Design of Turbojet Gas Turbine Engine
Design of Turbojet Gas Turbine Engine
Design of Turbojet Gas Turbine Engine
FACULTY OF ENGINEERING
DEPARTMENT OF MECHANICAL ENGINEERING
PROJECT - 3
(Final Report)
Özgür ATASEVER
1171601906
PROJECT ADVISOR:
Prof.Dr. Kamil KAHVECİ
May 2022
EDIRNE
TRAKYA UNIVERSITY
FACULTY OF ENGINEERING
DEPARTMENT OF MECHANICAL ENGINEERING
PROJECT - 3
(Final Report)
Özgür ATASEVER
1171601906
PROJECT ADVISOR:
Prof.Dr. Kamil KAHVECİ
May 2022
EDIRNE
i
ABSTRACT
Aircrafts used in the field of transportation as the fastest and most reliable mode of
transport since the beginning of twentieth century’s, which are widely used in many areas
not only in the field of transportation, but also health, military and firefighting.
In this study, theoretical researches has been made about Gas Turbine Engines, a turbojet
gas turbine engine has been analyzed theoretically and computationally in GasTurb
program, and a jet engine has been simply designed and presented in SolidWorks
program. CAD drawings of the parts are attached.
Keywords:
ii
CONTENTS
ABSTRACT...................................................................................................................... ii
CONTENTS..................................................................................................................... iii
LIST OF FIGURES ......................................................................................................... iv
LIST OF TABLES ........................................................................................................... vi
SYMBOLS...................................................................................................................... vii
1. INTRODUCTION ........................................................................................................ 1
2. BACKGROUND ......................................................................................................... 3
2.1. Project Background ............................................................................................... 3
2.2. History of Jet Engine ............................................................................................. 3
2.3. Application of Jet Engine ...................................................................................... 6
2.4. Working Principle of Jet Engine ........................................................................... 6
2.5. Components of Jet Engine ..................................................................................... 7
2.5.1. Fan .................................................................................................................. 7
2.5.2. Compressor ..................................................................................................... 8
2.5.3. Combustor ...................................................................................................... 9
2.5.4. Turbine ......................................................................................................... 10
2.5.5. Nozzle ........................................................................................................... 11
2.6. Types of Jet Engine ............................................................................................. 11
2.6.1. Turbojets ....................................................................................................... 11
2.6.2. Turboprops ................................................................................................... 12
2.6.3. Turbofans ...................................................................................................... 13
2.6.4. Turboshafts ................................................................................................... 13
2.6.5. Ramjets ......................................................................................................... 14
2.7. Reaction Engines ................................................................................................. 15
2.8. Cycle Overview ................................................................................................... 16
2.9. Compressors and Turbines .................................................................................. 20
2.10. Performance prediction for compressors and turbines ...................................... 20
2.11. Centrifugal compressor ..................................................................................... 22
2.12. Combustion champer ........................................................................................ 23
2.13. Gas Turbine Systems ......................................................................................... 26
2.13.1. Open Cycle ................................................................................................. 26
2.13.2. Power & district cooling application (Cogeneration) ................................. 27
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2.13.3. Combined heat & power application (Cogeneration) ................................. 28
2.14. GasTurb .............................................................................................................. 29
2.14.1. Cycle Design .............................................................................................. 29
2.14.1.1. Design Point Calculation .................................................................... 29
2.14.1.2. Parametric Studies .............................................................................. 30
2.14.1.3. Optimization ....................................................................................... 31
2.14.1.4. Secondary Air System ........................................................................ 31
2.14.1.5. Termodynamic Stations ...................................................................... 32
2.15. SolidWorks ........................................................................................................ 33
3. METHOD .................................................................................................................. 34
3.1. Given Data............................................................................................................ 34
3.2. Obtained Data from GasTurb .............................................................................. 37
4. MY JET ENGINE DESIGN WITH SOLIDWORKS ............................................... 42
5. CONCLUSIONS AND RECOMMENDATIONS .................................................... 43
6. REFERENCES ........................................................................................................... 44
APPENDIX: CAD DRAWINGS OF PARTS ............................................................... 45
RESUME ....................................................................................................................... 49
LIST OF FIGURES
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Figure 16. Nozzle ........................................................................................................... 11
Figure 17. Turbojet engine ............................................................................................. 12
Figure 18. Turboprop engine ......................................................................................... 12
