Gas turbines, also known as combustion turbines, are common prime
movers for many applications. Their historically high fuel consumption,
especially for small units (less than 10,000 kW), as well as at part load, and at
high inlet air temperatures has made them less desirable than diesel engines for
prime power plant applications. They have been utilized extensively in standby
and peaking applications where their relatively low installed cost outweighs
other factors. Open or simple cycle gas turbines are used in virtually all power
plant applications. Combined cycle systems, where heat is recovered from the
gas turbine exhaust and used to make steam which then drives a steam turbine,
have become much more common in recent years.
Gas turbines are based on the Brayton or Joule cycle which consists of four
processes: compression with no heat transfer, heating at constant pressure,
expansion with no heat transfer, and in a closed cycle system, cooling at
constant pressure. In open cycle gas turbines, the fourth step does not exist
since inlet air is taken from the atmosphere and the exhaust is dumped to
atmosphere. Due to its higher temperature, there is more energy available from
the expansion process than is expended in the compression. The net work
delivered to drive a generator is the difference between the two. The thermal
efficiency of the gas turbine is a function of the pressure ratio of the
compressor, the inlet temperature of the power turbine, and any parasitic losses
(especially the efficiency of the compressor and power turbine). Practical
limitations on thermal efficiency due to losses and materials technology yield a
maximum of about 40 percent at pressure ratios of 30 to 40 and temperatures
of approximately 2,500°F. These temperatures and pressure ratios are found
only in recently developed, large gas turbines. Typically pressure ratios of 5 to
20 and turbine inlet temperatures from 1,400 to 2,000°F are common in gas
turbines for this application, resulting in efficiencies from 20 to 33 percent. As
improved materials and cooling technologies are introduced to smaller units,
the efficiencies can be expected to improve if the cost is not prohibitive.
1.1. Theory of Operation
(Fig. 1) shows the processes that take place. To begin the cycle, air is
drawn in from atmosphere, through an air filter, into the compressor. As the
air passes through the compressor, pressure is raised within the range 150 - 250
psi, depending upon the particular unit. Where the pressure ratio is 10:1 the
pressure of the air as it leaves the compressor is 10 Atmospheres that is about
147 PSIA, 1035 Kilopascals or 10.35 Bar. The temperature of the air will also
rise as it is compressed typically to about 450°F (230°C). Atmospheric pressure
is about 14.7 PSIA. Air or gas pressures may be indicated in psi, kg per square
centimeter, kilopascals, atmospheres or bars. Appendix 1 has more details.
The compressed air is routed into the combustion section, where it is mixed
with incoming fuel. The fuel may be oil, or natural gas, or even gasified coal in
As a result of combustion, the temperature of the gas entering the turbine is
raised to say 1,800°F (980°C). On large modem machines the turbine inlet gas
temperature may be as high as 2,400°F, (1.300"C). As the gas passes through
the turbine, the temperature and pressure fall as energy is given-up to the
turbine to perform mechanical work-The turbine shaft is coupled to drive a
load such as an electric generator. Eventually the gas discharges at close to
atmospheric pressure and a temperature of, say, 1,000°F (537°C). These values
are given merely as an example. Make sure that you know the specific figures
related to your particular machine. The cycle described in this paragraph is
often called the simple cycle or the Brayton cycle.
The efficiency of the Brayton cycle depends upon the pressure ratio that is the
ratio between the pressure of air entering the compressor and the air leaving
see (Fig. 2). The compressor absorbs a large amount of energy; typically 50%
to 65% of the turbine output may be used to drive the compressor.
On most large gas turbines used for power generation the compressor is
mounted en the same shaft as the turbine. However on smaller units it is quite
common to have a two shaft arrangement as shown in (Fig. 3), in this
machine both the compressor and the generator are driven by separate
turbines. This has a certain advantage as it allows the compressor to run at
much higher speed where greater efficiency may be obtained.
The exhaust gas leaving the turbine contains a lot of heat. In order to recover
some of this heat, a heat recovery steam generator (HRSG) may be placed in
the gas passage. The aim is to lower the gas exhaust temperature to about
500°F or lower if possible, and produce steam in the process. This steam,
typically in the range of 100 to 1000 PSI, can then be used for industrial
purposes, district heating, or some other heating requirements.
One of the best applications is to use the steam to drive a steam turbine
generator. This is the well known combined cycle. (Fig. 4) show a combined
cycle arrangement where two 100 MW gas turbine generators provide exhaust
heat to run one 100 MW steam turbine generator. The advantage of this
arrangement is that the overall efficiency, (electrical energy out. compared with
heat energy in) is higher than either the standard steam turbine cycle or the gas
turbine cycle alone. The combined cycle also offers an economic gain as the
capital cost of installing a gas turbine generator, is much less than the
equivalent steam turbine installation, with its associated boiler and auxiliary
The majority of gas turbines currently in operation, range in size all the way
from 2 or 3 MW up to about 100 MW. A number of larger units under
construction range up to 150 and even 200 MW.
Many of the smaller units, say 15-20 MW in capacity are installed as stand-by
peaking units or to provide black-start capability. The gas turbines in these
units are usually multi-shaft derivatives of aircraft engines, typically Rolls Royce
or Pratt and Whitney. Although these units burn expensive light oil, they have
the advantage of being able to start-up quickly. This is ideal for stand-by
Aircraft type turbines are also used to drive larger generators say 15 to 30 MW.
