OVERVIEW
The major components of the engine are a compressor section, combustion section, and a turbine. The turbine is
mechanically coupled and drives the compressor by a drive shaft.
The compressor, combustor, and turbine are called the core of the engine, since all gas turbines have these components.
The core is also referred to as the gas generator (GG) since the output of the core is hot exhaust gas.
The gas is passed through an exhaust duct to atmosphere. On some types of applications, the exhaust gas is used to drive
an additional turbine called the power turbine which is connected to a piece of driven equipment (i.e. generators, pumps,
process compressors, etc).
Because of their high power output and high thermal efficiency, gas turbine engines are also used in a wide variety of
applications not related to the aircraft industry. Connecting the main shaft (or power turbine) of the engine to an electromagnet
rotor will generate electrical power. Gas turbines can also be used to power ships, trucks and military tanks. In
these applications, the main shaft is connected to a gear box.
The major components of the engine are a compressor section, combustion section, and a turbine. The turbine is
mechanically coupled and drives the compressor by a drive shaft.
The compressor, combustor, and turbine are called the core of the engine, since all gas turbines have these components.
The core is also referred to as the gas generator (GG) since the output of the core is hot exhaust gas.
The gas is passed through an exhaust duct to atmosphere. On some types of applications, the exhaust gas is used to drive
an additional turbine called the power turbine which is connected to a piece of driven equipment (i.e. generators, pumps,
process compressors, etc).
Because of their high power output and high thermal efficiency, gas turbine engines are also used in a wide variety of
applications not related to the aircraft industry. Connecting the main shaft (or power turbine) of the engine to an electromagnet
rotor will generate electrical power. Gas turbines can also be used to power ships, trucks and military tanks. In
these applications, the main shaft is connected to a gear box.
COMPRESSION – COMBUSTION – EXPANSION – EXHAUST
Four processes occur in gas turbine engines, as illustrated above. These processes, first described by George Brayton
and called the Brayton cycle, occur in all internal combustion engines. The Brayton steps are as follows:
Compression occurs between the intake and the outlet of the compressor (Line A-B). During this process, pressure and
temperature of the air increases.
Combustion occurs in the combustion chamber where fuel and air are mixed to explosive proportions and ignited. The
addition of heat causes a sharp increase in volume (Line BC).
Expansion occurs as hot gas accelerates from the combustion chamber. The gases at constant pressure and increased
volume enter the turbine and expand through it. The sharp decrease in pressure and temperature (Line C-D).
Exhaust occurs at the engine exhaust stack with a large drop in volume and at a constant pressure (Line D-A).
The number of stages of compression and the arrangement of turbines that convert the energy of accelerating hot gas into
mechanical energy are design variables. However, the basic operation of all gas turbines is the same.
Four processes occur in gas turbine engines, as illustrated above. These processes, first described by George Brayton
and called the Brayton cycle, occur in all internal combustion engines. The Brayton steps are as follows:
Compression occurs between the intake and the outlet of the compressor (Line A-B). During this process, pressure and
temperature of the air increases.
Combustion occurs in the combustion chamber where fuel and air are mixed to explosive proportions and ignited. The
addition of heat causes a sharp increase in volume (Line BC).
Expansion occurs as hot gas accelerates from the combustion chamber. The gases at constant pressure and increased
volume enter the turbine and expand through it. The sharp decrease in pressure and temperature (Line C-D).
Exhaust occurs at the engine exhaust stack with a large drop in volume and at a constant pressure (Line D-A).
The number of stages of compression and the arrangement of turbines that convert the energy of accelerating hot gas into
mechanical energy are design variables. However, the basic operation of all gas turbines is the same.
CONVERGENT AND DIVERGENT DUCTS
Compressors in gas turbine engines use convergent and divergent ducts to generate the high pressures necessary to (a)
provide a “wall of pressure,” preventing expanding hot gas from exiting through the engine inlet, as well as, through the
exhaust; and (b) provide the proper ratio of air-to-fuel for efficient combustion and cooling of the combustion chamber.
Pressure decreases through convergent ducts and increases through divergent ducts, a phenomenon which is
demonstrated in paint spray equipment. Compressed air, forced through a convergent duct, generates a lower pressure
through the narrow section to draw in paint.
Expansion through a divergent section then increases pressure and air volume, dispersing the paint in an atomized mist.
Compressors in gas turbine engines use convergent and divergent ducts to generate the high pressures necessary to (a)
provide a “wall of pressure,” preventing expanding hot gas from exiting through the engine inlet, as well as, through the
exhaust; and (b) provide the proper ratio of air-to-fuel for efficient combustion and cooling of the combustion chamber.
