‘splitter blade compressor’ type SR-30 engine. It also describes the
calculation of various parameters and has shown its dependency on rotational velocity. Cohen,
Rogers and Saravanamuttoo [5]&Ganesan [6], explains the theoretical validation of a turbojet engine,
the basic thermodynamic cycle for a turbojet engine and the thermodynamic process happened on the
working fluid when it passes through each of the components. It also explains the performance
analysis of each components and has explains the dependency of one parameter on the other The
engine in which the experiment has to do is SR30 Minilab micro turbojet engine. A micro turbojet
engine is a scaled down version of a normal jet engine used in air craft propulsion and it mimics a real
engine during operation.
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The main components of Minilab systems are a Bell-mouthed intake at inlet, a conventional
radial compressor.With no splitter blades, a reverse annular flow type combustion chamber, an axial
turbine and a thrust nozzle at the exit. The operating range of RPM for the Minilab system lies in
between 40000 to 90000. The engine thrust at 90000 RPM for the Minilab system is around 250N.In
Minilab system up to an ideal rpm of 40000 compressed air is been used and after that Jet A-1 fuel is
been used.
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Fig 1.SR-30 Minilab micro turbojet engine test rig
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The Minilab SR-30 engine has a fuel system of a capacity 26.5 liters and an oil system of capacity
3.8 liters placed below the engine cabinet. The fuel which used for the experiment is Jet-A1.To
measure experimental parameter, the system is provided with proper sensors. A load cell is been
used to measure the ‘thrust’ directly. For temperature measurement K-type thermocouple is been
used. To retrieve the reading there are standard data retrieval systems and for safety purpose, digital
reading meters are placed in the panel for sudden indication of values. The SR-30 engines in which
the testing has been done had some limitation beyond an RPM of 80000. So the maximum RPM for
which the experiments has done is nearer to 80000. The data which could be taken directly from the
test rig are RPM, Compressor inlet and exit pressures and temperatures, Turbine inlet and exit
pressures and temperatures, Nozzle exit pressure, Fuel flow, Exhaust Gas Temperature and Thrust.
2. EXPERIMENTAL PROCEDURE
•
•
•
•
•
Following is the experimental procedure to do testing on the SR-30 Turbojet Minilab system
Check all the necessary safety precautions and ensure that the proper amount of air and fuel is in
the system. This can be check either by manual inspection or from the indicator gauge on the
front panel.
During the operation at any time, if the operator suspects that some problem is there in the
system, suddenly turn the fuel switch to off.
After all safety precaution when the Master switch turned on, the power comes to the system.
When the engine panel gives a ‘ready flag’ signal, one can start the engine
Up to an idle stable RPM the engine will be operated by an external compressed air supply (idle
RPM will be around 30000 to 40000).
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Transactions on Engineering and Sciences
Vol. 2, Issue 2, February 2014
•
•
ISSN: 2347-1964 Online 2347-1875 Print
Before starting fuel supply makes sure that the TIT (Turbine Inlet Temperature) is stable.
After the display shows a ‘RUN’ signal on the front panel, the engine can start operating my
moving the position of the throttle lever from a minimum to maximum position.
In order to take a particular reading, the position of lever should be in the same position till all the
quantities come to a stable value (temperature readings are very difficult to be getting in to stable
position).
The engine can stop by the fuel cut off and readings can take from the data retrieval system. One
can also note the reading by checking the gauges in the engine panel.
The retrieved data can be used for further calculations.
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Fig 1. Sensor positioning of SR-30 turbo jet engine test rig . (Courtesy to Turbine Technologies ltd.)
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3. PARAMETER TO BE MEASURED AND THEORETICAL FORMULA, FOR THE
COMPONENT PERFORMANCE CALCULATION
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A. Compressor
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The compressor constitutes of both impeller and diffuser
• Type - Centrifugal ( radial) type
• RPM range - 40000 to 90000
• Number of blades - 12
• Vtip at maximum RPM - 473 m/s
• Maximum pressure ratio - 3.4
• η isentropic - 75%
a)
The System Parameters Measured in Compressor
•
•
•
•
•
•
P1 - The compressor inlet pressure, Pa
P2 - The compressor outlet pressure, Pa
T1 - The compressor inlet temperature, K
T2 - The compressor outlet temperature, K
T2’ – The compressor outlet temperature at isentropic condition, K
RPM - Measuring the output voltage of a generator mounted on the compressor
b) Formula Used for Calculation
1. Specific work done by the compressor
Wcomp = h02 – h01, KJ/kg
The values of h1&h2 can be found out by using the air tables
2. To perform the system performance calculation of the engine at compressor inlet
i.
Check the pitot static recorded pressure -Direct reading from the system
ii.
Density ρ
=
P
, kg/m3
RT
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Transactions on Engineering and Sciences
Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
2P
iii.