In this arrangement several turbines drive one generator.
For continuous operation gas turbines are usually heavier single shaft
1.2. Turbines Configuration
Gas turbines are lightweight in comparison to diesel engines, are very
compact, and due to their small, well-balanced rotating mass are able to operate
at very high speeds (from 10,000 to 25,000 rpm in sizes from 900 to 10,000
kW). Smaller gas turbines are usually single-shaft design that is the compressor
and power turbine are mounted on the same shaft. Larger gas turbines are
frequently two-shaft machines in which the power turbine is divided into two
sections, one of which drives the compressor and the other which drives the
generator. The two-shaft design allows the compressor section to be operated
at a variable speed (within limits) thus varying the flow to the power turbine
section as a function of load.
1.3. Turbines Designation
General electric designates its gas turbines according to frame size as
shown in (Fig. 5). It shows also a table of Westinghouse and ABB gas-turbine
ratings and model numbers. (Fig. 6) shows typical gas-turbine-generator
layouts for coupling at the hot or cold end.
Of course there are many other manufacturers offering similar size models
such as A.B.B., Mitsubishi, Solar, M.A.N., G.E.C., and so on. Most of these
units have similar, but not identical, features. This program cries to cover all of
the generic items associated with different manufacturers of gas turbines.
When comparing different units, the specified values of efficiency will often be
quoted. Usually around 35% to 40%. Take great care when comparing quoted
values of efficiency, to check that the numbers use the same basis of
measurement. Many manufacturers apply the "lower heating value" (T-HV) of
fuel and consequently arrive BE a higher efficiency number. Both natural gas
and oil (the major fuel sources for gas turbines) contain a considerable amount
of hydrogen. When combustion takes place this hydrogen mixes with oxygen
to form steam and heat is released. This steam passes through the turbine and
eventually exhausts to atmosphere. The steam is at too high a temperature to
condense and therefore the latent heat is not recovered (about 900 BTU per
pound). The value of this unrecoverable heat can be subtracted from the
measured calorific value of the fuel to give us the so called lower heating value
(LHV). (Fig. 7) shows typical heating values. Most North American standards
recommend that the higher heating value (HHV) be used to give a realistic
comparison of heat in against energy out.
Here are some acronyms you may find in manufacturers' literature. STAG is
G.E.'s proprietary name for their combined cycle power plant. It means steam
and gas. STIG is another G.E. term, used to describe a steam injected $as
turbine. Westinghouse uses the term PACE to describe their combined cycle
plants. PACE is derived from power and combined efficiency.
In this arrangement steam is injected into the gas stream as it leaves the
combustion chambers. It actually increases the mass flow through the turbine
and so increases the work performed at the turbine. This should not be
confused with steam which is injected into the combustion chamber for
control of nitrous oxides. The subject of "NO" and other environmental
considerations are discussed in other videotapes in this series.
Figure 1. Basic Gas Turbine
Figure 2. Brayton Cycle Efficiency VS Pressure Ratio
Figure 3. Two Shaft Arrangement
Figure 4. Combined Cycle
2. DESIGN AND CONSTRUCTION
2.1 Principles of Operation
The air entrance is usually identified as engine station one.
Understanding the function of the inlet duct and its importance to engine
performance makes it a necessary part of any on gas turbine design and
The air inlet to a turbine engine must furnish a uniform supply of air to the
compressor so the compressor can operate stall-free, and it must cause as little
drag as possible. It takes only a small obstruction to the airflow inside the duct
to cause a significant loss of efficiency. If the inlet duct is to deliver its full
volume of air with a minimum of turbulence, it must be maintained as closed
to its original Condition as possible, and any repair to the inlet duct must retain
its smooth aerodynamic shape.
The engine-drive gearbox is the main unit in accessory section.
Another common gearbox location is on the front or rear of the engine if the
inlet and exhaust locations permit. This location is particularly desirable
because it allows the narrowest engine diameter and thus the lowest drag
As a secondary function, the main gearbox acts as a collection point for oil
scavenged from the engine before it is pumped back to the oil tank. This allows
many of the internal gears and bearings to be lubricated by splash lubrication.
2.3 Compressor Section
The compressor section of a turbine engine houses the compressor rotor
and the stator vanes and it supplies air in sufficient quantity to satisfy the needs
of the combustor. The primary purpose of the compressor is to increase the
pressure of the mass of air entering the engine inlet and then to discharge it
into the diffuser and the combustors at the correct velocity, pressure and
temperature. The problems associated with these requirements are great
because the compressor must move air at a velocity of around 400 to 500 feet
per second and increase its pressure by perhaps 20 to 30 times in a space of
only a few feet.
The secondary purpose of the compressor is to supply engine bleed air to cool
the internal hot section, and supply heated air for inlet anti-icing. Air is also
extracted for such aircraft uses as cabin pressurization, air conditioning, fuel
system deicing heat, pneumatic engine starting, and various other functions that
require compressed air.
3. CENTRIFUGAL COMPRESSORS
The centrifugal compressor, sometimes referred to as a radial outflow
compressor is the oldest design and it still in use today Many of the smaller
flight engines as well as the majority of gas turbine auxiliary power units use
this design .
