Model prototype of a jet engine

PROTOTYPE OF A JET ENGINE
DEPARTMENT OF MECHANICAL ENGINEERING 1
Visvesvaraya Technological University, Belagavi.
PROJECT REPORT
on
“PROTOTYPE OF A JET ENGINE”
Project Report submitted in partial fulfillment of the requirement for the award
of the degree of
Bachelor of Engineering
in
Mechanical Engineering
For the academic year 2012-2016
Submitted by
RADHIKA VINOD KUMAR (1CR12ME128)
ROHITHA G (1CR12ME093)
V DHARSHAN (1CR12ME116)
DEEPESH P. JAIN (1CR12ME023)
Under the guidance of
Mr. Joseph Sajan
Asst. Professor
Department of ME
CMRIT, Bangalore
2015-2016
Department of Mechanical Engineering
CMR INSTITUTE OF TECHNOLOGY, Bangalore-560037

DEPARTMENT OF MECHANICAL ENGINEERING
CERTIFICATE
This is to Certify that the dissertation work “PROTOTYPE OF A JET ENGINE” carried out by
Radhika Vinod Kumar, USN: 1CR12ME128, Rohitha G, USN: 1CR12ME093, V Dharshan, USN:
1CR12ME116, Deepesh P. Jain, USN: 1CR12ME023, bonafide students of CMRIT in partial
fulfillment for the award of Bachelor of Engineering in Mechanical Engineering of the
Visvesvaraya Technological University, Belagavi, during the academic year 2015-16. It is
certified that all corrections/suggestions indicated for internal assessment have been incorporated
in the report deposited in the departmental library. The project report has been approved as it
satisfies the academic requirements in respect of Project work prescribed for the said degree.
Signature of External Guide Signature of Internal Guide Signature of HOD
_________________ _________________ __________________
Mr. Vinod K. Mr. Joseph Sajan Prof. Rajendra Prasad Reddy
Director Assistant Professor Head of the Department
Enweld Technologies Pvt.Ltd. Mechanical Engineering Mechanical Engineering
Malur Industrial Area CMRIT, Bangalore. CMRIT, Bangalore.
Kolar District- 563130
External Viva
Name of Examiners
1.
2

DECLARATION
We Radhika Vinod Kumar (1CR12ME128), Rohitha G
(1CR12ME093), V Dharshan (1CR12ME116), Deepesh P Jain
(1CR12ME023), students of 8th
semester BE in Mechanical
Engineering, CMR Institute of Technology, Bangalore, hereby
declare that the project work entitled “Prototype of a Jet Engine”
submitted to The Visvesvaraya Technological University during the
academic year 2015-2016,is a record of an original work done by us
under the guidance of Mr. Vinod K. Director, Enweld Technologies
Pvt. Ltd. This project work is submitted in partial fulfillment of the
requirements for the award of the degree of Bachelor of Engineering
in Mechanical Engineering. The results embodied in this thesis have
not been submitted to any other University or Institute for the award
of any degree.
Date:
Place: Bangalore
Radhika Vinod Kumar
Rohitha G
V Dharshan
Deepesh P Jain

ACKNOWLEDGEMENT
The satisfaction and euphoria that accompany the successful completion of any task
would be incomplete without the mention of people who made it possible, whose
consistent guidance and encouragement crowned our efforts with success.
I consider it as my privilege to express the gratitude to all those who guided in the
completion of the project.
I express my gratitude to Principal, Dr. Sanjay Chitnis, for having provided me the
golden opportunity to undertake this project work in their esteemed organization.
I sincerely thank Prof. Rajendra Prasad Reddy, HOD, Department of Mechanical
Engineering, CMR Institute of Technology for the immense support given to me.
I express my gratitude to my project guide Mr. Joseph Sajan Assistant Professor and
Mr. Vinod K. for their support, guidance and suggestions throughout the project work.
Last but not the least, heartfull thanks to my parents and friends for their support.
Above all, I thank the Lord Almighty for His grace to succeed in this endeavor.

CONTENTS
 INTRODUCTION 7
 TYPES
 Turbo jet 8
 Turbo prop 9
 Turbo fan 10
 Turbo shaft 11
 Fuels used 14
 Materials used 15
 Beayton cycle 16
 Design
 Air intake 17
 Compressor 18
 Combustion Chamber 19
 Turbine 20
 Nozzle 23
 After burner 24
 Spark plug 25
 Induction motor 26
 Fuel inlet 27
 Thrust 28
 Working
 Actual Jet engine working 29
 Prototype Working 31
 Gas turbine analysis 34
 After burner thrust 38
 Summary 40
List of figures

 Turbo jet
 Turbo prop
 Turbo fan
 Turbo shaft
 Beayton cycle
 Air intake
 Combuction Chamber
 Turbine
 Turbocharger
 Blades
 After Burner
 Spark Plug
 Induction Motor
 Fuel Inlet
 Working
 After Burner Thrust
INTRODUCTION