Air velocity into the compressor, v =
iv.
v.
vi.
Volumetric flow rate of air into the compressor,
Q = A× v , m3/s
Mass flow rate of air into the compressor,
ṁ = ρ ×A× v, kg/s
Thrust generated in this location and operating condition, T = ṁ × v, N
vii.
Mach number of the flow at this location,
ρ
, m/s
M=
V
γ RT
3. Air fuel ratio = mass flow rate of air / mass flow rate of fuel
=
m& a
m& f
4. Isentropic Efficiency, η Compressor =
h2 − h1
,%
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Combustion Chamber
zType - Reverse flow annular type
c) Material - Inconel sheet
• Number of nozzles - 6
• Fuel pressure - 150 psi
• Theoretical efficiency of combustion - 99%
• Theoretical pressure loss - 5%
System parameters used in Combustion Chamber Calculation
•
•
•
•
•
•
P2 - The compressor outlet pressure , Pa
P3 - The turbine inlet pressure , Pa
T2 - The compressor outlet temperature, K
T3 - The turbine inlet temperature , K
Fuel flow rate , kg/m3
Thrust force, N
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a)
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b) Calculations for Combustion Chamber
i.
Specific energy added by the fuel,
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Qadd = h3 –h2, KJ/kg
ii.
iii.
TSFC = Weight of the fuel burned per hour/ Thrust force, kg/kg.hr
ߟ Combustion Chamber
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B.
h2' − h1
= Heat generated/ heat supplied
=
m& (hout − hin )
m& f × Cv
C. Turbine
•
•
•
•
•
Type - Single stage axial flow
Number of blade - 26 in stator, 21 in rotor
Maximum pressure ratio - 1.92
ߟisentropic - 95%
Material - Inconel 718 (vane guide ring),CMR 247 (turbine)
a) System parameters used
•
•
•
•
•
•
T3 - Turbine inlet temperature, K
T4 - Turbine exit temperature, K
T4’- Turbine exit temperature (isentropic condition) K
P3 - Turbine inlet pressure, Pa
P4 - Turbine exit pressure, Pa
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Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
b) Calculation for turbine
i. The specific work output, Wspecific = h4- h3 KJ/kg.
ii. The turbine pressure ratio can be found out by WC = WT
m& a C p (T2 − T1 ) = (m& a + m& f )C p (T3 − T4 )
iii. The isentropic efficiency η
=
h3 − h4
h3 − h4′
D. Nozzle
Type - Convergent type
Maximum EGT (Exhaust Gas Temperature) - 720 ºC
Maximum thrust -180 N
a)
System Parameters used in Nozzle Calculation
•
•
•
•
T4 - Turbine exit temperature , K
T5 - EGT , K
P4 - Turbine exit pressure , Pa
P5 - Nozzle exit pressure , Pa
b)
Calculation for Nozzle
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•
•
•
i. Check the pitot static recorded pressure from the engine test rig
P
, KJ/kg
RT
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ii. Density, ρ =
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ρ
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iii. Air velocity through the nozzle, v = 2P , m/s
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iv. Volumetric flow rate of air from nozzle, Q = A× V, m3/s
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v. Mass flow rate of air from the nozzle, ṁ = ρ ×A×V, kg/s
vi. Thrust generated at this location and operating condition,
T = ṁ ×V, N
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vii. Mach number of the flow at this location, M =
V
γ RT
4. RESULTS AND DISCUSSION
A.Theoretically calculated values based on catalog details
TABLE 1 theoretically calculated values of SR-30 turbojet engine parameter, based
on catalogue details
Parameters
RPM
Inlet Pressure, P1
Inlet Temperature, T1
Compressor Isentropic
Efficiency
Turbine isentropic
efficiency
Combustion chamber
efficiency
Combustion chamber
pressure loss
Compressor pressure ratio
Value
90000
0.10072134 MPa
301 K
70 %
T2 ’
426.97 K
31
95 %
99%
5%
3.4
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Transactions on Engineering and Sciences
Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
478.33 K
893.32 K
750.18 K
741.35 K
0.3253299 MPa
0.16938972 MPa
1.92
646.557 K
508.484 m/s
257.34 N
0.98
0.8391 /kg.h
T2
T3
T4
T4 ’
P2
P4
Turbine pressure ratio
T5
Exit velocity
Thrust
Exit Mach number
TSFC
B.Experimental Calculation
This sample calculation is based on the real time test data. The values of the readings are given
below. The barometric pressure is 14.61 PSI.
56000
73000
75700
78780
11.62
9.516
2.029
21.96
19.80
2
24.91
26.51
25.65
29.63
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P4 in
PSIG
0.630
P5 in
PSIG
0.21
s.o
P3 in
PSIG
6.574
pt
47600
P2 in
PSIG
7.31
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P1 in
PSIG
0.
