MODULE – I

AIR-BREATHING ENGINES

Propulsion

Propulsion (Lat. pro-pellere, push forward) is making a body to move (against natural
forces), i.e. fighting against the natural tendency of relative-motion to decay. Motion is relative
to an environment. Sometimes, propulsion is identified with thrust, the force pushing a body to
move against natural forces, and one might say that propulsion is thrust (but thrust not
necessary implies motion, as when pushing against a wall; on the other hand, propulsion
implies thrust).

Sometimes a distinction is made from propulsion (pushing) to traction (pulling), but,

leaving aside internal stresses in the system (compression in the first case and tension in the
latter), push and pull motions produces the same effect: making a body to move against natural
forces. In other occasions 'traction' is restricted to propulsion by shear forces on solid surfaces.

The special case of creating aerodynamic thrust to just balance gravitational attraction
of a vehicle without solid contact (e.g. hovering helicopters, hovering VTOL-aircraft, hovering
rockets, hovering hovercraft...), is often included as propulsion (it is usually based on the same
engine), though its propulsion efficiency is zero.

Operational envelopes

Each engine type will operate only within a certain range of altitudes and Mach
numbers (velocities). Similar limitations in velocity and altitude exist for airframes. It is
necessary, therefore, to match airframe and propulsion system capabilities.

Figure1.1 shows the approximate velocity and altitude limits, or corridor of flight,
within which airlift vehicles can operate. The corridor is bounded by a lift limit, a temperature
limit, ·and an aerodynamic force limit. The lift limit is determined by the maximum level-flight
altitude at a given velocity. The temperature limit is set by the structural thermal limits of the
material used in construction of the · aircraft. At any given altitude, the maximum velocity
attained is temperature-limited by aerodynamic heating effects. At lower altitudes, velocity is
limited by aerodynamic force loads rather than by temperature.

The operating regions of all aircraft lie within the flight corridor. The operating region
of a particular aircraft within the corridor is determined by aircraft design, but it is a very small
portion of the overall corridor. Superimposed on "the flight corridor in Fig.1.1 are the
operational envelopes of various powered aircraft. The operational limits of each propulsion
system are determined by limitations of the components of the propulsion system and are
shown in Fig. 1.2.

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 Fig.1.1: Flight limits.

Fig.1.2:Engine operational limits.

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Classification of air-breathing engine

The air-breathing engine classified into following types:

1. turbojet
2. turbofan
3. turboprop
4. turboshaft
5. ramjet
6. scramjet

Gas Generator

The "heart" of a gas turbine type of engine is the gas generator. A schematic diagram of a gas
generator is shown in Fig. 1.3. The compressor, combustor, and turbine are the major
components of the gas generator which is common to the turbojet, turbofan, turboprop, and
turboshaft engines. The purpose of a gas generator is to supply high-temperature and high-
pressure gas.

Fig.1.3:Schematic diagram of gas generator.

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Turbojet engine

By adding an inlet and a nozzle to the gas generator, a turbojet engine can be constructed. A
schematic diagram of a afterburner is shown Fig. 1.4.The thrust of a turbojet is developed by
compressing air in the inlet and compressor, mixing the air with fuel and burning in the combustor, and
expanding the gas stream through the turbine and nozzle. The expansion of gas through the turbine
supplies the power to turn the compressor. The net thrust delivered by the engine is the result of
converting internal energy to kinetic energy.

The different processes in a turbojet cycle are the following:

a-1: Air from far upstream is brought to the air intake (diffuser) with some
acceleration/deceleration
1-2: Air is decelerated as is passes through the diffuser
2-3: Air is compressed in a compressor (axial or centrifugal)
3-4:The air is heated using a combustion chamber/burner
4-5: The air is expanded in a turbine to obtain power to drive the compressor
5-6: The air may or may not be further heated in an afterburner by adding further fuel
6-7: The air is accelerated and exhausted through the nozzle to produce thrust.