A centrifugal compressor performs its duties by receiving the air at its center
and accelerating it outward by centrifugal force. The air then expands into a
divergent duct called a diffuser, and as it spreads out, it slows and its static
Centrifugal compressors consist basically of an impeller rotor, a diffuser, and a
manifold. The impeller is usually forged from aluminum alloy, and can be either
single- or double-sided.
The diffuser acts as a divergent duct in which the air spreads out slows down
and increases I static pressure. The compressor manifold distributes the air in a
smooth flow to the compressor section.
A single-stage, dual-side impeller allows a high mass airflow from a small
diameter engine, and it has been used in a number of flight engines in the past
for those reasons. This design does not, however, receive the full benefit from
ram effect because of the corners the air must turn as it enters and leaves the
A single-sided impeller does benefit from ram intake and its less turbulent air
entry makes it well suited for aircraft installation.
Compression ratios attainable are about the same for both of the single-stage
types of centrifugal impellers. More than one stage of compression can be used
but the use of more than two stages of single-entry compressors is considered
impractical. The energy lost in the airflow as it slows down to make the turns
from one impeller to the next, the added weight, and the amount of power
needed to drive the compressor all seem to offset the benefits of additional
compression by using more than two stages.
The most generally used centrifugal compressor is the single-sided type with
either one or two stages.
Recent developments in centrifugal compressors have produced compression
ratios as high as 15:1. In past, pressures this high could be obtained only with
axial flow compressors.
Centrifugal compressors are shorter than axial flow compressors and because
of their spoke-like design they can accelerate air faster and immediately diffuse
it in the direction of flow. Tip speed of centrifugal impeller may reach speeds
as high as Mach I .3, but the pressure within the compressor casing prevents
airflow separation and provides a high transfer of energy into the airflow.
A centrifugal compressor may be used in combination with an axial
compressor, and this is done in some of the smaller flight engines, but all of the
larger engines today use axial flow compressors.
3.1 Advantages of centrifugal compressors
High pressure rise per stage-up to 10:1 to 15:1 in a dual stage.
Good efficiency over a wide rotational speed range, idle to
approximately Mach 1.3 tip speed.
Simplicity of manufacture and relatively low cost.
Low starting power requirements.
3.2 Disadvantages of a centrifugal compressor
a) Large frontal area for a given airflow.
b) More than two stages are not practical because of the energy
losses between the stages.
4. AXIAL FLOW COMPRESSORS
There are three types of compressors: single-spool, dual- spool, and
triple-spool. Single- and dual-spool compressors are used in turbojet and turbo
shaft engines while dual and triple- spool compressors are commonly used in
The front compressor is referred to as the low pressure low speed, or N 1
compressor. Its turbine is referred to in the same manner. The rear compressor
is called the high pressure, high speed, or N2 compressor.
The rotor arrangement we see in Fig. 19 is such that the fan is referred to as the
N1, or low speed compressor, the compressor next in line is called the N 2, or
intermediate compressor and the innermost compressor is the high pressure, or
Dual-and triple-spool compressors were developed for the operational
flexibility they afford the engine in the from of high compression ratios, quick
acceleration, and better control of the stall characteristics.
For any given power lever setting, the high pressure compressor speed is held
relatively constant by the fuel control governor. And, assuming that there is a
fairly constant energy level available at the turbine, the low pressure
compressors will speed up or slow down with changes in the aircraft inlet
conditions resulting from atmospheric changes or flight maneuvers. The N 1
compressor tries to supply the N2 compressor with a fairly constant air pressure
for each power setting by speeding up or slowing down to maintain a constant
mass airflow at the inlet of N2.
Low pressure compressors will speed up as altitude is gained, as the
atmosphere is less dense and more speed is needed to force to needed amount
of air through the engine. Conversely , as the aircraft descends, the air becomes
more dense and easier to compress so the N1 compressor slow down.
Figure shows a geared fan-type engine. This compressor was developed
for smaller engines so higher turbine speeds could be converted to
torque to drive the fan. Also since the fan is geared to the compressor,
the compressor is not restricted to fan speed.
Fan tip speed may be allowed to exceed Mach one so the compressor can
deliver the correct amount of air. The pressure within the fan duct helps to
retard airflow separation from the blades at speeds over Mach one so there is
an effective transfer of energy to the air at the required compression ratio.
There are several advantages of the axial flow compressor. They are:
High peak efficiencies from ram, created by its straight- through design.
High peak pressure attainable by addition of compression stages.
Small frontal area and resulting low drag.
4.1 The disadvantages of the axial flow compressor
Difficulty of manufacture and high cost.
Relatively high weight.
High starting power requirements.
Low pressure rise per stage, approximately 1.27:1.
Small gas turbine engines used in business jet aircraft may have a compression
ratio in the order of 6:1 in older models, up to as high as 18:1 in the newer
designs. By comparison, the engine used in some of the jumbo jets will
compress the air as much as 30 atmospheres.
Compression ratio is found by comparing the discharge pressure of the last
stage of compression with ambient air pressure. For example, if the ambient
pressure is 14.7 psi and the compressor static discharge is 97 psi, the
compression ratio is expressed as 6.6:1.
A typical example of the compression ratio of a dual-spool compressor may be
N1 compression ratio = 3:1.