An aircraft engine, or power plant, produces thrust to propel an aircraft. Reciprocating
engines and turboprop engines work in combination with a propeller to produce thrust.
Turbojet and turbofan engines produce thrust by increasing the velocity of air flowing
through the engine. All of these power plants also drive the various systems that support the
operation of an aircraft.
Jet engines operate according to Newton's third law of motion, which states that every force
acting on a body produces an equal and opposite force. The jet engine works by drawing in
some of the air through which the aircraft is moving, compressing it, combining it with fuel
and heating it, and finally ejecting the ensuing gas with such force that the plane is propelled
forward. The power produced by such engines is expressed in terms of pounds of thrust, a
term that refers to the number of pounds the engine can move.
Turbine Engines
An aircraft turbine engine consists of an air inlet, compressor, combustion chambers, a
turbine section, and exhaust. Thrust is produced by increasing the velocity of the air flowing
through the engine. Turbine engines are highly desirable aircraft power plants. They are
characterized by
• Smooth operation
• High power-to-weight ratio
• Readily available jet fuel.
Types of Turbine Engines

Turbine engines are classified according to the type of compressors they use. There are three
types of compressors—centrifugal flow, axial flow, and centrifugal-axial flow. Compression
of inlet air is achieved in a centrifugal flow engine by accelerating air outward perpendicular
to the longitudinal axis of the machine. The axial-flow engine compresses air by a series of
rotating and stationary airfoils moving the air parallel to the longitudinal axis. The
centrifugal-axial flow design uses both kinds of compressors to achieve the desired
compression.
The path the air takes through the engine and how power is produced determines the type of
engine. There are four types of aircraft turbine engines—turbojet, turboprop, turbofan, and
turboshaft
Turbojet
The turbojet engine consists of four sections: compressor, combustion chamber, turbine
section, and exhaust. The compressor section passes inlet air at a high rate of speed to the
combustion chamber. The combustion chamber contains the fuel inlet and igniter for
combustion. The expanding air drives a turbine, which is connected by a shaft to the
compressor, sustaining engine operation. The accelerated exhaust gases from the engine
provide thrust. This is a basic application of compressing air, igniting the fuel-air mixture,
producing power to self-sustain the engine operation, and exhaust for propulsion.

Turboprop
A turboprop engine is a turbine engine that drives a propeller through a reduction gear. The
exhaust gases drive a power turbine connected by a shaft that drives the reduction gear
assembly. Reduction gearing is necessary in turboprop engines because optimum propeller
performance is achieved at much slower speeds than the engine’s operating rpm. Turboprop
engines are a compromise between turbojet engines and reciprocating power plants.
Turboprop engines are most efficient at speeds between 250 and 400 mph and altitudes
between 18,000 and 30,000 feet. They also perform well at the slow airspeeds required for
takeoff and landing, and are fuel efficient.

Turbofan
Turbofans were developed to combine some of the best features of the turbojet and the
turboprop. Turbofan engines are designed to create additional thrust by diverting a secondary
airflow around the combustion chamber. The turbofan bypass air generates increased thrust,
cools the engine, and aids in exhaust noise suppression. This provides turbojet-type cruise
speed and lower fuel consumption.
The inlet air that passes through a turbofan engine is usually divided into two separate
streams of air. One stream passes through the engine core, while a second stream bypasses
the engine core. It is this bypass stream of air that is responsible for the term “bypass engine.”
A turbofan’s bypass ratio refers to the ratio of the mass airflow that passes through the fan
divided by the mass airflow that passes through the engine core.

Turboshaft
The fourth common type of jet engine is the turboshaft. It delivers power to a shaft that drives
something other than a propeller. The biggest difference between a turbojet and turboshaft
engine is that on a turboshaft engine, most of the energy produced by the expanding gases is
used to drive a turbine rather than produce thrust. Many helicopters use a turboshaft gas
turbine engine. In addition, turboshaft engines are widely used as auxiliary power units on
large aircraft.

Type Description Advantages Disadvantages
Turbojet
Generic term for
simple turbine
engine
Simplicity of design
Basic design, misses many
improvements in efficiency
and power
Turbofan
First stage
compressor
greatly enlarged
to provide bypass
airflow around
engine core
Quieter due to greater mass
flow and lower total exhaust
speed, more efficient for a
useful range of subsonic
airspeeds for same reason,
cooler exhaust temperature
Greater complexity (additional
ducting, usually multiple
shafts), large diameter engine,
need to contain heavy blades.
More subject to FOD and ice
damage. Top speed is limited
due to the potential for
shockwaves to damage engine

Ramjet
Intake air is
compressed
entirely by speed
of oncoming air
and duct shape
(divergent)
Very few moving parts,
Mach 0.8 to Mach 5+,
efficient at high speed (>
Mach 2.0 or so), lightest of
all airbreathing jets
(thrust/weight ratio up to 30
at optimum speed)
Must have a high initial speed
to function, inefficient at slow
speeds due to poor
compression ratio, difficult to
arrange shaft power for
accessories, usually limited to
a small range of speeds, intake
flow must be slowed to
subsonic speeds, noisy, fairly
difficult to test, finicky to kept
lit.
Scramjet
Similar to a
ramjet without a
diffuser; airflow
through the entire
engine remains
supersonic
Few mechanical parts, can
operate at very high Mach
numbers (Mach 8 to 15) with
good efficiencies.
Still in development stages,
must have a very high initial
speed to function (Mach >6),
cooling difficulties, very poor
thrust/weight ratio (~2),
extreme aerodynamic
complexity, airframe
difficulties, testing
difficulties/expense
Pulsejet
Air is
compressed and
combusted
intermittently
instead of
continuously.
Some designs use
valves.
Very simple design,
commonly used on model
aircraft
Noisy, inefficient (low
compression ratio), works
poorly on a large scale, valves
on valved designs wear out
quickly
Pulse
detonation
engine
Similar to a
pulsejet, but
combustion
occurs as a
Maximum theoretical engine
efficiency
Extremely noisy, parts subject
to extreme mechanical fatigue,
hard to start detonation, not
practical for current use