Fig.1.4: Turbo jet engine with after burner

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 (a) (b)

Fig.1.5: (a) Ideal turbojet cycle (without afterburning). (b) Ideal turbojet cycle with
afterburning

Afterburning: used when the aircraft needs a substantial increment in thrust.For eg.,To
accelerate to and cruise at supersonic speeds.Since the air-fuel ratio in gas turbine engines are
much greater than the stoichiometric values, there is sufficient amount of air available for
combustion at the turbine exit. There are no rotating components like a turbine in the
afterburner, the temperatures can be taken to much higher values than that at turbine entry.

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 Fig.1.6: Variation of flow properties across turbojet engine

The pressure, temperature, and velocity variations through a jet engine are shown in Fig. 1.6. In
the compressor section, the pressure and temperature increase as a result of work being done on
the air. The temperature of the gas is further increased by burning in the combustor. In the
turbine section, energy is being removed from the gas stream and converted to shaft power to
turn the compressor. The energy is removed by an expansion process which results in a
decrease of temperature and pressure. In the nozzle, the gas stream is further expanded to
produce a high exit kinetic energy. All the sections of the engine must operate in such a way as
to efficiently produce the greatest amount of thrust for a minimum of weight.

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Turbo fan engine

The turbofan engine consists of an inlet, fan, gas generator, and nozzle. A schematic diagram of
a turbofan is shown in Fig. 17. In the turbofan, a portion of the turbine work is used to supply
power to the fan. Generally the turbofan engine is more economical and efficient than the
turbojet engine in" a limited realm of flight. The thrust specific fuel consumption (TSFC, or
fuel mass flow rate per unit thrust) is lower for turbofans and indicates a more economical
operation. The turbofan also accelerates a larger mass of air to a lower velocity than a turbojet
for a higher propulsive efficiency. The frontal area of a turbofan is quite large compared to that
of a turbojet, and for this reason more drag and more weight result. The fan diameter is also
limited aerodynamically when compressibility -effects occur.

Variation of flow properties across jet engine

Fig.1.7: Schematic diagram of a high-bypass-ratio turbofan.

Bypass ratio

The bypass ratio (BPR) of a turbofan engine is the ratio between the mass flow rate of the
bypass stream to the mass flow rate entering the core. A 10:1 bypass ratio, for example, means
that 10 kg of air passes through the bypass duct for every 1 kg of air passing through the core.

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The Turboprop and Turboshaft engine

A gas generator that drives a propeller is a turboprop engine. The expansion of gas through the
turbine supplies the energy required to turn the propeller. A schematic diagram of the turboprop
is shown in Fig. 1.8. The turboshaftengine is similar to the turboprop except that power is
supplied to a shaft rather than a propeller. The turboshaft engine is used quite extensively for
supplying power for helicopters. The turboprop engine may find application in VTOL (vertical
takeoff and landing) transporters. The limitations and advantages of the turboprop are those of
the propeller. For low-speed flight and, short-field takeoff, the propeller has a performance
advantage. At speeds approaching the speed of sound, compressibility effects set in and the
propeller loses its aerodynamic efficiency. Due to the rotation of the propeller, the propeller tip
will approach the speed of sound before the vehicle approachesthe speed of sound. This
compressibility effect when one approaches the speed of sound limits the design of helicopter
rotors and propellers. At high subsonic speeds, the turbofan engine will have a better
aerodynamic performance than the turboprop since the turbofan is essentially a ducted
turboprop. Putting a duct or shroud around a propeller increases its aerodynamic performance.

Fig.1.8:Schematic diagram of a turboprop.

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The Ramjet

The ramjet engine consists of an inlet, a combustion zone, and a nozzle. Aschematic
diagram of a ramjet is shown in Fig. 1.9. The ramjet does not have the compressor and turbine
as the turbojet does. Air enters the inlet where it is compressed and then enters the combustion
zone where it is mixed with the fuel and burned. The hot gases are then expelled through the
nozzle, developing thrust. The operation of the ramjet depends upon the inlet to decelerate the
incoming air to raise the pressure in the combustion zone. The pressure rise makes it possible
for the ramjet to operate. The higher the velocity of the incoming air, the greater the pressure
rise. It is for this reason that the ramjet operates best at high supersonic velocities. At subsonic
velocities, the ramjet is inefficient, and to start the ramjet, air at a relatively higher velocity
must enter the inlet.