Figure 19. Turbofan engine ........................................................................................... 13
Figure 20. Turboshaft engine ......................................................................................... 14
Figure 21. Ramjets engine ............................................................................................. 14
Figure 22. Propulsive efficiency as a function of Ve / V0 ............................................... 16
Figure 23. A schematic view of a turbojet ..................................................................... 16
Figure 24. TS diagram of the ideal Brayton Cycle for Turbojet .................................... 17
Figure 25. TS diagram of actual Brayton Cycle ............................................................ 19
Figure 26. Velocity triangles for axial turbine. W is the relative stream velocity ......... 22
Figure 27. Velocity triangle for centrifugal compressor with radial vanes. (β2 = 90◦.) .. 23
Figure 28. Fuel-air ratio vs velocity................................................................................ 25
Figure 29. Cross-section of can, annular and can-annular type combustion chamber. .. 25
Figure 30. Open Cycle System ...................................................................................... 26
Figure 31. Gas District Cooling ..................................................................................... 27
Figure 32. Heat Recovery Steam Generator .................................................................. 28
Figure 33. GasTurb main window ................................................................................. 29
Figure 34. Design Point Calculation .............................................................................. 30
Figure 35. Parametric Studies ........................................................................................ 30
Figure 36. Optimization ................................................................................................. 31
Figure 37. Secondary Air System .................................................................................. 32
Figure 38. Termodynamic Stations ................................................................................ 32
Figure 39. SolidWorks ................................................................................................... 33
Figure 40. Combustion Temperature Distribution Coefficient ...................................... 35
Figure 41. Design Point Input Section ........................................................................... 37
Figure 42. Design Point Summary ................................................................................. 37
Figure 43. Design Point Stations ................................................................................... 38
Figure 44. T-S Diagram ................................................................................................. 38
Figure 45. H-S Diagram ................................................................................................. 39
Figure 46. P-V Diagram ................................................................................................. 39
Figure 47. External Factor Map ..................................................................................... 40
Figure 48. Parameter Selection ...................................................................................... 40
Figure 49. Efficiency Graph .......................................................................................... 41
Figure 50. Variation of net thrust with the altitude at a Mach number of 0.8 ................ 41
v
Figure 51. My jet engine design with SolidWorks ........................................................ 42
Figure 52. My jet engine design with SolidWorks (Transparent section view) ............. 42
LIST OF TABLE
vi
SYMBOLS
Symbol Name
Cp specific heat capacity at constant pressure
Cv specific heat capacity at constant volume
EW work
eW specific work
F force
h specific enthalpy
h0 specific stagnation enthalpy
LHV fuel heating value
ṁ mass flow rate
P static pressure
P0 stagnation pressure
Q heat
q specific heat
R gas constant
r radius
s specific entropy
T static temperature
T0 stagnation temperature
U blade velocity
V absolute flow velocity
W relative flow velocity
α absolute flow angle
β relative flow angle
γ ratio of specific heats
η efficiency
π pressure ratio
ρ density
τ temperature ratio
ω angular velocity
vii
1. INTRODUCTION
The gas turbine engine is a machine, which work according to the thermodynamic
Brayton Cycle by harnessing energy from a working fluid and converting the energy into
a useable form. Various types of gas turbines are designed to perform a range of tasks,
but all operate on similar principles. Air enters the engine, is compressed, mixed with
fuel, combusted, and then expanded through a rotating turbine. Due to the extreme
temperatures and high rotational speeds experienced by engine components, design and
construction of a gas turbine demands accuracy, informed material selection, knowledge
of thermodynamics, and the ability to model and machine metal components. From fine
tolerance in space to resilience to high temperatures and stress, the jet engine has gone
through a revolution over the years, with great improvements in performance, efficiency,
and reliability. The most known jet engines are the turbojet engine, the turboprop engine,
the turbofan engine, the turboshaft, and the ramjet engine. The major principle in all these
engines are the same and they work according to similar concepts as the internal
combustion engine:
i. Suck: The engine sucks in a large volume of air through the fan and compressor
stages. A typical commercial jet engine takes in 1.2 tons of air per second during
takeoff. The mechanism by which a jet engine sucks in the air is largely a part of the
compression stage. Some engines have an additional fan that is not part of the
compressor to draw additional air into the system. This part is on left side in figure 1.
1
ii. Squeeze: Aside from drawing air into the engine, the compressor also pressurizes the
air and delivers it to the combustion chamber. The compressor is shown in figure 1
just to the left of the fire in the combustion chamber and to the right of the fan. The
compression fans are driven from the turbine by a shaft. Compressors can achieve
compression ratios more than 40:1, which means that the pressure of the air at the end
of the compressor is over 40 times that of the air that enters the compressor.
Figure 2 shows the green fans that compose the compressor gradually get smaller and
smaller, as does the cavity through which the air must travel. The air must continue
moving to the right, toward the combustion chambers of the engine, since the fans are
spinning and pushing the air in that direction. The result is a given amount of air moving
from a larger space to a smaller one, and thus increasing in pressure.
iii. Bang: In combustion chamber fuel is mixed with air to produce the bang, which is
responsible for the expansion that forces the air into the turbine. Inside the typical
commercial jet engine, the fuel burns in the combustion chamber at up to 2000 degrees
Celsius. The temperature at which metals in this part of the engine start to melt is
1300 degrees Celsius, so advanced cooling techniques must be used.