detonation
instead of a
deflagration, may
or may not need
valves
Rocket
Carries all
propellants
onboard, emits jet
for propulsion
Very few moving parts,
Mach 0 to Mach 25+,
efficient at very high speed
(> Mach 10.0 or so),
thrust/weight ratio over 100,
no complex air inlet, high
compression ratio, very high
speed (hypersonic) exhaust,
good cost/thrust ratio, fairly
easy to test, works in a
vacuum-indeed works best
exoatmospheric which is
kinder on vehicle structure at
high speed.
Needs lots of propellant- very
low specific impulse typically
100-450 seconds. Extreme
thermal stresses of combustion
chamber can make reuse
harder. Typically requires
carrying oxidiser onboard
which increases risks.
Extraordinarily noisy.
Fuels Used
Jet fuel, aviation turbine fuel (ATF), or avtur, is a type of aviation fuel designed for use in
aircraft powered by gas-turbine engines. It is colourless to straw-colored in appearance. The
most commonly used fuels for commercial aviation are Jet A and Jet A-1, which are
produced to a standardized international specification. The only other jet fuel commonly used
in civilian turbine-engine powered aviation is Jet B, which is used for its enhanced cold-
weather performance.
Jet fuel is a mixture of a large number of different hydrocarbons. The range of their sizes
(molecular weights or carbon numbers) is restricted by the requirements for the product, for
example, the freezing point or smoke point. Kerosene-type jet fuel (including Jet A and Jet

A-1) has a carbon number distribution between about 8 and 16 (carbon atoms per molecule);
wide-cut or naphtha-type jet fuel (including Jet B), between about 5 and 15.[1]
MATERIALS USED
Strong, lightweight, corrosion-resistant, thermally stable components are essential to the
viability of any aircraft design, and certain materials have been developed to provide these
and other desirable traits. Titanium, first created in sufficiently pure form for commercial use
during the 1950s, is utilized in the most critical engine components. While it is very difficult
to shape, its extreme hardness renders it strong when subjected to intense heat. To improve its
malleability titanium is often alloyed with other metals such as nickel and aluminum. All
three metals are prized by the aerospace industry because of their relatively high
strength/weight ratio.
The intake fan at the front of the engine must be extremely strong so that it doesn't fracture
when large birds and other debris are sucked into its blades; it is thus made of a titanium
alloy. The intermediate compressor is made from aluminum, while the high pressure section
nearer the intense heat of the combustor is made of nickel and titanium alloys better able to
withstand extreme temperatures. The combustion chamber is also made of nickel and
titanium alloys, and the turbine blades, which must endure the most intense heat of the
engine, consist of nickel-titanium-aluminum alloys. Often, both the combustion chamber and
the turbine receive special ceramic coatings that better enable them to resist heat. The inner
duct of the exhaust system is crafted from titanium, while the outer exhaust duct is made
from composites—synthetic fibers held together with resins.

BEAYTON CYCLE
The Brayton cycle is a thermodynamic cycle that describes the workings of a constant
pressure heat engine. The original Brayton engines used piston-compressor and expander
systems, but more modern gas turbine engines and airbreathing jet engines also follow the
Brayton cycle. Although the cycle is usually run as an open system (and indeed must be run
as such if internal combustion is used), it is conventionally assumed for the purposes of
thermodynamic analysis that the exhaust gases are reused in the intake, enabling analysis as a
closed system.
The engine cycle is named after George Brayton (1830–1892), the American engineer who
developed it originally for use in piston engines , although it was originally proposed and
patented by Englishman John Barber in 1791. There are two types of Brayton cycles, open to
the atmosphere and using internal combustion chamber or closed and using a heat exchanger.
E

MODEL PROTOTYPE OF A JET ENGINE
DESIGN
Air intake
An intake, or tube, is needed in front of the compressor to help direct the incoming air
smoothly into the moving compressor blades. Older engines had stationary vanes in front of
the moving blades. These vanes also helped to direct the air onto the blades. The air flowing
into a turbojet engine is always subsonic, regardless of the speed of the aircraft itself.
The intake has to supply air to the engine with an acceptably small variation in pressure
(known as distortion) and having lost as little energy as possible on the way (known as
pressure recovery). The ram pressure rise in the intake is the inlets contribution to the
propulsion system overall pressure ratio and thermal efficiency.
The intake gains prominence at high speeds when it transmits more thrust to the airfame than
the engine does.

