Turbine Technologies, Ltd. SR-30 Turbojet Engine Analysis
University of Minnesota Duluth,
Department of Mechanical and Industrial Engineering
November 19, 2006
2/6/2016
Drew Hauck
_______________________
Craig Bourassa
_______________________
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Table of Contents:
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Objective:
The objective of the turbojet analysis is to calculate the actual performance of the engine
based on experimental measurements and compare it with theoretical results.
Summary:
Analysis of the SR-30 Turbojet engine was performed after the data was gathered during
a run of the Turbine Technologies, Ltd. Model LX4000 Gas Turbine Laboratory
apparatus. The pressures and temperatures along with fuel flow rate, thrust force and
engine revolutions per minute (RPM) were recorded by the data acquisition system
during the experiment. Initial review of the data indicated good results but the analysis
revealed incorrect data and or assumptions. *****************
Theory and Analysis:
The SR-30 Turbojet Engine is theoretically modeled by the Brayton cycle. The Brayton
cycle is made up of 4 processes:

Isentropic compression in the compressor

Constant-pressure heat addition in the combustor

Isentropic expansion in the turbine

Constant-pressure heat rejection across the nozzle
Figure 1: The Brayton Cycle [1]
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Figure 2: T-s Diagram of ideal Brayton Cycle [1]
Fresh air is drawn in to the compressor through the nozzle. The compressor increases the
pressure of the air before it is sent into the combustion chamber. In the combustor, the
air is mixed with the injected fuel and burned at constant pressure. The high pressure and
temperature air is forced across the turbine blades which spin the turbine. The air exiting
the turbine is accelerated across the nozzle to produce thrust. This analysis has a few
overriding assumptions:
1. One dimensional analysis
2. Working fluid is air treated as an ideal gas
3. All processes are internally reversible
4. Combustion is modeled as heat addition from external source
5. Constant specific heats
6. Steady state
7. Steady flow
The stagnation temperature for isentropic compression is calculated as Equation 1.
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kair 1
 p 02 
T02s  T01 

 p 01 
kair
Equation 1: Stagnation Temperature for Isentropic Compression
Equation 1 is used for each location using subscript 02 as the exit and subscript 01 as the
inlet. The density anywhere in the system can be calculated from the ideal gas law using
the gas constant for air (Rair = 287J/kg K).
 1 
p1
Rair T1
Equation 2: Density from Ideal Gas Law
The entropy in the system is calculated using Equation 3. This is assuming the flow is
classified as Rayleigh flow. Rayleigh flow is steady, one dimensional flow of an ideal
gas having constant specific heats passing through a constant area duct with heat transfer
and negligible friction.
T 
p 
s c cp  T  p  s ref  cp ln 
 Rair ln


 Tref 
 pref 


Equation 3: Entropy in the System
Compressor:
The basic operation of the compressor consists of a stationary
casing and a rotating impeller producing high velocity air.
The compressor rotates at a very high rate imparting a high
speed to the outer edge of the vanes. The high speeds present
require high strength materials to be used in the
Figure 3: Engine inlet and
Compressor [5]
construction. The compressor has a diffuser section which
decelerates the air thus creating high pressure. The compressor contains relatively low
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temperatures and can be considered to have a constant heat capacity cp=1.005 kJ/kg K
and ratio of specific heat capacities k=1.4. The compressor is assumed to be isentropic,
meaning constant entropy from the compressor inlet to compressor outlet / turbine inlet.
Compressor work can be calculated from equation 1.


W'c  m'air h02  h01
Equation 4: Compressor work
The efficiency of the compressor is calculated from the ideal isentropic temperature and
the actual temperature at the stagnation state.
 compressor 
T01  T02s
T01  T02
Equation 5: Compressor efficiency
Combustor:
The SR-30 Turbojet engine uses a reverse flow annular combustion chamber. The
annular type combustor is used because of the low frontal area and weight for the given
volume. The reverse type allows the turbine to be radially inboard of the combustor
permitting a shorter engine. The air entering into the combustor is at a high pressure
from the compressor. The high pressure air mixes with the fuel as it is sprayed in. The
fuel is assumed to burn at constant pressure and the resulting high temperature, high
pressure air is exited to the turbine. The combustion process has a high air to fuel ratio
because the excess air is used to cool the turbine. The air to fuel ratio in the combustor
can be calculated as:
AF 
m'air
m'fuel
Equation 6: Air to Fuel Ratio
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The total energy produced from the combustion process can be calculated using the lower
heating value of the fuel and the mass flow rate of the fuel.
Q'in  m'fuel LHVfuel
Equation 7: Combustion Energy From Fuel
Turbine:
Like the compressor, the turbine has a very high rotational rate. The added effect of the
high temperature from the combustion gasses make the strength of the blades crucial.
The strength of the turbine blades is the limiting factor to how fast the engine can spin.
The axial flow type turbine is used in the SR-30 to convert the high pressure gasses
exiting the combustor to rotational energy. The axial flow turbine is used in aircraft
because of the lower frontal area required for a given mass flow and pressure ratio. The
turbine has high temperatures and the heat capacity is assumed to be cp=1.148 kJ/kg K
and the ratio of specific heat capacities is k = 1.333. Turbine efficiency can be calculated
from the ideal isentropic temperature and the actual temperature at the stagnation state.
 turbine 
T03  T04
T03  T04s
Equation 8: Turbine Efficiency
Turbine work is a function of the total mass flow rate, specific heat and the change in
temperature across the turbine.


W't  m'tot cpp T03  T04
Equation 9: Turbine Work
Nozzle:
The purpose of the nozzle in an aircraft turbine is to accelerate the exhaust gasses to
produce thrust. In subsonic flow, i.e. Ma < 1, a converging nozzle is employed to
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accelerate the air. For supersonic flow, i.e. Ma > 1, a diverging nozzle is used to
accelerate the air. The exit velocity in the SR-30 Turbojet is less than the speed of sound
so a converging nozzle is used. Like the turbine, the nozzle has high temperatures also.
The specific heat capacity and the ratio is the same at the turbine. Nozzle efficiency is
calculated like the rest of the components with the ideal isentropic temperature and the
actual temperature at the stagnation state.
 nozzle 
T04  T5
T04  T5s
Equation 10: Nozzle Efficiency
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