The combustion process in an ordinary ramjet takes place at low subsonic velocities. At
high supersonic flight velocities, a very large pressure rise is developed that is more than
sufficient to support operation of the ramjet. Also, if the inlet has to decelerate a supersonic
high-velocity airstream to a subsonic velocity, large pressure losses can result. The deceleration
process also produces a temperature rise, and at some limiting flight speed, the temperature will
approach the limit set by the wall materials and cooling methods. Thus when the temperature
increase due to deceleration reaches the limit, it may not be possible to burn fuel in the
airstream.

Fig.1.9:Schematic diagram of a ramjet.

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SCRAMJET

The ramjet with supersonic combustion is known as the scramjet (supersonic combustion
ramjet). By using a supersonic combustion process, the temperature rise and pressure loss due
to deceleration in the inlet can be reduced.

Fig.1.10:Schematic diagram of a Scramjet.

Fig.1.11:Schematic diagram of a Scramjet with station number.

The scramjet engine belongs to the family of Brayton cycles, which consist of two adiabatic
and two constant-pressure processes. A simplified schematic of a scramjet equipped vehicleis
shown above Fig.1.10.

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Station 0 represents the free-stream condition.

Station 1 represents the beginning of the compression process. Hypersonic shock-wave

angles are small, resulting in long compression ramps (or spikes if an
axisymmetric configuration is used) that, in many of the suggested
configurations, begin at the vehicle‘s leading edge. Additional compression
takes place inside the inlet duct.

Station 2.1 represents the entrance into the isolator section. The role of the isolator is to
separate the inlet from the adverse effects of a pressure rise that is due to
combustion in the combustion chamber. The presence of a shock train in the
isolator further compresses the air before arriving at the combustion chamber.
Thermodynamically the isolator is not a desirable component, because it is a
source of additional pressure losses, increases the engine cooling loads, and adds
to the engine weight. However, operationally it is needed to include a shock
train that adjusts such that it fulfills the role just described.

Station 3 is the combustion chamber entrance. Unlike the turbojet engine cycle, in which
the air compression ratio is controlled by the compressor settings, in a fixed-
geometry scramjet the pressure at the combustion chamber entrance varies over
a large range.

Station 4 is the combustion chamber exit and the beginning of expansion.

Station 10 is the exit from the nozzle; because of the large expansion ratios the entire aft
part of the vehicle may be part of the engine nozzle.

The thrust equation

The following assumptions are made:

1. The flow is steady within the control volume; thus all the properties within the control
do not change with time.

2. The external flow is reversible; thus the pressures and velocities are constant over the
control surface except over the exhaust area Pe of the engine.

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 Fig.1.12:Generalized thrust producing device.

Engine performance parameters

The engine performance is described by different efficiency definitions, thrust and the fuel
consumption. The efficiency definitions that we shall now be discussing are applicable to an
engine with a single propellant stream (turbojets or ramjets). For other types of jet engines
(turbofan, turboprop) the equations need to be appropriately modified.

Thermal efficiency: The ratio of the rate of production of propellant kinetic energy to the total
energy consumption rate.

For a turboprop or turboshaft engine, the output is largely shaft power. In this case,

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Propulsion efficiency: The ratio of thrust power to the rate of production of propellant kinetic
energy. The propulsive efficiency ηp can be defined as the ratio of the useful propulsive energy
or thrust power (F. u) to the sum of that energy and the unused kinetic energy of the jet.

If we assume that f«1 and the pressure thrust term is negligible

Overall efficiency: The product of thermal efficiency and propulsion efficiency.