COMPRESSOR
The compressor is driven by the turbine. It rotates at high speed, adding energy to the airflow
and at the same time compressing it into a smaller space. Compressing the air increases its
pressure and temperature. The smaller the compressor the faster it turns. At the large end of
the range the GE-90-115 fan rotates at about 2,500 RPM while a small helicopter engine
compressor rotates at about 50,000 RPM.
Turbojets supply bleed air from the compressor to the aircraft for the Environmental Control
System, anti-icing and fuel tank pressurization, for example. The engine itself needs air at
various pressures and flow rates to keep it running. This air comes from the compressor and
without it the turbines would overheat, the lubricating oil would leak from the bearing
cavities, the rotor thrust bearings would skid or be overloaded and ice would form on the
nose cone. The air from the compressor is called secondary air and is used for turbine
cooling, bearing cavity sealing, anti-icing and ensuring that the rotor axial load on its thrust
bearing will not wear it out prematurely. Supplying bleed air to the aircraft decreases the
efficiency of the engine because it has been compressed but then does not contribute to
producing thrust. Bleed air for aircraft services is no longer needed on the turbofan-powered
Boeing 787.
Compressor types used in turbojets were typically axial or centrifugal.
Early turbojet compressors had low pressure ratios up to about 5:1. Aerodynamic
improvements including splitting the compressor into two separately rotating parts,
incorporating variable blade angles for entry guide vanes and stators and bleeding air from
the compressor enabled later turbojets to have overall pressure ratios of 15:1 or more. For
comparison, modern civil turbofan engines have overall pressure ratios of 44:1 or more.
After leaving the compressor, the air enters the combustion chamber.

COMBUSTION CHAMBER
In a turbojet the air and fuel mixture burn in the combustor and pass through to the turbine in
a continuous flowing process with no pressure build-up. Instead there is a small pressure loss
in the combustor.
The fuel-air mixture can only burn in slow moving air so an area of reverse flow is
maintained by the fuel nozzles for the approximately stoichiometric burning in the primary
zone. Further compressor air is introduced which completes the combustion process and
reduces the temperature of the combustion products to a level which the turbine can accept.
Less than 25% of the air is typically used for combustion, as an overall lean mixture is
required to keep within the turbine temperature limits.

TURBINE
Hot gases leaving the combustor expand through the turbine. The hottest turbine vanes and
blades in an engine have internal cooling passages. Air from the compressor is passed
through these to keep the metal temperature within limits. The remaining stages don't need
cooling.
In the first stage the turbine is largely an impulse turbine and rotates because of the impact of
the hot gas stream. Later stages are convergent ducts that accelerate the gas. Energy is
transferred into the shaft through momentum exchange in the opposite way to energy transfer
in the compressor. The power developed by the turbine drives the compressor as well as
accessories, like fuel, oil, and hydraulic pumps that are driven by the accessory gearbox.

In a turbocharger, the wheels are designed to work 'radially' that is, the gases exit the
compressor and enter the turbine in a radial direction, i.e. at right angles to their axis of
rotation, which is the reason for the 'snail shell'-like shape to the housings. The reason for
this is efficiency, radial compressors and turbines work more efficiently below a certain size,
above these size axial compressors and turbines are used, but this is not an issue for us apart
from one of design compactness.
The third element that we need to build our jet engine requires us to build some form of
suitable combustion chamber. A turbocharger, when bolted to an engine is almost behaving
like a jet engine already, it provides compressed air to the engine's combustion chamber
where fuel is burned, the resulting gases then being forced out of the chamber by the piston
which spins the turbine wheel and hence driving the compressor. When we introduce a jet
engine style combustion chamber we effectively replace the engine and it's cylinders for the
burning of our fuel, turning the discrete 'suck, squeeze, bang, blow' cycle into a continuous
one as in a real turbojet. The combustion chamber will essentially be a large can into which
the fuel is sprayed and burned. The air from the turbocharger's compressor is fed in, fuel is
added, burned and the resulting hot expanding gases exit the combustion chamber through a
pipe connected to the inlet of the turbochargers turbine thereby completing the loop.

Combustion chambers have been constructed using a variety of basic materials, built up from
tubular steel or from modified fire extinguishers using mild or sometimes stainless steel for
durability.
Because of the inherent design of the turbocharger ( radial inflow wheels as opposed to the
more normal axial flow wheels ) and the fact that on most DIY jet engines we are using it 'as
is', the combustion chamber needs to be constructed 'outside' of the turbocharger as a separate
unit. This leads to the construction of a jet engine that is bulkier, heavier and far less
efficient, thrust for thrust, than their more streamlined commercial brethren (both full-size
and model jets) but is the price we have to pay in order to reduce complexity and cost to
achieve a real working jet.

NOZZLE
After the turbine, the gases are allowed to expand through the exhaust nozzle to atmospheric
pressure, producing a high velocity jet in the exhaust plume. In a convergent nozzle, the
ducting narrows progressively to a throat. The nozzle pressure ratio on a turbojet is usually
high enough for the expanding gases to reach Mach 1.0 and choke the throat. Normally, the
flow will go supersonic in the exhaust plume outside the engine.
If, however, a convergent-divergent de Laval nozzle is fitted, the divergent (increasing flow
area) section allows the gases to reach supersonic velocity within the nozzle itself. This is
slightly more efficient on thrust than using a convergent nozzle. There is, however, the added
weight and complexity since the convergent-divergent nozzle must be fully variable in its
shape to cope with changes in gas flow caused by engine throttling
AFTER BURNER
An afterburner or "reheat jetpipe" is a device added to the rear of the jet engine. It provides a
means of spraying fuel directly into the hot exhaust, where it ignites and boosts available
thrust significantly; a drawback is its very high fuel consumption rate. Afterburners are used
almost exclusively on supersonic aircraft – most of these are military aircraft. The two
supersonic civilian transports, Concorde and the TU-144, also utilized afterburners but these
two have now been retired from service. Scaled Composites White Knight, a carrier aircraft
for the experimental SpaceShipOne suborbital spacecraft, also utilizes an afterburner.
Intake exhaust
Air Intake = exhaust (thrust)
greater the air intake, greater the thrust