N2 compression ratio 4:1
Total compression ratio = 12:1
Note that the compression ratio of one compressor is multiplied by the other
to get the total compression ratio. Normally N2 will turn at a higher speed than
N1, and because of its small diameter, it will have a higher compression ratio.
Therefore, a dual compressor having 10 stages would have a higher
compression ratio than 10 stages in single-spool compressor. This is one of the
principle advantages of dual -spool compressors.
Compression ratio of the fan section may be found by dividing the fan
discharge pressure by the fan inlet pressure. An example of compression ratio
of a fan might be:
inlet pressure = 14.7 psi.
Fan discharge pressure = 26.5 psi.
Fan compression ratio = 1.8: 1
The rated compression ratio of a gas turbine engine is calculated at standard
conditions and full takeoff power and any deviation from these conditions will
affect the discharge pressure.
If the mass airflow eating for given engine is 50 pounds per second under
standard-day conditions at full power, the mass airflow will increase as the
compressor discharge pressure increases. In fact, the mass airflow change
which occurs after takeoff will require the pilot to reduce power to keep from
over boosting the engine. The climb to altitude will result in a drop in ambient
pressure, but an airplane flying at an altitude where the ambient pressure is only
about one third of sea level pressure has ram compression in the engine inlet of
approximately 1.5:1, and this helps compensate for the decreased air density.
Actually, the engine is designed to operate most efficiently with this lower mass
airflow at cruise altitude by a corresponding reduction in fuel flow.
4.2 BLADES AND VANES
An axial flow compressor has two main elements, the rotor and the
stator. The rotor blades force air rearward through each stage which consists of
one set of rotor blades and the following set of stator vanes. The speed of the
rotor determines the air velocity in each stage. As the velocity increases, kinetic
energy is added to the air. The stator vanes are placed to the rear of the rotor
blades to receive the high velocity air and act as diffusers, changing the kinetic
energy of velocity into potential energy of pressure the stators also serve a
secondary function of directing the airflow into the next stage of compression.
Compressor blades are constructed with a varying angle of incidence, or twist.
This twist compensates for the blade velocity change caused by its radius. The
further from the axis of rotation, the faster the blade section travels. The blades
also decrease in size, from the first stage to the last to accommodate the
converging, or tapering, shape of the space in which they rotate.
The length, chord, thickness, and aspect ratio (ratio of the length to the width)
to the compressor blade are designed to suit the performance factors required
for a particular engine and aircraft combination.
Axial flow compressors normally have from 10 to 18 stages of compression
with the fan considered to be the first stage of compression. Some tong fan
blades have a mid-span shroud each blade to from a circular ring which helps
support the blades against the bending forces from the air stream. The shrouds,
however, block some of the airflow, and the aerodynamic drag they produce
reduces the efficiency of the fan. The section of the fan blade, from the midspan shrouds to the root, is the compressor blade section for the core engine.
The roots of the compressor blades are often loosely fitted into the compressor
disk for ease of blade assembly and for the vibration damping it provides. As
the compressor rotates, centrifugal force keeps the blade in its correct position,
and the air stream over the airfoil provides a shock absorbing or cushioning
effect. These blades are attached to the disk with a dovetail and are secured
with a pin and a lock tab or lock wire.
Some blades are cut off square at the tip and these are referred to as flat
machine tips. Other blades have a reduced thickness at the yips and these are
called profile, or squealer, tips. All rotating machinery has a tendency to vibrate,
and profiling a compressor blade increases the natural frequency of the blade.
By raising the natural frequency of the blade beyond the frequency of rotation,
the vibration tendency is reduced.
Profiling changes the aerodynamics at the tip to produce a smooth axial airflow
even if the tip is rotating at speeds beyond the speed of sound and flow
separation has started to occur.
On some of the newer engines, the compressor rotor tips are designed to have
a tight running clearance and rotate within a shroud tip strip of abradable
material. This strip will wear away rather than cause blade damage if contact
takes place and the strip is replaced when the engine is overhauled. A high
pitch noise can be head on coast down if the compressor blade touches the
shroud tip strip and this is the reason profile tips are called squealer tips.
4.3 Variable vanes
Stator vanes may either be stationary or may have their angle: variable.
The inlet guide vanes which are the vanes immediately in front of the first stage
rotor blades may also be stationary or variable. The function of the inlet guide
vanes is to direct the airflow into the compressor at the most desirable angle.
Exit guide vanes are placed at the compressor discharge to remove the
rotational moment imparted to the air by the compressor.
5. COMBINATION COMPRESSORS
To take advantage of the points of both the centrifugal and the axial
flow compressor and eliminate some of their disadvantages, the combination
axial-centrifugal compressor was deigned. This application is currently being
used in many small turbine engines installed in business jet airplanes and
The combination compressor is especially well suited to engines using reverseflow annular combustors.
The engine diameter is wider to accommodate this type of combustor so there
is no disadvantage in using the centrifugal compressor which, by the nature of
its design, is much wider than a comparable axial flow compressor.
6. INTER-STAGE AIRFLOW
Compressor airfoils experience an infinite variety of angles of attack and
air densities, and controlling the angle of attack is a design function of the inlet
duct, the compressor and the fuel control sensors.
The rotating compressor blades speed up the air in the inlet duct and the air
passes through the inlet guide vanes, which changes its angle of flow, but does
not change either its velocity or its pressure. The amount the inlet guide vanes
change the angle of the air entering the compressor is determined by their
position and by the curvature of the vanes. Note that the entering and exiting
arrows are the same length, showing that there has been no change in velocity,
but only a change in direction.