The combustion chamber has the difficult task of burning large quantities of fuel, supplied
through fuel spray nozzles, with extensive volumes of air, supplied by the compressor,
and releasing the resulting heat in such a manner that the air is expanded and accelerated
to give a smooth stream of uniformly heated gas. This task must be accomplished with
the minimum loss in pressure and with the maximum heat release within the limited space
available. The amount of fuel added to the air will depend upon the temperature rise
required.
2
iv. Blow: The fourth part is focused on the outlet of the engine, the reaction of the
expanded gas the mixture of fuel and air being forced through the turbine, drives the
fan and compressor and blows out of the exhaust nozzle providing the thrust.
Thus, the turbine has the task of providing power to drive the compressor and accessories,
by extracting energy from the hot gases released from the combustion system and
expanding them to a lower pressure and temperature. To produce the driving torque, the
turbine may consist of several stages, each employing one row of moving blades and one
row of stationary guide vanes to direct the air as desired onto the blades. The number of
stages depends on the relationship between the power required from the gas flow, the
rotational speed at which it must be produced, and the diameter of turbine permitted.
2. BACKGROUND
In this section history of jet engines, applications, working principle and types of jet
engines are discussed.
The basic principle used in jet engines has been known for a long time. It dates to around
150 BC when the principle was used in the Aeolipile as shown in figure 3, which is a
simple construction using a radial steam turbine. The steam exits through a nozzle
creating a spinning motion of a ball. All according to Newton’s third law.
The key to a practical jet engine was the gas turbine, extracting power from the engine
itself to drive the compressor. The gas turbine was not a new idea: the patent for a
stationary turbine was granted to John Barber in England in 1791. The first gas turbine to
successfully run self-sustaining was built in 1903 by Norwegian engineer Ægidius Elling.
The interest continued during the 1800s. But it wasn’t until Sir Frank Whittle of the Royal
Air Force in the 1930s made the first patent for the jet engine and showed the possibilities
through reliable energy conversion. He made the first static test in 1937. Two years later,
in 1939, it was a German physicist named Hans von Ohain who made the first jet-
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powered-flight and demonstrated the possibilities of the jet engines. The ideas came about
improving the propeller driven aircrafts of the time, where the main problem was the
speed of the aircraft. The aircraft of the time were closing in on the speed of sound, and
sometimes getting too close, which would result in shock waves being created, causing
the propeller to shatter.
The jet engine allowed a continuous combustion and airflow. It was a big change from
the piston engines dominating the industry. At the time, the greatest struggle the engineers
had was to create a material that could withstand the temperatures generated in the
combustion chamber, since it would often lead to the turbines melting. The development
of the jet engine took off during World War II and performance was quickly raised
because of the efforts made to try to get any advantage possible. Thus, paving the way
for the modern jet engines.
Following the end of the war the German jet aircraft and jet engines were extensively
studied by the victorious allies and contributed to work on early Soviet and US jet
fighters. The legacy of the axial-flow engine is seen in the fact that practically all jet
engines on fixedwing aircraft have had some inspiration from this design. By the 1950s
the jet engine was almost universal in combat aircraft, except for cargo, liaison and other
specialty types. The efficiency of turbojet engines was still rather worse than piston
engines, but by the 1970s, with the advent of high-bypass turbofan jet engines fuel
efficiency was about the same as the best piston and propeller engines.
4
Figure 4. Albert Fonó's ramjet-cannonball from 1915
Figure 5. Heinkel He 178, the world's first aircraft to fly purely on turbojet power
5
2.3. Application of Jet Engine
Jet engines power jet aircraft, cruise missiles and unmanned aerial vehicles. In the form
of rocket engine, jet engine power fireworks, model rocketry, spaceflight, and military
missiles. Jet engine designs are frequently modified for non-aircraft applications, as
industrial gas turbines or marine powerplants. These are used in electrical power
generation, for powering water, natural gas, or oil pumps, and providing propulsion for
ships and locomotives. Industrial gas turbines can create up to 50,000 shaft horsepower.
Jet engines are also sometimes developed into, or share certain components such as
engine cores, with turboshaft and turboprop engines, which are forms of gas turbine
engines that are typically used to power helicopters and some propeller-driven aircraft.
All jet engines work on the same principle. The engine sucks air in at the front with a fan.