095
0
.142
0.
318
0.359
0.388
5.134
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RPM
Pressure and fuel flow rate variations with RPM
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TABLE I.
7.110
7.263
Fuel flow
in GPH
2.2
1.42
2.91
3.40
4.21
4.09
4.31
4.46
4.79
T2 in
ºC
T3 in
ºC
T4 in
ºC
EGT
in ºC
Thrust
In lbs
47600
56000
28.3
28
121
129
607
620
534
539
528
531
1.43
9.86
73000
28.2
177
681
561
536
23.31
75700
78780
27.7
28
192
205
702
720
571
580
538
546
26.04
29.96
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T1 in
ºC
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TABLE II. Temperature and thrust variations with RPM
TABLE III. The calculation result for SR-30 turbojet based on Test Data Parameters for 73000
RPM
Parameter
RPM
Mass
flow
rate
at
compressor inlet
Air-fuel ratio
Compressor inlet velocity
Value
73000
0.22469 kg/s
63.44
60.69 m/s
58.12%
compressor
88.32%
turbine
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Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
Combustion
chamber
efficiency
Compressor pressure ratio
93.72%
Pressure
drop
across
combustion chamber
Turbine pressure ratio
Exit velocity of nozzle
Calculated thrust
6.2%
TSFC
Measured thrust
2.45
1.710
296.09m/s
117.1949 N
1.205kg fuel /kg.hr
thrust
104.06N
A. Component efficiencies and thermal efficiency
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As the RPM value changes from 47600 to 78780, the compressor isentropic efficiency varies from
38.63 % to 60.70 %, the turbine isentropic efficiency varies from 91.02% to 86.65 %. The average
efficiency of combustion chamber lies above 92%. The maximum thermal efficiency is 28.41%, and is
at the peak RPM.
52.58
73000
58.12
75700
78780
59.02
60.70
Thermal efficiency
in %
91.02
91.15
1
89.93
86.20
9
88.32
93.72
22.68
87.13
86.65
96.24
94.53
27.855
28.41
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38.63
56000
Turbine
isentropic
efficiency in %
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47600
Efficiency of
combustion
chamber in %
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Compressor
isentropic efficiency
in %
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RPM
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TABLE IV. The table shows the variation of component efficiencies and thermal efficiency
with respect to RPM
Fig 3. RPM vs. Compressor efficiency
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Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
From Fig.3 It can be noted that the compressor efficiency increases as RPM increases from 47600
to 75700. This is a typical character of a rotating machine especially compressor. When the system
attains a speed near to its maximum, the performance of the machine will be best.
Fig 4. RPM vs. Turbine isentropic efficiency
The combustion chamber efficiency vs. RPM
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Fig 5.
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From Fig.4, As the RPM increases turbine efficiency decreased from 91.02% to 86.65%. The
decrement in turbine efficiency is less than 5 %.
From Fig.5, Except for the second, for all other reading the combustion efficiency lies over
90%.The typical value of good combustion efficiencies are above 90%. Theoretically we are assuming
that the combustion efficiency is been 99% but in actual case this is not possible owing to various
losses. The average efficiency of nozzle and combustion chamber is well over 90%.So it is in par with
the theoretical values. The value of thermal efficiency is very poor at low RPM and as the engine RPM
increases, it get to a value of 28.41%, which is comparable with the standard thermal efficiency of a
turbo jet engine.From table III, it can be noted that the TIT increases as the RPM increases. The
maximum value of TIT is at the peak RPM. At 78780 the value of TIT is 720 ºC
B.Thrust and TSFC
TABLE VI. The variation between theoretical and actual thrust, TSFC with change in RPM
RPM
Thrust measured
using sensor in N
Thrust by theoretical
calculation in N
Actual TSFC
in
kg fuel /kg.hr thrust
47600
56000
73000
75700
6.3910
44.067
104.017
116.25
7.2402
48.3071
117.319
141.484
10.26
1.97
1.21
1.14
78780
133.75
148.38
1.07
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Transactions on Engineering and Sciences
Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
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Fig 7.Actual and theoretical thrust vs. RPM
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Fig 6. TSFC vs. RPM
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B. Mass flow rate of air and fuel and air-fuel ratio
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The air-fuel ratio, has an average value of 63.11 over the RPM range. Both air flow rate and fuel
flow rate increases as the RPM goes to higher values. The maximum value of air flow rate is 0.24886
kg/s and the maximum value of fuel flow rate is 0.004029kg/s .Both is at the maximum engine RPM.
Variation of air flow rate, fuel flow rate, air-fuel ratio with the change in RPM
Fuel flow
rate in
Kg/s
0.0018506
Air-Fuel
ratio
47600
Air flow
rate in
kg/s
0.12186
56000
0.14930
0.002447
60.99
73000
0.22469
0.003541
63.44
75700
0.23849
0.003751
63.56
78780
0.24886
0.004029
61.7616
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RPM
Fig 8.