There are two vector forces acting on airflow. One vector is the ram effect
giving velocity to the air entering the compressor. This is shown by the arrow
labeled “inlet guide vane effect”. The other force is created by the aerodynamic
shape of the airfoil (rotor blade). Air is pulled into the compressor and flow
over the airfoil in the direction opposite to the blade rotation. This vector is
labeled “rotor speed effect.
The resultant of the two force vectors gives us the angle of attack, or the angle
between the resultant vector and the chord line of the blade. Airflow through
the compressor, in spite of the spinning rotor, is relatively straight, with no
more than about 180 degrees of rotation as the air passes through the engine.
Air leaving the last stage of compression passes through a stationary set of exit
guide vanes which straightens the air flow before it enters the combustor.
The duct, formed by the top, or cambered, side of one blade and the bottom
side of the adjacent blade is diverging in shape, and air passing through a
diverging duct has its static pressure increased, and work done on the air by the
rotating blades increases its velocity.
AS the air leaves the trailing edge of the compressor rotor blades, it flows
through a row of stator vanes which also form diverging ducts. But since no
energy is added by the stator, the velocity of the air decreases, and its static
pressure increases. The action of the compressor rotor blades and the stator
vanes continues through all of the stages of compression, and when the air
leaves the compressor it has approximately the same velocity it had when it
started, but it has a much a higher static pressure.
The air can flow rearward against the ever increasing pressure only because
energy is transferred from the turbine as it drives the compressor. When a
compressor blade or a stator vane has a positive angle of attack, the pressure
on the bottom of its airfoil shape is higher than the pressure on the top.
The high and low pressure areas formed on the airfoils allow air passing
through one stage to be influenced by the next stage. This is called the cascade
Fig. 24. shows the high-pressure area of the first stage blade being pulled into
the low-pressure area of the stator. Notice that the leading edge of the stator
vane faces in the opposite direction as the leading edge of the rotor blade. This
arrangement produces a pumping action. The high-pressure area of the firststage stator vane pumps air into the low- pressure area of the second-stage
rotor blade and so on through the compressor.
According to- Bernoulli’s principle, as pressure builds up in the rear stages of
the compressor, its velocity should decrease. To keep the velocity constant, the
shape of the path of the gas through the compressor converges.
7. ANGLE OF ATTACK AND COMPRESSOR STALL
As we see in Fig. 25, the angle of attack of the compressor blade is
determined by the inlet air velocity and the compressor RPM. These two forces
combine to form a vector force which gives us the angle of attack of the airfoil.
A compressor stall is a condition of airflow when the angle of attack becomes
A compressor stall causes the airflow to slow down, stop, or even reverse its
direction, depending upon the severity of the still. Stalls can range from a slight
air vibration, or fluttering sound, to a louder pulsating sound, or even to a
violent backfire or explosion. Quite often the gages in the cockpit do not show
a transient stall condition, and these stalls are usually not to an engine. They
often correct themselves after one or two pulsations. But severe stalls, called
hung stalls, can significantly impair engine performance, cause loss of power
and can, even damage the engine. The pilot can identify a stall con4tion by its
audible noise, by fluctuations of the RPM, by an incr1ase In the exhaust gas
temperature, or by a combination of these clues.
Compressor stalls may be caused by:
a) Turbulent or disrupted airflow to the engine inlet, which red the velocity
b) Excessive fuel flow caused by abrupt engine acceleration. This increases
the back pressure of the combustor and reduces the velocity vector.
c) Contaminated or damaged compressor blades or stator vanes.
d) Damaged turbine components which cause a loss of shaft horsepower
delivered to the compressor. This decreases the compressor speed and
reduces the velocity vector.
The remedy for an acceleration stall is to reduce the power and allow the inlet
air velocity and the engine RPM to get back into their proper relationship.
In case of a severe compressor stall or surge caused by a fuel control
malfunction or from foreign object ingestion, a reversal of airflow can occur
with such force that the compressor blades may be bent enough to cause them
to contact the stator vanes: This extreme condition will result in disintegration
of the compressor and complete engine failure.
7.1 Stall, or surge, margin curie
Another way to describe the compressor stall is by the use of a stall
surge, margin curve. A stall is a localize condition, while a surge occurs across the
entire compressor. Every compressor has a best operation condition for a given
compression ratio, speed, and mass airflow. This is commonly called the design
In (Figure 26 above), the steady-state operating line indicates that the engine
will perform without stall at the various compression pressure speeds, and mass
airflow along the line. This line falls well below the stall zone. The design point
is located on this line and-it represents the conditions under which the engine
will for most of its life-that is, at high altitude cruise
From, this curve, it may be seen that for any given compressor speed, only a
narrow band of compressor pressure ratios are acceptable. Only those between
the steady-state line and the stall zone will provide satisfactory engine
This band of compression ratios is called the stall margin.
Also, for any given mass airflow, there exists only a narrow band of
compressor pressure ratios which will allow the engine to operate stall free.
If compressor contamination, cold- or hot- section damage, incorrect fuel
scheduling or other engine malfunctions cause a significant change in any of
three parameters on the chart, a stall or occur, or if not a stall.