A compressor made with many blades attached to a shaft spin at high speed and raises the
pressure of the air by compress or squeeze the air. The compressed air is then sprayed
with fuel and an electric spark lights the mixture. The burning gases expand and blast out
through the nozzle, at the back of the engine. The engine and the aircraft get thrust
forward as the jets of gas shoot backward. As the hot air is going to the nozzle, it passes
through another group of blades called the turbine. The turbine is attached to the same
shaft as the compressor. Spinning the turbine causes the compressor to spin. Figure 9
shows the air flow through the engine.
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Figure 9. Air flow through the Jet engine
2.5.1. Fan
The fan is the first component of the engine which spins and draws a large quantity of air.
The fan is mainly made of titanium. The air gets divided into two halves. One half goes
through the core of the engine where it passes through the compressor, combustor and
turbine. The other half goes from the outer duct of the engine. This is known as the
‘bypass’ duct which covers the compressor, combustor and turbine of engine. After
flowing through the bypass duct, this cold air mixes with the hot air from the core of the
engine in the nozzle and produces the thrust for the plane.
7
Figure 11. Fan of jet engine
2.5.2. Compressor
After the fan, the air enters the compressor. Every jet engine has a compressor to increase
the pressure of the air. Jet engines mainly have only two types of compressors, axial
compressor and centrifugal compressor. In the most common jet engines i.e. two spool
turbofan engines, there are two compressors present. One is low pressure compressor,
which is just beside the Fan and the other is a high pressure compressor which is present
after the low pressure compressor. A high pressure compressor and high pressure turbine
are connected with one shaft, whereas a low pressure turbine, low pressure compressor
and fan are connected to another shaft in a two spool turbofan.
8
Figure 13. Compressor map
The compressor’s performance is decided by the compressor map on which it works. The
compressor map has a mass flow axis as x-axis and pressure ratio as y-axis. The left hand
boundary of the map is known as the surge line and the right hand boundary is the choke
line. There are concentric circles in the graph which are the efficiency islands and the
lines intersecting these islands are turbo speed lines, as shown in the Figure 13.
2.5.3. Combustor
In the combustor the air is mixed with fuel and then ignited. There are as many as 20
nozzles to spray fuel into the airstream. The mixture of air and fuel catches fire. This
provides a high temperature, high-energy airflow. The fuel burns with the oxygen in the
compressed air, producing hot expanding gases. The inside of the combustor is often
made of ceramic materials to provide a heat-resistant chamber. The heat can reach 2700°.
9
Figure 14. Combustion chamber
2.5.4. Turbine
The high-energy airflow coming out of the combustor goes into the turbine, causing the
turbine blades to rotate. The turbines are linked by a shaft to turn the blades in the
compressor and to spin the intake fan at the front. This rotation takes some energy from
the high energy flow that is used to drive the fan and the compressor. The gases produced
in the combustion chamber move through the turbine and spin its blades. The turbines of
the jet spin around thousands of times. They are fixed on shafts which have several sets
of ball-bearing in between them.
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2.5.5. Nozzle
The nozzle is the exhaust duct of the engine. This is the engine part which produces the
thrust for the plane. The energy depleted airflow that passed the turbine, in addition to the
colder air that bypassed the engine core, produces a force when exiting the nozzle that
acts to propel the engine, and therefore the airplane, forward. The combination of the hot
air and cold air are expelled and produce an exhaust, which causes a forward thrust.
The velocity of the air entering the nozzle is low, about Mach 0.4, a prerequisite for
minimizing pressure losses in the duct leading to the nozzle. The temperature entering the
nozzle may be as low as sea level ambient for a fan nozzle in the cold air at cruise
altitudes. It may be as high as the 1000 K exhaust gas temperature for a supersonic
afterburning engine or 2200K with afterburner lit. The pressure entering the nozzle may
vary from 1.5 times the pressure outside the nozzle, for a single stage fan, to 30 times for
the fastest manned aircraft at mach 3.
2.6.1. Turbojets
The concept of the turbojet aircraft engine entails taking air in from the engine’s rear side
and then compressing it in the compressor. In addition, fuel is added to the combustion
chamber and burned to raise the fluid mixture temperature to about 1000 degrees.
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The hot air that is produced is then pushed through a turbine that rotates the compressor.
The pressure at the discharge of the turbine is generally almost twice the pressure in the
atmosphere. However, that depends on the efficiency level of an aircraft engine. The
excessive pressure then moves to the nozzle that then generates gas streams, which are
responsible for creating a thrust.
2.6.2. Turboprops
A turboprop engine is basically a gas generator which drives a propeller. Hot gases that
are produced by gas generator makes the turbine at the back turn and that turns a shaft
which is connected to the propeller. Turboprop engines are efficient for low-speed flights
and short-field takeoff. When speed of flight approaches the speed of sound, turboprop
engines lose their efficiency since there is a loss of aerodynamic efficiency of propeller.