Specific Thrust

Specific thrust is a term used in gas turbine engineering to show the relative thrust per air mass
flow rate of a jet engine (e.g. turbojet, turbofan, etc.) and is defined as the ratio: net thrust/total
intake airflow.

Why are we interested in specific thrust?

First, it is an indication of engine efficiency. Two different engines have different values of
specific thrust. The engine with the higher value of specific thrust is more efficient because it
produces more thrust for the same amount of airflow.

It gives us an easy way to "size" an engine during preliminary analysis. The result of our
thermodynamic analysis is a certain value of specific thrust. The aircraft drag defines the
required value of thrust. Dividing the thrust required by the specific thrust tells us how much
airflow our engine must produce and this determines the physical size of the engine.

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Specific impulse

Specific impulse (usually abbreviated Isp) is a measure of how effectively a rocket uses
propellant or jet engine uses fuel. By definition, it is the total impulse (or change in
momentum) delivered per unit of propellant consumed.

Thrust Specific fuel consumption

TSFC or SFC for thrust engines (e.g. turbojets, turbofans, ramjets, rocket engines, etc.) is the
mass of fuel needed to provide the net thrust for a given period.SFC varies with throttle setting,
altitude and climate. For jet engines, flight speed also has a significant effect upon SFC; SFC is
roughly proportional to air speed.

Component performance

Air intake performance

Inlet losses arise due to wall friction and shock waves (in a supersonic inlet).
•These result in a reduction in total pressure.
•The flow is usually adiabatic as it flows through the intake.
•Performance of intakes arecharacterized using total pressure ratio and isentropic efficiency.
Isentropic efficiency, ηd, of the diffuser is

This efficiency can be related to the total pressure ratio (πd) and Mach number

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Compressor/fan performance
Compressors are to a high degree of approximation, adiabatic. Compressor performance is
evaluated using the isentropic efficiency.

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The isentropic efficiency is thus a function of the total pressure ratio and the total temperature
ratio.

Combustion chamber performance

In a combustion chamber (or burner), there are two possibilities of losses, incomplete
combustion and total pressure losses.
 Combustion efficiency can be defined by carrying out an energy balance across the
combustor.
 Two different values of specific heat at constant pressure: one for fluid upstream of the
combustor and the other for fluid downstream of the combustor

Total pressure losses arise from two effects:

 viscous losses in the combustion chamber
 total pressure loss due to combustion at finite Mach number.

Turbine performance
The flow in a turbine is also assumed to be adiabatic, though in actual engines there could be
turbine blade cooling.Isentropic efficiency of the turbine is defined in a manner similar to that
of the compressor.

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Nozzle performance
The flow in the nozzle is also adiabatic,however losses in a nozzle could occur due to
incomplete expansion process (under or over-expansion). Friction may reduce the isentropic
efficiency.

The efficiency is defined by

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Afterburner performance
Afterburner is thermodynamically similar to a combustion chamber. The performance
parameters for an afterburner are thus the combustion efficiency and the total pressure loss. In
case of engines with afterburning, the corresponding performance parameters for an afterburner
need to be taken into account.
As stated above the propeller rotates at very low speed compared to its driving turbine. The
speedreduction may be 1:15. This speed reduction is necessary owing to two reasons:
1. A large centrifugal force arises from the rotation of the large diameter (2–4 m or
evenmore) propeller blades. These blades are fixed to the propeller hub in a cantilever
fixedendconfiguration. Consequently, such a centrifugal force generates a large tensile
stressat blade root. Stress limitations require that the large diameter propeller rotates at
a muchslow speed. It is a fact that no propeller can withstand the tensile force (and
stress) that isgenerated when it is turned at the same speed of the turbine.
2. Owing to the rotation of the propeller, the relative velocity at the propeller tip will
approach the speed of sound before the aircraft approaches the speed of sound. This
compressibility effect when approaching the speed of sound limits the design of
propellers. At high subsonic flight speeds (M > 0.7), the tips of blades may approach
supersonic speeds. If this happens, the flow may separate and shock waves may form.
As a consequence, the performance of turboprop engine deteriorates due to both the
poor propeller efficiency and the decrease in air flow rate into the engine.