The inter-stage airflow we have discussed has all been at
sonic speed. But some of the newer and more advanced compressor have
supersonic airflow with speeds up to Mach 1.3 over some of the airfoil
sections of both blades and vanes in compressors of this design have thin
leading edges and a twist that forms a slightly convergent-divergent airflow
passage, similar to the C-D flight inlet duct we have discussed.
7.2 Comparison of axial flow with centrifugal flow
Axial flow and centrifugal flow compressors both raise the pressure of
the air inside the engine. The centrifugal compressor raises the pressure by
accelerating the air outward into a single divergent-duct diffuser where
according to Bernoulli’s principle the air spreads out and slows down, and its
An axial compressor raises the air pressure by accelerating the air rearward
through many small diffuser or divergent ducts formed by the shape and
position of the rotor blades and stator vanes. The trailing edges of the blades
pairs also form divergent ducts which start the rise in pressure prior to entry
into the stators.
8. DIFFUSER SECTION
The diffuser, located directly behind the compressor, provides the space
for the air leaving the compressor to spread out. It is in the form of a divergent
duct, and is usually a separate section, bolted to the compressor case. The
pressure in the diffuser is the highest in the engine, and this high-pressure gives
the combustion products something to push against.
Figure or Typical diffuser section located between the compressor and
the combustion The air is at hi heat pressure ü the diffuser, and it is
from air that bleed air is taken.
Fig 28 shows Through-flow combustors
9. COMBUSTION SECTION
The combustors, or burners, in a gas turbine engine have an outer
casing, an inner perforated liner, usually made of stamped sheet metal, a fuel
injection system, and an ignition system for starting Heat energy is added to the
flowing gases in the burners, and this energy expands the gases and accelerates
them as they leave the engine.
When heat energy from the fuel is added, the gases expand, but since the area
through which the gas must flow remains the same, the flowing gases speed up.
Most combustors are of the through-flow configuration which is sometimes
called a through flow combustor. Gases entering from the diffuser are ignited
and then pass directly through the combustor into the turbine section. The
multiple can, annular, and an-annular combustors are of this type.
Another configuration is the reverse-flow annular type where gases leaving the
diffuser flow to the rear of the combustor where fuel is sprayed in and ignited.
Then the burning gases following a reverse-S path into the turbine section.
To function properly, the combustors must mix the air and the fuel for
efficient combustion. Then it must lower the temperature of the hot
combustion products enough that they will not overheat the turbine
components. To do this, the air flow through the combustor is divided into
primary and secondary air paths. Approximately 25 to 35 percent of the air is
routed to the area around the fuel nozzle for combustion. This is the primary
air. The secondary air, or the remaining 65 to 75 percent, forms a cooling air
blanket on either side of the liner and centers the flames so they do not contact
the metal. The secondary air also dilutes and cools the hot primary air to a
temperature that will not shorten the service life of the turbine components.
Developments in recent years have brought out the smokeless or reduced
smoke combustor. Incomplete combustion in the early engines left unburned
fuel in the tail pipe where it entered the atmosphere as smoke. By shortening
the flame pattern and using new materials that can with stand higher operating
temperatures, manufacturers have been able to almost completely eliminate the
smoke emissions from turbine engines as more complete combustion occurs.
Figure 29 shows Reverse-/tow combustor
Figure shows secondary airflow in the combustor cool is burning gases
enough that they will not damage.
The secondary air in the combustors may flow at a velocity of up to several
hundred feet per second, but the primary airflow is slowed down by swirl
vanes, which gives the air a radial motion and retards its axial velocity to about
five or ix feet per second before it is mixed with the fuel and burned. The
vortex created in the flame area provides the required turbulence to properly
mix the fuel and the air. This reduction in the airflow velocity is important
because of the slow flame propagation rate of kerosene-type fuels. If the
primary airflow velocity was too high, it would literally blow the flame out of
the engine. As it is, the combustion process is complete in the first third of the
combustor length, and the burned and unburned gases then mix to provide an
even distribution of heat at the turbine nozzle.
Although flameout is uncommon in modern engines, combustion instability
still occurs and, occasionally, a complete flameout. Turbulent weather, high
altitude) slow acceleration during maneuvers and high-speed maneuvers are
some of the typical conditions which induce combustor instability which could
lead to flameout
There are two types of flameouts: a lean flameout usually occurs at low engine
speed and low fuel pressure, at high altitude where the flame from a weak
mixture can be blown out by the normal airflow. A rich flameout occurs during
rapid engine acceleration where an overly-rich mixture causes the combustion
pressure to increase so much that the compressor airflow stagnates and slows
down or even stops. The interruption of the airflow then causes the flame to
go out. Turbulent inlet conditions and violent flight maneuvers can also cause
compressor stalls which could result in airflow stagnation and flameout.
Combustor instability sometimes causes small gas pressure fluctuations. These
low pressure cycles cause high fuel flow pulsations, which increase the
combustor instability until the pilot makes the necessary adjustments to the
flight conditions or to engine controls.
Combustor efficiency ranges between 95 and 99 percent, which means that 95
to 99 percent of the heat energy in the fuel is released. The combustor
efficiency is high, but only about one third of the mass airflow is used for
combustion, with the remainder of it used for cooling, to keep the
temperatures within acceptable limits for the combustor and the turbine.
