DEPARTMENT OF AERONAUTICAL
MR.SURESH CHANDRA KHANDAI
FUNDAMENTALS OF GAS TURBINE ENGINES
• Comprehend the thermodynamic processes
occurring in a gas turbine.
• Comprehend the basic components of gas
turbine engines and their basic operation.
• Comprehend the support systems associated
with gas turbine engines.
ADVANTAGES OF GTE’s
Weight reduction of 70%
Reduced manning requirements
Quicker response time
DISADVANTAGES OF GTE’s
Many parts under high stress
High pitched noise
Needs large quantities of air
Large quantities of hot exhaust (target)
Cannot be repaired in place
• Supplies high pressure air for combustion
• Compressor types
– Radial/centrifugal flow compressor
– Axial flow compressor
• Radial/centrifugal flow
– Adv: simple design, good
for low compression ratios
– Disadv: Difficult to stage,
• Axial flow
– Good for high compression
– Most commonly used
THE THRUST EQUATION
FACTORS AFFECTING THRUST
METHODS OF THRUST AUGMENTATION
INJECTION OF WATER & ALCOHOL MIXTURE
BLEED BURN CYCLE
SUBSONIC & SUPERSONIC INLETS FOR
Inlets are very important to the overall jet
engine performance & will greatly
influence jet engine thrust output.
The faster the airplane goes the more
critical the inlet duct design becomes.
Engine thrust will be high only if the inlet
duct supplies the engine with the required
airflow at the highest possible pressure.
The nacelle/duct must allow the engine to
operate with minimum stall/surge
tendencies & permit wide variation in
angle of attack & yaw of the aircraft.
For subsonic aircraft, the nacelle shouldn’t
produce strong shock waves or flow
separations & should be of minimum
weight for both subsonic & supersonic
For certain military applications, the radar
cross sectional control or radar reflectance
is a crucial design requirements.
Inlet ducts add to parasite drag skin friction+
viscous drag) & interference drag.
It must operate from static ground run up to
high aircraft Mach number with high duct
efficiency at all altitude, attitudes & flight
It should be as straight & smooth as
possible & designed such a way that
Boundary layer to be minimum.
It should deliver pressure distribution evenly
to the compressor.
Spring loaded , blow-in or such-in-doors are
sometimes placed around the side of the
inlet to provide enough air to the engine at
high engine rpm & low aircraft speed.
It must be shaped such a way that ram
velocity is slowly & smoothly decreases
while the ram pressure is slowly &
The duct pressure efficiency ratio is defined
as the ability of the duct to convert the
kinetic or dynamic pressure energy at the
inlet of the duct to the static pressure
energy at the inlet of the compressor
without a loss in total pressure . It is in
order of 98% if there is less friction loss.
RAM RECOVERY POINT
The ram recovery point is that aircraft speed
at which the ram pressure rise is equal to
the friction pressure losses or that aircraft
speed at which the compressor inlet total
pressure is equal to the outside ambient
A good subsonic duct has 257.4 km/h.
SINGLE ENTRANCE DUCT
VARIABLE GEOMETRY DUCT FOR
NORMAL SHOCK RELATION
OBLIQUE SHOCK RELATIONS
Combustion in normal, open cycle, gas turbine is a
continuous process in which fuel is burned in the
air supplied by the compressor, an electric spark
is required only for initiating the combustion
process and thereafter the flame must be selfsustaining.
Over the years combustion chamber
geometry has evolved considerably with
respect to the objective of:
1.Improving flame stability both at sea
2.Reducing chamber size (while still burning all the
fuel and maintaining reasonable pressure drop)
3.Reducing emissions of the oxides of nitrogen,
carbon monoxide and unburned hydrocarbons.
4. Increasing chamber life.
5. Controlling the temperature distribution at inlet to
TYPES OF COMBUSTION CHAMBER
• CAN TYPE
• CAN-ANNULAR TYPE
• ANNULAR TYPE
• The earliest aircraft engines made use of Can
(or tubular) combustor in which the air leaving
the compressor splits into a number of separate
streams, each supplying a separate chamber.
• These chambers are spread around the shaft
connecting the compressor & turbine, each
chamber having its own fuel jet fed from a
common supply line.
• This arrangement was well suited to engines
with centrifugal compressors, where the flow
was divided into separate streams in the
diffuser. Example Rolls-Royce Dart
• Development could be carried out on a single
can using only a fraction of the overall airflow
and fuel flow.
• Ease of control of the fuel air ratio and simplicity.
• Low cost of replacement of a damaged liner.
• Relatively large size & weight of the chamber
• Relatively large pressure drop & need to
provide numerous igniters.
• It is not easy to provide nearly simultaneous
ignition of all chambers.