Pressure decreases through convergent ducts and increases through divergent ducts, a phenomenon which is
demonstrated in paint spray equipment. Compressed air, forced through a convergent duct, generates a lower pressure
through the narrow section to draw in paint.
Expansion through a divergent section then increases pressure and air volume, dispersing the paint in an atomized mist.
INLET GUIDE VANES
Inlet guide vanes direct, or align, airflow into the first rotating blade section where velocity is increased by the addition of
energy. The following stator vane section is divergent, providing an increase in static pressure and a decrease in air
velocity. Airflow then enters the second stage at a higher initial velocity and pressure than at the inlet to the preceding
stage. Each subsequent stage provides an incremental increase in velocity and static pressure until the desired level of
pressure and velocity is reached.
Some compressor stator vanes are designed to move, changing their divergence, allowing regulation of compressor outlet
pressure and velocity to achieve the proper ratio of air for fuel combustion and cooling versus engine speed and power
output.
Inlet guide vanes direct, or align, airflow into the first rotating blade section where velocity is increased by the addition of
energy. The following stator vane section is divergent, providing an increase in static pressure and a decrease in air
velocity. Airflow then enters the second stage at a higher initial velocity and pressure than at the inlet to the preceding
stage. Each subsequent stage provides an incremental increase in velocity and static pressure until the desired level of
pressure and velocity is reached.
Some compressor stator vanes are designed to move, changing their divergence, allowing regulation of compressor outlet
pressure and velocity to achieve the proper ratio of air for fuel combustion and cooling versus engine speed and power
output.
COMPRESSORS
Compressors in gas turbine engines use convergent and divergent ducts to generate the high pressures necessary to (a)
provide a “wall of pressure,” preventing expanding hot gas from exiting through the engine inlet as well as through the
exhaust; and (b) provide the proper ratio of air-to-fuel for efficient combustion and cooling of the combustion chamber.
Pressure decreases through convergent ducts and increases through divergent ducts, a phenomenon which is
demonstrated in paint spray equipment. Compressed air, forced through a convergent duct, generates a lower pressure
through the narrow section to draw in paint. Expansion through a divergent section then increases pressure and air
volume, dispersing the paint in an atomized mist.
All turbine engines have a compressor to increase the pressure of the incoming air before it enters the combustor.
Compressor performance has a large influence on total engine performance. There are two main types of compressors:
axial and centrifugal.
In the illustration, the example on the left is called an axial compressor because the flow through the compressor travels
parallel to the axis of rotation. An apparent contradiction in the operation of the axial-flow compressor is that high pressure
is generated, although the overall divergent shape would appear to cause a lower output pressure. Output pressure is
increased by divergence in each static inter-stage section. Rotating compressor blades between each static stage
increases the velocity that is lost by injecting energy.
The compressor on the right is called a centrifugal compressor because the flow through this compressor is turned
perpendicular to the axis of rotation. Centrifugal compressors, which were used in the first jet engines, are still used on
small turbojets and turbo-shaft engines. Modern large turbojet, turbofan, and turbo-shaft engines usually use axial
compressors
Compressors in gas turbine engines use convergent and divergent ducts to generate the high pressures necessary to (a)
provide a “wall of pressure,” preventing expanding hot gas from exiting through the engine inlet as well as through the
exhaust; and (b) provide the proper ratio of air-to-fuel for efficient combustion and cooling of the combustion chamber.
Pressure decreases through convergent ducts and increases through divergent ducts, a phenomenon which is
demonstrated in paint spray equipment. Compressed air, forced through a convergent duct, generates a lower pressure
through the narrow section to draw in paint. Expansion through a divergent section then increases pressure and air
volume, dispersing the paint in an atomized mist.
All turbine engines have a compressor to increase the pressure of the incoming air before it enters the combustor.
Compressor performance has a large influence on total engine performance. There are two main types of compressors:
axial and centrifugal.
In the illustration, the example on the left is called an axial compressor because the flow through the compressor travels
parallel to the axis of rotation. An apparent contradiction in the operation of the axial-flow compressor is that high pressure
is generated, although the overall divergent shape would appear to cause a lower output pressure. Output pressure is
increased by divergence in each static inter-stage section. Rotating compressor blades between each static stage
increases the velocity that is lost by injecting energy.
The compressor on the right is called a centrifugal compressor because the flow through this compressor is turned
perpendicular to the axis of rotation. Centrifugal compressors, which were used in the first jet engines, are still used on
small turbojets and turbo-shaft engines. Modern large turbojet, turbofan, and turbo-shaft engines usually use axial
compressors
COMBUSTORS
All turbine engines have a combustor, in which the fuel is combined with high pressure air and burned. The resulting high
temperature exhaust gas is used to turn the turbine and produce thrust when passed through a nozzle.