9.1 Multiple-can combustor
This older type of combustion chamber (not commonly used today)
consists of a series of outer housings, each with its own perforated inner liner.
Each of the multiple combustor cans is actually a separate burner unit, with all
of them discharging into the open area at the turbine nozzle inlet. The
individual combustors are interconnected with small flame propagation tubes
so that when combustion starts in the two having igniter plugs, the flame will
travel through the tubes and air mixture in the other cans.
Figure 31 shows Multiple-can combustor
Figure 32 shows annular combustor
9.2 Annular Combustor
The annular combustor consists of an outer housing and a perforated
inner liner called a basket, with both parts encircling the engine. Multiple fuel
spray nozzles stick out into the basket and both primary and secondary air for
combustion and cooling flow through it in the same way as in the other
Annular combustors are in common use today in both small and large engines.
They are the most efficient type from the standpoint of both thermal efficiency
and weight, and they are also shorter than the other types. The small amount of
surface area requires less cooling air, and makes the best use of the available
space, especially for large engines where other types of combustors would be
much heavier for the large mass airflow these engines use.
9.3 Can-Annular Combustor
The can-annular combustor is used for commercial aircraft powered by
Pratt and Whitney engines. This type of combustor consists of an outer case
with multiple inner liners located radially around the axis of the engine. Flame
propagation tubes connect the individual liners and two igniter plugs are used
The combustor in figure below uses eight cans, and each can has its own fuel
nozzle supporting its forward end. The outlet duct with its eight openings
supports the cans at cans at their aft end. An advantage of this type of
combustor is its ease of on-the-wing maintenance because the forward half of
the outer combustor casing may be unfastened to slide rearward exposing the
cans for inspection.
Figure 33 shows Can-annular combustor
9.4 Reverse-Flow Annular Combustor
This design is used by the Pratt and Whitney PT6 and the Garrett TFE
731 and several of the other engines instanced in business aircraft. The reverseflow combustor serves the same function as the through-flow combustor but it
differs by the air flowing around the chamber and entering from the rear,
causing the combustion gas flow to be in the opposite direction as the normal
airflow through the engine.
Notice in Figure below that the turbine wheels are inside the combustor area
rather than in tandem, as they are with through- flow combustors. This allows
for a shorter and lighter engine and it also uses the hot gases to preheat the
compressor discharge air. These factors help make up for the loss of efficiency
caused by the gases having to reverse their direction as they pass through the
Figure 34 shows Reverse-flow combustor
10. TURBINE SECTION
The turbine section is bolted to the combustor and it contains the
turbine wheels and the turbine stators. The turbine transforms a portion of the
kinetic energy and heat energy in the exhaust gases into mechanical work, so it
can drive the compressor and the accessories.
The compressor adds energy to the air by pressure, and the turbine extracts
energy by pressure of the flowing gases.
Figure 35 shows an impulse turbine is driven by the impulse of hot gases
on the blade.
Figure shows When the hot high velocity gases flow through an in pulsar
action turbine, an aero dynaink force as well as the impulse force moves
the, blades in the direction needed to spin the wheel
This is done by converting pressure into velocity at the nozzles formed at the
training edge of the stator vanes and rotor blades. The airflow is directed in a
tangential rather than axial direction, and this slows the gas flow and reduces its
reactive power, but it adds torque power to the rotor system. The mass of the
airflow is not changed by the transfer of energy to the rotor system, but the
velocity of the air flowing through the engine is decreased as power is taken to
drive the compressor. Turboprop and turbo shaft engine have very little
reactive thrust in the tail pipe after their multi-stage turbines extract power.
An efficient turbine performs a maximum amount of work with the least fuel
consumption, and efficient turbines operate at their design point of
temperature and RPM when the compressor is operating at its design point of
compression ratio and mass airflow.
The turbine absorbs most of the energy released by the combustion, and it is
the most highly stress component in the engine. The disk, usually a heavy
forging of a nickel alloy, must withstand extremely high temperature loading
and high rotational speed.
The blades are held in the disk by a method similar to the compressor blades.
The stator vanes for the turbine are located ahead of the rotor rather than
behind the rotor, as they are in the compressor. The compressor stators act as
diffusers to decrease to velocity and increase the pressure of the gas, but the
turbine stators act as nozzles, to increase the velocity and decrease the pressure.
Most turbine nozzles operate in a choked condition from cruise to takeoff
power. This provides a fairly constant flow of energy to the turbine wheel over
its normal operating range. Since the nozzle is choked, the velocity of the gases
depends upon their temperature, which affects the local speed of sound. The
downstream pressure has little effect on a choked turbine nozzle, as it creates
its own back pressure. The turbine stator also directs the gases at the optimum
angle to the turbine blades so the wheel will turn with maximum efficiency.
The gas flowing through the turbine stator is at the highest velocity of any
point in the engine, and this velocity is controlled by the total area of the
opening between the vanes.
Turbine is classified as impulse, reaction, or a combination impulse-reaction
type. An impulse turbine causes no net change in the pressure across the
turbine wheel. The nozzle guide vanes are so shaped that they change the
direction of the gases so they will strike the turbine wheel at the correct angle
and at an increased velocity.