• The disadvantages have out weighted the
advantages & tubular chambers are now seldom
used for large gas turbines for aircraft
circumferentially uniform fuel-air ratio & outlet
• A failure of the liner in one spot means
replacement of relatively expensive component.
• There tend to be a heavy buckling load, due to
thermal expansion on the outer surface of the
• Requires large airflow rates during testing.
• Compact dimensions, in which maximum use is
made of the space available within the specified
diameter , this should reduce the pressure loss
& results in an engine of minimum diameter.
• It is more difficult to obtain an even fuel/air
distribution & an even outlet temperature
• Structurally weaker & it is difficult to avoid
buckling of the hot flame tube walls.
• For testing full air mass flow is required.
• Annular combustors are universally used in
modern aircraft engines. Some of the examples
are, The Olympus 593. PT-6, PW 530, V2500
FACTOS AFFECTING COMBUSTOR
1. The temperature of the gas after combustion
must be comparatively low to suit the highly
stressed turbine materials. Development of
improved materials & method of blade cooling,
temperatures to raise from about 1100K to as
much as 1850K for aircraft applications.
2. At the end of the combustion space the
temperature distribution must be of known form
if the turbine blades are not suffer from local
overheating. In practice, the temperature can
increase with radius over the turbine annulus,
because of the strong influence of temperature
on allowable stress and the decrease of blade
centrifugal stress from toot to tip.
3.Combustion must be maintained in a stream of
air moving with a high velocity in the region of
30-60m/s, and stable operation is required over
a wide range of air/fuel ratio from full load to
4. The formation of carbon deposits must be
avoided, particularly the hard brittle variety.
Small particles carried into the turbine in the high
velocity gas stream can erode the blades and
block cooling air passages. Aerodynamically
excited vibration the combustion chamber might
cause sizeable pieces of carbon to break free
resulting in even worse damage to the turbine.
5. In aircraft gas turbines, combustion must also be
stable over a wide range of chamber pressure
because of the substantial change in this
parameter with altitude &forward speed. It also
capable of relighting at high altitude in the event
of an engine flame out.
6.Avoidance of smoke in the exhaust is of major
importance of all types of gas turbine. Smoke
trails in flight were a problem for military aircraft,
permitting them to be seen from a great
7. Although gas turbine combustion systems
operate at extremely high efficiencies, they
produce pollutants such as oxides of nitrogen,
carbon monoxides and unburned hydrocarbons
and these must be controlled to very low levels.
ZONES OF THE COMBUSTOR
• The function of recirculation zone is to
evaporate, partly burn, and prepare the fuel for
rapid combustion within the remainder of the
• The function of the burning zone to burn all the
• The function of dilution zone is solely to mix the
hot gas with the dilution air.
The main factors of importance in assessing
combustion chamber performance are
•Outlet temperature distribution
Overall stagnation Pressure loss can be regarded
as the sum of the fundamental loss( a small
component which is a function of T02/T01) & the
Pressure loss factor,
1.With the help of swirl vanes surrounding the fuel
nozzle, strong vortex flow occurs in the combustion air
in the combustion region.
2.A low pressure region is created at the combustor axis,
which causes recirculation of the flame toward the fuel
3.At the same time, radial holes around the liner supply
air to the center of the vortex, making the flame grow
to some extent.
4.Jet angles and penetration from the holes are such
that jet impingement along the combustor axis results
in upstream flow.
5.The upstream flow forms a torroidal recirculation zone,
Flame stabilization created by impinging jets and general airflow pattern.
1.Use of rich primary zone in which little NO formed,
followed by rapid dilution in the secondary zone.
2.Use of a very lean primary zone to minimize peak
flame temperature by dilution.
3.Use of water or stream admitted with the fuel for
cooling the small zone downstream from the fuel
4.Use of inert exhaust gas recirculated into the reaction
5.Catalytic exhaust cleanup.
BASIC REQUIREMENTS OF FUELS
4.Deposition and fouling tendencies
The losses associated with subsonic ramjet combustion
can be substantial.
If ramjets are applied to hypersonic flight, additional
problems arise because of extremely high temperature
at the entrance of the combustion chamber.
This not only makes vehicle cooling very difficult, but it
leads to severe combustion loss due to dissociation.
For hypersonic flight (M>8) the temperature of the air in
the chamber is quite dependent on the pressure.
The higher the pressure, the less dissociation and the
higher the temperature of the mixture.
The temperature of the combustion products is likewise
For combustion pressure of 10atm and a flight Mach
number of 10, there is no temperature rise due to
All of the combustion energy is absorbed by dissociation.
At sufficiently high Mach number, the temperature of the
combustion products can be lower than that of the
Consideration with the speed with which the fuel and air
can be converted into dissociation products may show
that there is sufficient residence time in the
combustion chamber to approach equilibrium
During the subsequent expansion in the nozzle, it is
quite possible that the expansion will be too rapid for
the composition to readjust, after each step of
temperature and pressure reduction, to a new
If the expansion is extremely rapid, the mixture may be
effectively “frozen” with the initial(high temperature)
This would mean that little of the combustion energy of
the fuel would be available for acceleration of the
combustion product to provide thrust.