A

SPARK PLUG
A spark plug (sometimes, in British English, a sparking plug,[1] and, colloquially, a plug) is
a device for delivering electric current from an ignition system to the combustion chamber of
a spark-ignition engine to ignite the compressed fuel/air mixture by an electric spark, while
containing combustion pressure within the engine. A spark plug has a metal threaded shell,
electrically isolated from a central electrode by a porcelain insulator. The central electrode,
which may contain a resistor, is connected by a heavily insulated wire to the output terminal
of an ignition coil or magneto. The spark plug's metal shell is screwed into the engine's
cylinder head and thus electrically grounded. The central electrode protrudes through the
porcelain insulator into the combustion chamber, forming one or more spark gaps between
the inner end of the central electrode and usually one or more protuberances or structures
attached to the inner end of the threaded shell and designated the side, earth, or ground
electrode(s).

INDUCTION MOTOR
An electric motor is an electrical machine that converts electrical energy into mechanical
energy. The reverse of this would be the conversion of mechanical energy into electrical
energy and is done by an electric generator.
In normal motoring mode, most electric motors operate through the interaction between an
electric motor's magnetic field and winding currents to generate force within the motor. In
certain applications, such as in the transportation industry with traction motors, electric
motors can operate in both motoring and generating or braking modes to also produce
electrical energy from mechanical energy.

FUEL INLET
The inlet where the LPG connection is given to run the engine

THRUST
The net thrust F N of a turbojet is given by:
F N = ( ma i r + mf ) Vj −ma i rV
where
m a i r is the air
mf
is the rate of flow of fuel entering the engine
V j is the speed of the jet (the exhaust plume) and is assumed to be les
velocity
V is the true airspeed of the aircraft
(ma i r +mf ) Vj
represents the nozzle gross thrust
ma i r V represents the ram drag of the intake

WORKING
Actual Jet Engine working
• For a jet going slower than the speed of sound, the engine is moving through the air at
about 1000 km/h (600 mph). We can think of the engine as being stationary and the
cold air moving toward it at this speed.
• A fan at the front sucks the cold air into the engine and forces it through the inlet. This
slows the air down by about 60 percent and its speed is now about 400 km/h (240
mph).
• A second fan called a compressor squeezes the air (increases its pressure) by about eight
times, and this dramatically increases its temperature.
• Kerosene (liquid fuel) is squirted into the engine from a fuel tank in the plane's wing.
• In the combustion chamber, just behind the compressor, the kerosene mixes with the
compressed air and burns fiercely, giving off hot exhaust gases and producing a huge
increase in temperature. The burning mixture reaches a temperature of around 900°C
(1650°F).
• The exhaust gases rush past a set of turbine blades, spinning them like a windmill. Since
the turbine gains energy, the gases must lose the same amount of energy—and they do

so by cooling down slightly and losing pressure.
The turbine blades are connected to a long axle (represented by the middle gray line) that
runs the length of the engine. The compressor and the fan are also connected to this axle. So,
as the turbine blades spin, they also turn the compressor and the
 The hot exhaust gases exit the engine through a tapering exhaust nozzle. Just as water
squeezed through a narrow pipe accelerates dramatically into a fast jet (think of what
happens in a water pistol), the tapering design of the exhaust nozzle helps to
accelerate the gases to a speed of over 2100 km/h (1300 mph). So the hot air leaving
the engine at the back is traveling over twice the speed of the cold air entering it at the
front—and that's what powers the plane. Military jets often have an after burner that
squirts fuel into the exhaust jet to produce extra thrust. The backward-moving exhaust
gases power the jet forward. Because the plane is much bigger and heavier than the
exhaust gases it produces, the exhaust gases have to zoom backward much faster than
the plane's own speed.
CONNECTIONS

 The induction motor is connected to the AC current and negative of the coil is
connected to the induction motor.
 Negative of the coil is connected to the battery.Positive of the battery is connected to
the one of the spark plugs.
 Air is supplied to the inlet through a blower.
 LPG is supplied at two spots,one at the after burner and the second inlet at the
compressor.
 Sparl plugs are connected to the induction motor.
How the prototype works
 A gas turbine, also called a combustion turbine, is a type of internal combustion
engine. It has an upstream rotating compressor coupled to a downstream turbine, and
a combustion chamber in-between.
Energy is added to the gas stream in the combustor, where fuel is mixed with air and ignited.
In the high pressure environment of the combustor, combustion of the fuel increases the
temperature. The products of the combustion are forced into the turbine section. There, the
high velocity and volume of the gas flow is directed through a nozzle over the turbine's
blades, spinning the turbine which powers the compressor and, for some turbines, drives their
mechanical output. The energy given up to the turbine comes from the reduction in the
temperature and pressure of the exhaust gas
 Preceding the compressor is the air intake (or inlet). It is designed to be as efficient as
possible at recovering the ram pressure of the air stream tube approaching the intake.
The air leaving the intake then enters the compressor. The stators (stationary blades)
guide the airflow of the compressed gases.
 The compressor is driven by the turbine. The compressor rotates at very high speed,
adding energy to the airflow and at the same time squeezing (compressing) it into a
smaller space. Compressing the air increases its pressure and temperature.In most
turbojet-powered aircraft, bleed air is extracted from the compressor section at
various stages to perform a variety of jobs including air conditioning/pressurization,