Reaction turbines produce their turning force by an aerodynamic action. The
nozzle guide vanes are shaped in such a way that they only direct the gas in the
correct direction; they do not increase the velocity of the gas. This gas passes
between the blades of the turbine, which form converging nozzles that further
increase its velocity. As the gases flow over the airfoil- shaped blades, a force
component in direction of the plane of rotation causes the turbine to spin.
Most turbines are neither totally of the reaction nor of the impulse type, but are
of the impulse-reaction type that causes the turbine to spin because of a
combination of the impulse pressure and the reaction force of the flowing
11. Gas Turbine Auxiliary Systems
Gas turbine auxiliary systems include starting, fuel supply, lubrication,
governor/controls, speed reduction gear, inlet air, and engine exhaust.
11.1. Starting System
Gas turbines utilize a variety of starting systems based on size of the unit
and other considerations. Common starting methods include compressed air,
direct current (DC) electric motors with dedicated batteries, or a hydraulic
pump driven by an alternating current (AC) motor, small gas turbine, or diesel
engine, which in turn drives the hydraulic motor on the gas turbine. Where
used, an auxiliary gas turbine or diesel engine also requires a starting system,
usually a DC motor and batteries. Regardless of the equipment used, the
starting system brings the unit up to a minimum speed at which the burners
may be ignited and the turbine is then brought up to operating speed.
Gas turbines installed in power plants may be started with compressed air,
DC motors, or an engine driven hydraulic system. Dedicated compressors
typically provide starting air at pressures from 150 to 500 psig, depending on
the specific requirements of the gas turbine. The system must provide adequate
storage of compressed air to allow multiple attempts to start the engines. DC
motors are driven from batteries located at the engine skid, which are charged
by a dedicated battery charger. Hydraulic systems are composed of a prime
mover, usually a diesel engine or small gas turbine, hydraulic pump, drive
motor, and accessories, including hydraulic reservoir, air cooled heat exchanger,
11.2. Fuel System
Although gas turbines are capable of burning either gas or liquid fuels,
only liquid fuels are addressed in this chapter since they are preferred for
standby power generation. The following fuel system components are
commonly provided as part of the gas turbine package: motor driven booster
pump, low-pressure duplex fuel filter, main turbine driven fuel pump, high
pressure filter, main fuel control valve (regulated by the governor), fuel
manifold and injectors at the combustor, and igniter.
The gas turbine is dependent on the fuel oil system to provide fuel to the
engine skid. The fuel oil must have the proper characteristics required for the
specific engine installation. In general, gas turbines can utilize a wider range of
liquid fuels than diesel engines. Most facilities use kerosene, No. 1 fuel oil, or
No. 1 diesel, but some use No. 2 fuel (if acceptable to the manufacturer) since
it is less expensive than the lighter grades of fuel.
11.3. Lubrication System
Most gas turbines are provided with complete lubrication systems which
include a cooler (air cooled), filter, pre/post lube pumps, engine driven main
lube oil pump, alarms, oil storage tank (located in engine skid), and heater. The
system is usually packaged with the gas turbine and only the lube oil cooler is
remotely located. The lube oil system may supply the speed reduction gear and
generator in addition to the gas turbine.
The proper lubrication of the moving parts inside a gas turbine is critical to
obtain satisfactory operation of the engine and maximum life of its
components. The lube oil must be approved by the engine manufacturer and
analyzed on a regular basis to determine the optimum interval for changing the
lube oil. Lube oil change intervals are much longer than those for diesel
engines, since the oil does not become contaminated by products of
combustion. Lube oil systems cool and filter the lube oil to provide both
proper lubrication and cooling of critical components within the engine.
The gas turbine speed and fuel flow are controlled by the governor in
response to load changes. Typically two types of governors are used on gas
turbines driving electric generators: self-contained mechanical-hydraulic type or
remote electronic governor with separate engine mounted actuator. Electronic
governor systems with load sharing capability are the usual choice for multiple
engine plants. Plants with multiple engines must have compatible governors to
ensure proper operation of engines in parallel.
The basic control of the engine is maintained by the governor during operation
and the control is independent for each engine. The overall control of a
multiple engine power plant can be relatively simple or very sophisticated.
Possible control options range from local or manual starting and
synchronization of each engine to automatic starting, synchronization, and load
sharing of the engine generators.
11.5. Speed Reduction Gear System
The high operating speeds of most gas turbines require that a speed
reduction gear be installed to drive the generator at the appropriate
synchronous speed, usually 1,200 to 1,800 rpm. The reduction gear is typically
an epicyclic design that permits a straight-through shaft arrangement, thus
simplifying alignment. A variety of epicyclic designs are used and depending on
the speed of the gas turbine, a two-stage reduction may be required. Two
common designs are the standard planetary system and the star compound
system. The reduction gear is typically lubricated by the main lube oil system.
11.6. Inlet and Exhaust Systems
Gas turbines require significantly more combustion air than diesel
engines. Flows are typically four to five times as much as that required by a
diesel engine of the same capacity. This leads to much larger air filters, intake
ducts, and exhaust ducts. Proper air filtration is critical to gas turbine
performance. Deposits on compressor and turbine blades can significantly
The engine intake and exhaust systems provide filtered air to the engine and
remove products of combustion from the engine room. These systems may be
very simple or relatively complex, incorporating such features as preheating or
pre-cooling of the intake air, or hardened design. Restrictions or blockage of
either the intake or exhaust systems will severely impact engine performance.