The propeller is pitch controlled to be suitable for a wider range of satisfactory applications.
If the shaft of a free turbine is used to drive something other than an aircraft propeller, the
engineis called a turboshaft engine. This is one of the turboshaft engines that will be discussed
later in thischapter. Turboshaft engines are similar to turboprop engines, except that the hot
gases are expandedto a lower pressure in the turbine, thus providing greater shaft power and
low exhaust velocity.
Examples of turboshaft engines are those used in helicopters.
Now, let us discuss the advantages of turboprop engines:
1. Turboprops have high fuel efficiency, even greater than turbofan engines. This is due to
thesmall amount of air flow burned inside the engine. Turboprop engines can then
generatea lot of thrust at low fuel consumption.
2. Turboprop engines may find application in vertical takeoff and landing (VTOL).
TheOsprey V-22 aircraft as shown in Figure 6.5 is one of the famous VTOL aircraft
that ispowered by a turboprop engine.
3. Turboprop engines have high takeoff thrust that enables aircraft to have a short
fieldtakeoff.
4. They have the highest propulsive efficiency for flight speeds of 400 mph compared
toturbofan and turbojet engines.
However, turboprop engines have several disadvantages:
1. The noise and vibration produced by the propeller is a significant drawback.
1. Turboprop engines are limited to subsonic flights (< 400 mph) and low altitudes
(below30,000 ft).

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 2. The propeller and its pitch control mechanism as well as the power turbine
contributeadditional weight, so the turboprop engine may be 1.5 times as heavy as a
conventionalturbojet of the same gas generator size.

CLASSIFICATION OF TURBOPROP ENGINES

1. Based on engine–aircraft configuration
Turboprop engines may be either of the tractor (sometimes identified as puller) or pusher types.
Bythe word puller (tractor) it is meant a turboprop engine with a propeller that precedes the
intake andcompressor. The thrust force (mostly generated by the propeller) is thus a pulling
force. Most aircraftare powered by puller (tractor) configuration; Figure 1.13, for example. If
the propeller is downstreamthe inlet and compressor then this turboprop is identified as pusher
as shown in Figure 1.14.

Fig 1.13:Russian TU-95.

Fig 1.14: Pusher configuration of Jetcruzer 500, AASI airplane.

.

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Advantages of pusher turboprop engines are as follows:
a. A higher quality (clean) airflow prevails over the wing.
b. Engine noise in the cabin area is reduced.
c. The pilot‘s front field of view is improved.
Disadvantages are as follows:
a. The heavy gearbox is at the back, which shifts the center of gravity rearward and
thusreduces the longitudinal stability.
b. Propeller is more likely to be damaged by flying debris at landing.
c. Engine cooling problems are more severe.
The ―pusher‖ configuration is not very common.

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Problem 1: A gas turbine operating at a pressure ratio of 11.314 produces zero net work output
when 473.35 kJ of heat is added per kg of air. If the inlet air temperature is 300 K and the
turbine efficiency if 71%, find the compressor efficiency.

Since the net work output is zero

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Problem 2:An aircraft flies at a Mach number of 0.75 ingesting airflow of 80 kg/s at an
altitude where the ambient temperature and pressure are 222 K and 10 kPa, respectively. The
inlet design is such that the Mach number at the entry to the inlet is 0.60 and that at the
compressor face is 0.40. The inlet has an isentropic efficiency of 0.95. Find (a) the area of the
inlet entry (b) the inlet pressure recovery (c) the compressor face diameter.

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Problem 3:A turbojet engine operates at an altitude where the ambient temperature and
pressure are 216.7 K and 24.444 kPa, respectively. The flight Mach number is 0.9 and the inlet
conditions to the convergent nozzle are 1000 K and 60 kPa. If the nozzle efficiency is 0.98, the
ratio of specific heat is 1.33, determine whether the nozzle is operating under choked condition
or not. Determine the nozzle exit pressure.

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