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2.6.3. Turbofans
A turbofan engine has a large fan at the front, which sucks in air. Most of the air flows
around the outside of the engine, making it quieter and giving more thrust at low speeds.
In a turbojet all the air entering the intake passes through the gas generator, which is
composed of the compressor, combustion chamber, and turbine. In a turbofan engine only
a portion of the incoming air goes into the combustion chamber. The remainder passes
through a fan, or low-pressure compressor, and is ejected directly as a "cold" jet or mixed
with the gas-generator exhaust to produce a "hot" jet. The objective of this sort of bypass
system is to increase thrust without increasing fuel consumption. It achieves this by
increasing the total air-mass flow and reducing the velocity within the same total energy
supply.
2.6.4. Turboshafts
This is another form of gas-turbine engine that operates much like a turboprop system.
The difference between them is that in turboshaft engines, power is delivered to a shaft
instead of the propeller. This shaft is generally gives power to a helicopter rotor. Like for
the other jet engine types, all steps in the gas generator are applied to the air and high-
energy gases are produced. However, on a turboshaft engine most of this energy is utilized
to drive the turbine instead of producing thrust.
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Figure 20. Turboshaft engine
2.6.5. Ramjets
The ramjet is the simplest jet engine and has no moving parts. The speed of the jet "rams"
or forces air into the engine. It is essentially a turbojet in which rotating machinery has
been omitted. Its application is restricted by the fact that its compression ratio depends
wholly on forward speed.
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2.7. Reaction Engines
In order to provide thrust all aircraft engines work by imparting rearward momentum to
one or more streams of gas, this gas is the reaction mass and such engines are known
collectively as reaction engines. From Newtons laws of motion we know that force is
equal to change in momentum, so net thrust, Fn for a jet engine can be written as follows,
know as the general thrust equation:
V0 is the free stream velocity of the air coming in to the engine, same
as the true airspeed of the aircraft
Since ṁair >> ṁfuel we can make the simplification that the mass flow rate entering the
engine is the same as the mass flow rate exiting the engine, and from there we get a simple
expression for the propulsive efficiency, ηp, which is the ratio between the work done on
the aircraft compared to the kinetic energy imparted to the air stream flowing through the
engine:
(2)
From this the conclusion can be made that the highest efficiency is achieved when the
velocity of the exhaust is nearly the same as that of the aircraft, shown in fig. 2. Using a
larger mass flow rate instead of higher exhaust velocities is therefor preferable.
Depending on the range of speeds an aircraft is designed to operate in the preferred engine
type varies, with slow aircraft the choice is propellers, followed by high-bypass turbofan
and then progressively lower bypass ratios as the speed increases.
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Figure 22. Propulsive efficiency as a function of Ve / V0
The turbojet engine is a heat engine, which works by converting heat to useful mechanical
work. In this case, the heat from combustion is used to propel the air rearwards. With the
help of a thermodynamic cycle an ideal case of the jet engine’s operation can be shown.
This is a good way to calculate the overall efficiency of the engine.
The turbojet engine consists of five main regions. Diffuser, compressor, combustion
chamber, turbine and nozzle. There are also numbered station which are used for
describing the state of the flow at different points in the engine. A schematic view is
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shown in figure 23. The zeroth station is far enough up stream before the intake that
ambient conditions apply. The region between stage one and two is the diffuser, where
the stream is slowed down, and the pressure rises. The region between 2 and 3 is the
compressor, where energy is added to the flow, idealized as an adiabatic process where
the pressure and the temperature increases and the volume decreases. Between station 3
and 4 the combustion takes place, heat is added and the volume and the entropy increases
and the temperature reaches its peak. Between station 4 and 5 is the turbine, where the
pressure and the temperature decreases and the volume increases while the air flows
through the turbine, converting heat to mechanical work. Finally, between 5 and 6, the air
goes through a nozzle back to ambient pressure, while accelerating.
The ideal open Brayton cycle consists of 4 processes, isentropic compression, isobaric
heat addition and isentropic expansion and then an isobaric heat rejection between exit
and inlet. We can see these relationships:
Work done by compressor:
EW,c = ṁa (h03− h02) (3)
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Work done on turbine:
Using conservation of energy, assuming that the change in potential energy due to
elevation is negligible and assuming conservation of mass (neglecting fuel flow since ṁair
>> ṁfuel )
) (7)
Using the thrust equation from the previous section, eq 1, the kinetic energy term can be
written as
(8)
Substituting back into eq 7, and assuming eW,in − eW,out = 0, since in the ideal case no work
is done by the shaft, gives an expression for thrust
(9)
Define the over all efficiency of the engine, the ratio of work done by the exiting stream
to the rate of heat added
(10)
18
(11)
Both heat in and heat out are constant pressure processes, approximating constant specific
heat Cp = Cp,avg, equation 11 becomes
(12)
Compression and expansion is assumed to be isentropic and we assume polytropic gas.