engine inlet anti-icing and turbine cooling. Bleeding air off decreases the overall
efficiency of the engine, but the usefulness of the compressed air outweighs the loss
in efficiency.Several types of compressor are used in turbojets and gas turbines in
general: axial, centrifugal, axial-centrifugal, double-centrifugal, etc.Early turbojet
compressors had overall pressure ratios as low as 5:1 (as do a lot of simple auxiliary
power units and small propulsion turbojets today). Aerodynamic improvements, plus
splitting the compression system into two separate units and/or incorporating variable
compressor geometry, enabled later turbojets to have overall pressure ratios of 15:1 or
more. For comparison, modern civil turbofan engines have overall pressure ratios of
44:1 or more.After leaving the compressor section, the compressed air enters the
combustion chamber.
 The burning process in the combustor is significantly different from that in a piston
engine. In a piston engine the burning gases are confined to a small volume and, as
the fuel burns, the pressure increases dramatically. In a turbojet the air and fuel
mixture passes unconfined through the combustion chamber. As the mixture burns its
temperature increases dramatically, but the pressure actually decreases a few
percent.The fuel-air mixture must be brought almost to a stop so that a stable flame
can be maintained. This occurs just after the start of the combustion chamber. The aft
part of this flame front is allowed to progress rearward. This ensures that all of the
fuel is burned, as the flame becomes hotter when it leans out, and because of the
shape of the combustion chamber the flow is accelerated rearwards. Some pressure
drop is required, as it is the reason why the expanding gases travel out the rear of the
engine rather than out the front. Less than 25% of the air is involved in combustion, in
some engines as little as 12%, the rest acting as a reservoir to absorb the heating
effects of the burning fuel.Another difference between piston engines and jet engines
is that the peak flame temperature in a piston engine is experienced only momentarily
in a small portion of the full cycle. The combustor in a jet engine is exposed to the
peak flame temperature continuously and operates at a pressure high enough that a
stoichiometric fuel-air ratio would melt the can and everything downstream. Instead,
jet engines run a very lean mixture, so lean that it would not normally support
combustion. A central core of the flow (primary airflow) is mixed with enough fuel to
burn readily. The cans are carefully shaped to maintain a layer of fresh unburned air
between the metal surfaces and the central core. This unburned air (secondary

airflow) mixes into the burned gases to bring the temperature down to something a
turbine can tolerate.
 Hot gases leaving the combustor are allowed to expand through the turbine. Turbines
are usually made up of high temperature metals such as inconel to resist the high
temperature, and frequently have built-in cooling channels.In the first stage the
turbine is largely an impulse turbine (similar to a pelton wheel) and rotates because of
the impact of the hot gas stream. Later stages are convergent ducts that accelerate the
gas rearward and gain energy from that process. Pressure drops, and energy is
transferred into the shaft. The turbine's rotational energy is used primarily to drive the
compressor. Some shaft power is extracted to drive accessories, like fuel, oil, and
hydraulic pumps. Because of its significantly higher entry temperature, the turbine
pressure ratio is much lower than that of the compressor. In a turbojet almost two-
thirds of all the power generated by burning fuel is used by the compressor to
compress the air for the engine.
 After the turbine, the gases are allowed to expand through the exhaust nozzle to
atmospheric pressure, producing a high velocity jet in the exhaust plume. In a
convergent nozzle, the ducting narrows progressively to a throat. The nozzle pressure
ratio on a turbojet is usually high enough for the expanding gases to reach Mach 1.0
and choke the throat. Normally, the flow will go supersonic in the exhaust plume
outside the engine.If, however, a convergent-divergent de Laval nozzle is fitted, the
divergent (increasing flow area) section allows the gases to reach supersonic velocity
within the nozzle itself. This is slightly more efficient on thrust than using a
convergent nozzle. There is, however, the added weight and complexity since the
convergent-divergent nozzle must be fully variable in its shape to cope with changes
in gas flow caused by engine throttling.
 An afterburner or "reheat jetpipe" is a device added to the rear of the jet engine. It
provides a means of spraying fuel directly into the hot exhaust, where it ignites and
boosts available thrust significantly; a drawback is its very high fuel consumption
rate. Afterburners are used almost exclusively on supersonic aircraft – most of these
are military aircraft. The two supersonic civilian transports, Concorde and the TU-
144, also utilized afterburners but these two have now been retired from service.
Scaled Composites White Knight, a carrier aircraft for the experimental
SpaceShipOne suborbital spacecraft, also utilizes an afterburner.