The chemical kinetics of the recombination processes in
the nozzle have a strong effect on the thrust and
propulsive efficiency of the engine.
1.For resistance against fatigue, Nimonic 75 has been used with
Nimonic 80 and Nimonic 90.
2.Nimonic 75 is an 80-20 nickel-chromium alloy stiffened with a
small amount of titanium carbide.
3.Nimonic 75 has excellent oxidation and corrosion resistance at
elevated temperatures, a reasonable creep strength, and good
fatigue resistance. It is easy to press, draw and mold.
4.HA-188, a Cr, Ni-based alloy, has been employed in the latter
section of some combustion liners for improved creep rupture
5.In addition to the base material changes, many of today’s
combustors also have Thermal Barrie Coatings(TBCs), which
have an insulation layer of the total thickness used is 0.4-0.6mm
and based on ZrO2-Y2O3 and can reduce metal temperature by
2 MARK QUSTIONS
1. What is need for supersonic combustion?
2. Define equivalence ratio and stoichiometric fuel air
3. Define efficiency of the combustor.
4. What is the purpose of primary air in combustion
5. What is the purpose of secondary air in combustion
6. What is the purpose of dilution air in combustion
7. Define combustion intensity?
8. State the advantages and disadvantages of annular
LONG TYPE QUESTIONS
1.What are the important factors affecting combustor design?
2.Write down the methods of flame stabilization and explain with
3.What are the three types of combustion chamber? Compare its
advantages and disadvantages.
4.Name the material used for combustion chamber and discuss
the special qualities of the material used for combustion
5. With the aid of a simplified picture explain the operation of a
6.With a neat sketch explain the working of a combustion
7. Consider n-decane fuel, balance the chemical equation for the
stoichiometric combustion of this fuel in air and find the
stoichiometric fuel-to-air ratio.
• The primary objective of a nozzle is to
expand the exhaust stream to atmospheric
pressure, and form it into a high speed jet
to propel the vehicle. For air breathing
engines, if the fully expanded jet has a
higher speed than the aircraft's airspeed,
then there is a net rearward momentum
gain to the air and there will be a forward
thrust on the airframe.
• Many military combat engines incorporate
an afterburner (or reheat) in the engine
exhaust system. When the system is lit,
the nozzle throat area must be increased,
to accommodate the extra exhaust volume
flow, so that the turbo machinery is
unaware that the afterburner is lit. A
variable throat area is achieved by moving
a series of overlapping petals, which
approximate the circular nozzle crosssection.
• At high nozzle pressure ratios, the exit pressure is
often above ambient and much of the expansion
will take place downstream of a convergent nozzle,
which is inefficient. Consequently, some jet engines
(notably rockets) incorporate a convergentdivergent nozzle, to allow most of the expansion to
take place against the inside of a nozzle to
maximise thrust. However, unlike the
fixed con-di nozzle used on a conventional rocket mo
, when such a device is used on a turbojet engine it
has to be a complex variable geometry device, to
cope with the wide variation in nozzle pressure ratio
encountered in flight and engine throttling. This
further increases the weight and cost of such an
• The simpler of the two is the ejector nozzle,
which creates an effective nozzle through a
secondary airflow and spring-loaded petals. At
subsonic speeds, the airflow constricts the
exhaust to a convergent shape. As the aircraft
speeds up, the two nozzles dilate, which allows
the exhaust to form a convergent-divergent
shape, speeding the exhaust gasses past Mach
1. More complex engines can actually use a
tertiary airflow to reduce exit area at very low
speeds. Advantages of the ejector nozzle are
relative simplicity and reliability. Disadvantages
are average performance (compared to the other
nozzle type) and relatively high drag due to the
secondary airflow. Notable aircraft to have
utilized this type of nozzle include the SR-71,
Concorde, F-111, and Saab Viggen
1-D ANALYSIS OF GAS
MASS FLOW RELATION
– Draws in air & compresses it
• Combustion Chamber
– Fuel pumped in and ignited to burn with
– Hot gases converted to work
– Can drive compressor & external load
• Controlling Load on Compressor
– To ensure maximum efficiency and allow for
flexibility, compressor can be split into HP &
– Vane control: inlet vanes/nozzle angles can
be varied to control air flow
• Compressor Stall
– Interruption of air flow due to turbulence
USE OF COMPRESSED AIR
• Primary Air (30%)
– Passes directly to combustor for combustion
• Secondary Air (65%)
– Passes through holes in perforated inner shell
& mixes with combustion gases
• Film Cooling Air (5%)
– Insulates/cools turbine blades