And P03 = P04 and P06 = P00 so
(13)
where πo = P03/P00 = P04/P06 is the overall pressure ratio of the engine. As fuel efficiency
is directly related to thermal efficiency, as high a pressure ratio, and temperature ratio as
possible is desired.
In the actual cycle there are a number of differences with regard to the ideal case.
Compression and expansion is not isentropic, and there are pressure losses along the
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whole path that the air takes in the engine. The isentropic efficiency for the compressor
and turbine is an expression for the differences between the ideal case and the actual case,
a meassure of how much of the theoretically available heat that is converted to work on
the rotor in the case of the turbine, and for compressors how much of the mechanial work
done by the rotor leads to increase in energy of the gas.
(14)
(15)
There are mainly two types of compressors and turbines used in jet engines and power
turbines: axial and radial/centrifugal. Both are of the continuous flow variety, where a
rotating mechanical part exchanges energy on a continuous flow of air. Both are
composed of two main parts, a rotating and a static part. The rotating part, rotor, transfers
kinetic energy to/from the fluid. In the the static part, stator, the kinetic energy is
converted to pressure by redirecting the flow and by increasing the flow area to slow the
fluid down. Or vice versa, to convert pressure to kinetic energy. Each pair of rotor and
stator is called a stage and are compounded in order to achieve greater pressure
differentials. In axial compressors, the pressure rise per stage is usually in the range 1.1:1
to 1.4:1, whereas centrifugal regularly operate around 3:1 and in extreme cases up to 12:1.
The stages are usually compounded and some designs use a mixture of radial and axial.
In high performance applications, such as in modern aircraft engines, many stages are
used to achieve total pressure ratios of up to 40:1. For turbines the axial type is almost
exclusively used as the pressure differential needed is lower, and isentropic efficiency is
more important.
Using a model of simple one dimensional flow, a way to describe the gas stream is to use
three basic equations. These are derived from conservation of mass, conservation of
momentum and conservation of energy.
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The mass of the gas is conserved, with the fluid seen as a continuum this can be
formulated as:
ṁ = ρAV (16)
where
(17)
(18)
the product of the torque and the angular velocity is the rate of energy transfer
(19)
(20)
where U1 and U2 is the tangential velocity of the rotor, blade velocity, at respective radii.
Note the sign, by convention work done by the fluid is defined as positive, so for a
compressor this expression would be negative.
In the case of an axial rotor, the blade velocity is constant from inlet to outlet, U1 = U2, so
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(21)
And for the case of a centrifugal compressor with no pre-whirl, the incoming flow has no
tangential component, vθ,1 = 0, so
(22)
Figure 26. Velocity triangles for axial turbine. W is the relative stream velocity
Centrifugal compressors are often used in small turbines both in aircraft and industrial
uses. One reason is that, compared to axial flow compressors, they offer higher pressure
ratios per stage. Centrifugal compressors also provide a wider range of operation, offering
better off-design performance. The drawback is lower isentropic efficiencies than axial
compressors, with higher fluid losses. And a larger frontal area, which can be undesirable
for aircraft application. Since the same pressure rise can be accomplished with fewer
stages centrifugal compressors are the choice for compact units and the larger operational
span make them appropriate for applications where adjustable output and robustness is
priority.
The main parts of the centrifugal compressor are the rotor and stator, often called impeller
and diffuser respectively. Fluid flows axially into the center of the impeller and is then
turned radial and flung outwards, kinetic energy being imparted to the flow. In the diffuser
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the cross sectional area of the flow is increased, decreasing the velocity and increasing
pressure. The compressor used for this project is the centrifugal compressor.
Figure 27. Velocity triangle for centrifugal compressor with radial vanes. (β2 = 90◦.)
Centrifugal compressors can be divided into three categories depending on the angle of
the vanes at the outlet: forward curved, β2 > 90◦, radial, β2 = 90◦ and backward curved, β2
< 90◦. As can be seen from figure 20 the tangential component of V2 varies with β2 in such
a way that with forward curved blades it is greater, and with backwards curved it is lesser.
That means that for forward curved vanes the energy transfer is larger, but at the price of
larger fluid losses. And respectively backwards curved vanes trade smaller losses, and
higher efficiency, with a lower energy transfer. Radial vanes are a compromise between
the too and have the added benefit of being easier to manufacture, with simpler geometry
and no bending stresses to take in to account.