 Realistic Simple-Cycle Gas Turbine Analysis
The preceding analysis of the Air Standard cycle assumes perfect turbomachinery, an
unachievable but meaningful ideal, and room-temperature heat capacities. Realistic
quantitative performance information can be obtained by taking into account efficiencies of
the compressor and the turbine, significant pressure losses, and more realistic thermal
properties.
Properties for Gas Turbine Analysis
It is pointed out in reference 1 that accurate gas turbine analyses may be performed using
constant heat capacities for both air and combustion gases. This appears to be a
specialization of a method devised by Whittle (ref. 4). The following properties are therefore
adopted for all gas turbine analyses in this book:
Air:
cp = 0.24 Btu/lbm-R or 1.004 kJ /kg-K k =
1.4 implies k /(k 1) = 3.5
Combustion gas:
cp,g = 0.2744 Btu/lbm-R or 1.148 kJ /kg-K kg =
1.333 implies kg/(kg 1) = 4.0
The properties labeled as combustion gas above are actually high-temperature-air properties.
Because of the high air-fuel ratio required by gas turbines, the thermodynamic properties of
gas turbine combustion gases usually differ little from those of high-temperature air. Thus
the results given below apply equally well to closed-cycle machines using air as the working
fluid and to open-cycle engines.
Analysis of the Open Simple-Cycle Gas Turbine
A simple-cycle gas turbine has one turbine driving one compressor and a power-consuming
load. More complex configurations are discussed later. It is assumed that the compressor
inlet state, the compressor pressure ratio, and the turbine inlet temperature are known, as
before. The turbine inlet temperature is usually determined by the limitations of the high-
temperature turbine blade material. Special metals or ceramics are usually selected for their

ability to withstand both high stress at elevated temperature and erosion and corrosion
caused by undesirable components of the fuel.
As shown in Figure 5.3, air enters the compressor at a state defined by T1 and p1.
The compressor exit pressure, p2, is given by
p2 = rp1 [lbf /ft2 | kPa] (5.10)
where r is the compressor pressure ratio. The ideal compressor discharge temperature, T2s is
given by the isentropic relation
T2s = T1 r (k 1)/k [R | K] (5.11)
The compressor isentropic efficiency, defined as the ratio of the compressor isentropic work
to the actual compressor work with both starting at the same initial state and ending at the
same pressure level, may be written as
c = ( h1 h2s )/( h1 h2 ) = ( T1 T2s )/( T1 T2 ) [dl] (5.12)
Here the steady-flow energy equation has been applied to obtain expressions for the work
for an irreversible adiabatic compressor in the denominator and for an isentropic compressor
in the numerator. Solving Equation (5.12) for T2, we get as the actual compressor discharge
temperature:
T2 = T1 + ( T2s T1 ) / c [R | K] (5.13)
Equation (5.3) then gives the work needed by the compressor, wc:
wc = cp ( T1 T2 ) = cp ( T1 T2s )/ c [Btu /lbm | kJ/kg] (5.14)
Note that the compressor work is negative, as required by the sign convention that defines
work as positive if it is produced by the control volume. The compressor power requirement
is, of course, then given by mawc [Btu/hr | kW], where ma is the
compressor mass flow rate [lbm / hr | kg / s].
After leaving the compressor at an elevated pressure and temperature, the air then enters
the combustion chamber, where it completely oxidizes a liquid or a gaseous fuel injected
under pressure. The combustion process raises the combustion gas temperature to the turbine
inlet temperature T3. One of the goals of combustion chamber design is to minimize the
pressure loss from the compressor to the turbine. Ideally, then, p3 = p2, as assumed by the

Air Standard analysis. More realistically, a fixed value of the combustor fractional pressure
loss, fpl, (perhaps about 0.05 or 5%) may be used to account for burner losses:
fpl = (p2 p3)/p2 [dl] (5.15)
Then the turbine inlet pressure may be determined from
p3 = (1 fpl) p2 [lbf /ft2 | kPa] (5.16)
Rather than deal with its complexities, we may view the combustion process simply as one
in which heat released by exothermic chemical reaction raises the temperature of
combustion gas (with hot-air properties) to the turbine inlet temperature. The rate of heat
released by the combustion process may then be expressed as:
Qa = ma(1 + f )cp,g(T3 T2) [Btu/hr | kW] (5.17)
where f is the mass fuel-air ratio. The term ma(1 + f) is seen to be the sum of the air and fuel
mass flow rates, which also equals the mass flow rate of combustion gas. For gas turbines it
will be seen later that f is usually much less than the stoichiometric fuel-air ratio and is often
neglected with respect to 1 in preliminary analyses.
The turbine in the open-cycle engine operates between the pressure p3 and atmospheric
pressure, p4 = p1, with an inlet temperature of T3. If the turbine were isentropic, the
discharge temperature would be
T4s = T3( p4 /p3 ) (kg 1) / kg [R | K] (5.18)
From the steady-flow energy equation, the turbine work can be written as wt = cp,g (T3
T4 ) = t cp,g( T3 T4s) [Btu/lbm | kJ/kg] (5.19)
referenced to unit mass of combustion gas, and where t is the turbine isentropic efficiency.
The turbine power output is then ma(1 + f)wt, where, as seen earlier, ma(1 + f) is the mass
flow rate of combustion gas flowing through the turbine. The net work based on the mass of
air processed and the net power output of the gas turbine, Pn, are then given by
wn = (1 + f )wt + wc [Btu/lbm air| kJ/kg air] and (5.20)
Pn = ma [(1 + f )wt + wc ] [Btu/hr | kW] and the
thermal efficiency of the engine is
(5.21)
th = Pn /Qa [dl] (5.22)