65.84
The fuel flow rate vs. RPM
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Transactions on Engineering and Sciences
Vol. 2, Issue 2, February 2014
ISSN: 2347-1964 Online 2347-1875 Print
Fig 9. Air- fuel ratio vs. RPM
The Fig 9.shows that the air –fuel ratio is not changing much with respect to RPM and it
almost lies in 60 to 65 ranges
C. Performance analysis in compressor inlet
The mass flow rate of air and inlet Mach number increases slightly as the RPM get increased.
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TABLE VII. The variation of inlet air velocity, mass flow rate, Mach number with increment in
RPM.
Velocity of air into the
compressor. In m/s
Mass flow rate of air
into compressor in
kg/s
47600
56000
73000
75700
78780
33.43
40.78
60.69
64.39
66.86
0.12186
0.14930
0.22469
0.23850
0.24886
Mach number of
flow into
compressor
pt
s.o
RPM
ch
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0.096
0.1172
0.1744
0.185
0.192
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C. Pressure ratio and work done by the compressor and turbine
Compresso
r pressure
ratio
1.49
1.78
2.45
2.69
2.95
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TABLE IX. The variation of compressor and turbine pressure ratio with change in RPM
47600
56000
73000
75700
78750
Fig 10.
Turbine
pressure
ratio
1.39
1.45
1.71
1.82
1.88
Compresso
r work in
KJ
11.738
15.636
35.119
41.266
46.530
Turbine
work in
KJ
11.911
15.916
35.4034
41.659
46.833
Pressure ratio of compressor and turbine
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ISSN: 2347-1964 Online 2347-1875 Print
D. Pressure ratio needed for the given efficiency
From the analysis of slip factor, one can be able to calculate the theoretical pressure ratio needed
for the given efficiency.
TABLE X Actually obtained pressure ratio vs. Theoretically obtained pressure ratio.
RPM
Theoretical
tip velocity
in m/s
Pressure ratio
actually
obtained
Theoretically
obtained pressure
ratio for the same
efficiency.
47600
56000
73000
75700
78780
254.08
298.92
389.67
404.08
420.53
1.49
1.78
2.45
2.69
2.95
1.266
1.54
2.16
2.3130
2.51
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5. CONCLUSION
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Sensor type
Setra differential
Setra Model 209
Setra Model 209
Setra Model 209
Setra Model 209
Pulse counting
Futek LLB400
K-Type thermocouple
K-Type thermocouple
K-Type thermocouple
K-Type thermocouple
K-Type thermocouple
s.o
Probe
Pitot/static
Stagnation
Static
Stagnation
Stagnation
Tachometer generator
Button type load cell
Static
Stagnation
Stagnation
Stagnation
Stagnation
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System parameter
P1, Compressor inlet
P2, Compressor outlet pressure
P3, Turbine inlet
P4, Turbine exit
P5, Exhaust gas
RPM
Thrust
T1, Compressor inlet
T2, Compressor exit
T3,Turbine inlet
T4,Turbine exit
T5,Exhaust gas temperature
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Sensor details of SR-30 turbojet engine test rig
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TABLE XII.
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The experiments which were conducted in the micro turbojet engine setup allow one to reveal the
thermodynamics behind the operation of a turbojet engine. The results obtained through the
experiments are in par with the standard values and existing theories. By checking the result, one can
predict that the machines like compressor and turbine is relying heavily on their rotational speed. The
components will give a good performance as the engine speed is approaching near the peak.
REFERENCES
A.Pourmovahed, C.M.Jeruzal, K.D.Brinker; Development of a Jet Engine experiment for the
Energy System Laboratory. Proceedings of the 2003 ASME International Mechanical Engineering
Congress, Washington D.C 2003 .IMECE 2003-43638
[2] W.William Liou, Chin Hoong Leong; Gas Turbine Engine Testing Education at Western Michigan
University. 45th AIAA Aerospace Sciences Meeting and Exhibit, Nevada 2007. AIAA-2007-703
[1]
Witkowski, T. White, S.Oritz Duenas; Characterizing the Performance of the SR-30 Turbojet
Engine, Proceedings of the 2003 American Society for Engineering Education Annual Conference
and Exposition. 2003-1133
[4] H.Perez-Blanco; Activities around the SR-30 MINILAB at PSU. Proceedings of the 2003 American
Society for Engineering Education Annual Conference and Exposition.2003-1133
[5] . CFC Rogers, H Cohen, HIH Saravanamuttoo; Gas Turbine Theory, Fifth edition. Prentice Hall
2001.
[6] V Ganesan; Gas Turbines, Third edition. McGraw Hill 2010.
[3]
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