In the combustion chamber is where the combustion takes place. Here, heat is added to
the jet engine in the Brayton Cycle. Compressed air flows from the compressor into the
chamber and ignites after being mixed with the fuel. The efficiency of the combustion is
given by
(23)
23
ṁfuel mass flow of fuel
h3 enthalpy of gas after combustor
h2 enthalpy of gas before combustor
LHV fuel heating value
The actual change of enthalpy in the chamber is given by ∆hactual and is divided by the
theoretical change of enthalpy, ∆htheoretical, given by the energy added by the fuel. For this
formula we assume an adiabatic process where no heat flows through the boundary of the
chamber ∆Q = 0.
The efficiency given is to see how much of the fuel that takes part in the combustion. Fuel
that is unburned is wasted and therefore reduces the efficiency. A major problem in
maintaining a high efficiency is loss of pressure. In an ideal Brayton cycle the pressure is
kept at a constant level through the combustion chamber. But with pressure losses from
such things as wall friction, turbulence and heat loss, it is not possible. There are three
stages in the combustion chamber. The recirculation zone, the burning zone and the
dilution zone. Here the fuel gets, respectively, evaporated and partially burned, and then
completely burned, and last mixed with bypass air to provide proper cooling. About 25%
to 35% of the incoming air is entered directly into the flame tube, where the combustion
takes place. The rest of the air is bypassed and used for cooling of the housing and to keep
a steady flame. An important part of the chamber is the diffuser, located before the liner,
which is used to slow the compressed air down to a speed better suited for combustion.
In addition to the diffuser the bypassed air is also used to create turbulence in the liner
which also slows the flow.
Figure 8 shows how the velocity of the gas relates to the fuel-air ratio, this shows the
important of velocity for a good combustion and also that there is an upper limit. A high
velocity and a high fuel-air ratio will give a rich blowout. Which means that oxygen is
displaced by fuel, which lowers the temperature of the flame and in some cases
distinguishes it. A lean blowout is where not enough fuel is given to the flame and could
also cause it to be distinguished, this is also used for lowering the engine RPM. The peak
velocity and optimal fuel-air ratio gives the best combustion.
Fuel-air ratio when burning propane, which is the fuel chosen for this project, is
approximately 1 kg propane per 12 m3 of air.
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Figure 28. Fuel-air ratio vs velocity.
Different from an internal combustion engine, in which ignition is needed at every cycle,
the jet engine works with a continuous flow. The igniter needs only to create a spark at
the start-up. Once the air and fuel-mixture has been ignited the combustion will be self-
sustained. There are three different types of combustion chambers used in aircrafts.
Annular, can and can-annular, as shown in cross-section in figure 28. All three types have
the same function, to increase the temperature of the high-pressure gas. The combustion
chamber used for this project is the annular combustion chamber.
Figure 29. Cross-section of can, annular and can-annular type combustion chamber.
Can and can-annular work in similar ways. The combustion takes place in several cans,
placed symmetrically around the shaft. The difference lies in the forming casing. The can-
annular has a more evenly structure, keeping the cans together. While the can type is kept
together by a ring-type structure.
The annular combustion chamber is the most common type. It is the most efficient of the
three and has the simplest structure.
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2.13. Gas Turbine Systems
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2.13.2. Power & district cooling application (Cogeneration)
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2.13.3. Combined heat & power application (Cogeneration)
As in figure 32 above the combined heat and power generation system whereby the
exhaust flow is now channeled to a HRSG which can be known as boiler as well. The
heat from the exhaust will heat up and tend to boil the coil of water in the tubing and
will change it to become steam and the steam will be used as the input for other
combined system (steam turbine) or for industries usage such as paper products. This
system is one of the ways to increase the efficiency of the gas turbine which it can
gives 80-85% of efficiency because the waste energy (exhaust) has been utilized
through the attached system.
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2.14. GasTurb
GasTurb is a gas turbine cycle program which is very powerful and versatile. It has a user
friendly graphical interface for simulating the different types of aircraft and power
generation gas turbines. Nomenclature used is clear and no cryptic abbreviations are used.
With GasTurb, it is easy to evaluate the thermodynamic cycles both for engine design and
off-design of the most common gas turbine architectures. In GasTurb even off design
transient mode simulations are possible.
In the gas turbine design process, many alternative thermodynamic cycles are evaluated.
Ultimately, a cycle will be selected which constitutes the cycle design point (cycle
reference point) of the gas turbine.
The mass flows, total pressures and total temperatures at the inlet and exit of all
components of the engine will then be determined on the basis of this design point.
Selection of appropriate Mach numbers and hub-tip ratios at the component boundaries
sets the aero-thermodynamically important dimensions of the gas turbine. Therefore by
establishing a cycle design point, the geometry of the flow annulus may be defined and
the weight of the gas turbine may be estimated.