5.1
A simple-cycle gas turbine has 86% and 89% compressor and turbine efficiencies,
respectively, a compressor pressure ratio of 6, a 4% fractional pressure drop in the
combustor, and a turbine inlet temperature of 1400 F. Ambient conditions are 60 F and one
atmosphere. Determine the net work, thermal efficiency, and work ratio for the engine.
Assume that the fuel-mass flow rate is negligible compared with the air flow rate.
THRUST

To move an airplane through the air, thrust is generated by some kind of propulsion system.
Most modern fighter aircraft employ an afterburner on either a low bypass turbofan or a
turbojet. On this page we will discuss some of the fundamentals of an afterburning turbojet .
In order for fighter planes to fly supersonically , they have to overcome a sharp rise in drag
near the speed of sound. A simple way to get the necessary thrust is to add an afterburner to a
core turbojet. In a basic turbojet, some of the energy of the exhaust from the burner is used to
turn the turbine. The afterburner is used to put back some energy by injecting fuel directly
into the hot exhaust. On the schematic, you'll notice that the nozzle of the basic turbojet has
been extended and there is now a ring of flame holders, colored yellow, in the nozzle. When
the afterburner is turned on, additional fuel is injected through the hoops and into the hot
exhaust stream of the turbojet. The fuel burns and produces additional thrust, but it doesn't
burn as efficiently as it does in the combustion section of the turbojet. You get more thrust,
but you burn much more fuel. With the increased temperature of the exhaust, the flow area of
the nozzle has to be increased to pass the same mass flow. Therefore, afterburning nozzles
must be designed with variable geometry and are heavier and more complex than simple

turbojet nozzles. When the afterburner is turned off, the engine performs like a basic turbojet.
You can investigate nozzle operation with our interactive nozzle simulator.
The nozzle of a turbojet is usually designed to take the exhaust pressure back to free stream
pressure. The thrust equation for an afterburning turbojet is then given by the general thrust
equation with the pressure-area term set to zero. If the free stream conditions are denoted by a
"0" subscript and the exit conditions by an "e" subscript, the thrust F is equal to the mass
flow rate m dot times the velocity V at the exit minus the free stream mass flow rate times
the velocity.
F = [m dot * V]e - [m dot * V]0
This equation contains two terms. Aerodynamicists often refer to the first term (m dot * V)e
as the gross thrust since this term is largely associated with conditions in the nozzle. The
second term (m dot * V)0 is called the ram drag and is usually associated with conditions in
the inlet. For clarity, the engine thrust is then called the net thrust. Our thrust equation
indicates that net thrust equals gross thrust minus ram drag.
Afterburners are only used on fighter planes and the supersonic airliner,
Concorde. The Concorde turns the afterburners off once it gets into cruise.
Otherwise, it would run out of fuel before reaching Europe. Afterburners offer
a mechanically simple way to augment thrust and are used on both turbojets
and turbofans.
SUMMARY

Phase 1 - Beginnings...
Phase 1 document my first faltering steps, blind alleys, false
starts, crude designs, naive assumptions and a rush to get
things going spurred on by turbine fever...
Phase 2 - Better equipped...
Phase 2 followed a period of rest, a step back and a more
measured approach to things. I wanted to build something that
worked first time and looked right. I still had a way to go and
the design was constantly changing, resulting in new parts to
find and re-making old parts, so was a constant two steps
forward and a one stepback process!
Phase 3 - New Ideas...
Having partially documented Phase 2, I've since changed my
designs to incorporate a better oil/fuel systemwithout the need
for heavy batteries using a model aero engine, an idea for oil
cooling using the fuel itself, an on-board starting device and
my own design for an adjustable jet nozzle for easing the
process of fine tuning for maximum thrust!

Phase 4 – Free power!
Moving a vehicle using pure thrust is fine, but for faster
acceleration and making more efficient use of the exhaust
gases from a gas turbine requires the use of a free power
setup...
Phase 5 - Electronics
Any vehicle needs devices to monitor the operating conditions
of the various systems. A DIY jet powered vehicle is no
exception. Sensors are needed to measure and display
temperatures, pressures, speed, RPM, etc. These
measurements can also be used to control other systems, e.g.
switching on cooling fans, controlling start up procedures or
shutting down the engine if critical conditions occur. At the
very least temperature measurement of the hot exhaust gases
impinging on the turbine wheel is essential to avoid meltdown
and some form of RPM sensor for indicating the speedof the
turbo without which over-speedcan easily lead to a destructive
scenario! The ideal solution would be to have some form of
complete process control system. All of the desired
measurements would be fed in to a microcontroller type
processing device or PC/laptop where display, logging and
monitoring can be performed, start-up sequences can be
programmed and decisions taken to avoid disastrous
situations. This is a huge project in itself so I am starting with
the minimum requirements, an RPM sensor(s) and exhaust gas
temperature measurement.

Model prototype of a jet engine

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