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Figure 34. Design Point Calculation
Parametric studies give an overview of the design space. Two parameters can be varied
simultaneously generating a range of possible outcomes, visualized in a carpet plot. The
calculation can be completed quickly and the results plotted in various formats.
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2.14.1.3. Optimization
Numerical optimization can be employed for finding the best cycle. The optimization
algorithm searches for the best solution taking into account up to 12 optimization
variables and 12 constraints.
Any cycle output parameter, including the composed values, can be selected as a
figure of merit that can be maximized on the one hand (e.g. specific thrust) or
minimized on the other (e.g. specific fuel consumption).
The secondary air system used in GasTurb is able to simulate internal flows used
for turbine cooling and sealing as well as external flows for aircraft systems.
An engine diagram explains the secondary air system and the location of the
thermodynamic stations, for which all the temperatures, pressures, velocities, etc.
are shown in a table.
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Figure 37. Secondary Air System
GasTurb identifies the major thermodynamic stations between components using the
nomenclature described in SAE AS755. The values at each station are summarized in a
table.
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2.15. SolidWorks
SolidWorks is structured in three basic types: part mode, assembly mode and drawing
mode. Part mode is the basic building block in this software. For example, you must have
to create a part before you create assembly. Assembly mode contains parts or other
assemblies, called sub assemblies.
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3. METHOD
* The Overall Temperature Distribution Factor (OTDF) value must be found using the
formula below.
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Figure 40. Combustion Temperature Distribution Coefficient
The non-design point requirements of the engine to be designed are given in the table
below. Lines specified as non-design point inputs represent analysis inputs, and lines
specified as non-design point requirements represent target values to be reached as a result
of the analysis. The altitude of 5000 meters outside the design point is determined as 0.8
Mach speed.
(Standard day; static temperature at 5000 meters condition will be used from the ISA
table.)
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Parameter Value Instruction
Flight Speed (Mach) 0,8 The speed of the motor to be
designed outside the design
point is 0.8 Mach.
Flight Altitude (meter) 5000 The flight altitude of the
Design engine to be designed is 5000
Inputs meters.
JP8 Fuel Calorific 43124 The heating value of the JP8
Value (kj/kg) fuel to be used in the engine to
be designed will be 43124
kj/kg.
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3.2. Obtained Data from GasTurb
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Figure 43. Design Point Stations
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Figure 45. H-S Diagram
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Figure 47. External Factor Map
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Figure 49. Efficiency Graph
Figure 50. Variation of net thrust with the altitude at a Mach number of 0.8
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4. MY JET ENGINE DESIGN WITH SOLIDWORKS
Figure 52. My jet engine design with SolidWorks (Transparent section view)
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5. CONCLUSIONS AND RECOMMENDATIONS
The Jet engine's invention changed the future. With jet engines, planes can carry more
cargo, fly faster, and go farther than any propeller plane. The jet engine is a fascinating
piece of machinery, and though it’s basic principles are very simple there are many
variables to control for and many hurdles to overcome to build one of your own design.
In this study, a turbojet gas turbine engine was analyzed theoretically and computationally
in GasTurb program and a simple turbojet engine was designed in SolidWorks program.
Focusing on the studies in this type of computer modeling will be the basis for the
development of gas turbines and their digital control systems in the future. The science
of jet engines is still to a large degree an empirical one, where experimental data and a
good amount of trial and error is needed to push the field forward. The literature on this
subject is often dense and not always easy to navigate. Still, i have learned a lot about
both the theory, about the design processes.
43
6. REFERENCES
44
Name Özgür Atasever Scale Part Name Trakya University
Faculty of Engineering
Number 1171601906 Jet Engine Assembly Department of
1:200
Date May 2022 Mechanical Engineering
Part No 1
45
25,00
25,00
580,00
46
Name Özgür Atasever Scale Part Name Trakya University
Faculty of Engineering
Number 1171601906 Department of
1:100 Jet Engine
Date May 2022 Mechanical Engineering
Part No 3
47
Name Özgür Atasever Scale Part Name Trakya University
Faculty of Engineering
Number 1171601906 Department of
1:150 Shaft
Date May 2022 Mechanical Engineering
Part No 4
48
RESUME
Personal details
Name Özgür ATASEVER
Email adress ataseverozgur@gmail.com
Date of birth December 24th, 1997
Place of birth Şile / İstanbul
Education
2011 - 2015 Umraniye Technical High School, Istanbul
Computer aided manufacturing – Computer Numeric
Control (CNC)
Employment
2017 – 2017 Intern
Kormas Electric Motors
Skills
SolidWorks
AutoCad
Microsoft Office
Languages
Turkish
English
Hobbies
Computer
Design
Football
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