The combustor is located between the compressor and the turbine. The combustor is arranged like an annulus, or a
doughnut, as shown by illustrations above. The central shaft that connects the turbine and compressor passes through the
center hole. Combustors are made from materials that can withstand the high temperatures of combustion. The liner is often
perforated to enhance mixing of the fuel and air.
There are three main types of combustors, and all three designs are found in gas turbines:
• The combustor at the right is an annular combustor with the liner sitting inside the outer casing which has been peeled
open in the drawing. Many modern combustors have an annular design.
• The combustor on the left is an older can or tubular design. Each can has both a liner and a casing, and the cans are
arranged around the central shaft.
• A compromise design (not shown) is a can-annular design, in which the casing is annular and the liner is can-shaped. The
advantage to the can-annular design is that the individual cans are more easily designed, tested, and serviced.
Turbine blades exist in a much more hostile environment than compressor blades. Located just downstream of the
combustor, turbine blades experience flow temperatures of more than a thousand degrees Fahrenheit. Turbine blades must
be made of special materials that can withstand the heat, or they must be actively cooled. In active cooling, the nozzles and
blades are hollow and cooled by air which is bled off the compressor. The cooling air flows through the blade and out
through the small holes on the surface to keep the surface cool
All turbine engines have a combustor, in which the fuel is combined with high pressure air and burned. The resulting high
temperature exhaust gas is used to turn the turbine and produce thrust when passed through a nozzle.
The combustor is located between the compressor and the turbine. The combustor is arranged like an annulus, or a
doughnut, as shown by illustrations above. The central shaft that connects the turbine and compressor passes through the
center hole. Combustors are made from materials that can withstand the high temperatures of combustion. The liner is often
perforated to enhance mixing of the fuel and air.
There are three main types of combustors, and all three designs are found in gas turbines:
• The combustor at the right is an annular combustor with the liner sitting inside the outer casing which has been peeled
open in the drawing. Many modern combustors have an annular design.
• The combustor on the left is an older can or tubular design. Each can has both a liner and a casing, and the cans are
arranged around the central shaft.
• A compromise design (not shown) is a can-annular design, in which the casing is annular and the liner is can-shaped. The
advantage to the can-annular design is that the individual cans are more easily designed, tested, and serviced.
Turbine blades exist in a much more hostile environment than compressor blades. Located just downstream of the
combustor, turbine blades experience flow temperatures of more than a thousand degrees Fahrenheit. Turbine blades must
be made of special materials that can withstand the heat, or they must be actively cooled. In active cooling, the nozzles and
blades are hollow and cooled by air which is bled off the compressor. The cooling air flows through the blade and out
through the small holes on the surface to keep the surface cool
The flame stabilizing and general-flow patterns are illustrated above for a typical “can-type” combustion chamber.
Although modern engines use one continuous annular combustion chamber, the can-type simplifies illustration of the
cooling and combustion techniques used in all combustion chambers.
The temperature of the flame illustrated in the center of the combustor is approximately 3200°F at its tip when the engine is
operating at full load. Metals used in combustion chamber construction are not capable of withstanding temperatures in
this range; therefore, the design provides airflow passages between the inner and the outer walls of the chamber for
cooling and flame shaping.
Air flowing into the inner chamber is directed through small holes to shape the flame centering it within the chamber, to
prevent its contact with the chamber walls. Approximately 82% of the airflow into combustion chambers is used for cooling
and flame shaping; only 18% is used for fuel combustion. Regulation of fuel flow determines engine speed. Stator vane
control in the compressor controls pressure and velocity into the combustion chamber as a function of compressor speed.
Although modern engines use one continuous annular combustion chamber, the can-type simplifies illustration of the
cooling and combustion techniques used in all combustion chambers.
The temperature of the flame illustrated in the center of the combustor is approximately 3200°F at its tip when the engine is
operating at full load. Metals used in combustion chamber construction are not capable of withstanding temperatures in
this range; therefore, the design provides airflow passages between the inner and the outer walls of the chamber for
cooling and flame shaping.
Air flowing into the inner chamber is directed through small holes to shape the flame centering it within the chamber, to
prevent its contact with the chamber walls. Approximately 82% of the airflow into combustion chambers is used for cooling
and flame shaping; only 18% is used for fuel combustion. Regulation of fuel flow determines engine speed. Stator vane
control in the compressor controls pressure and velocity into the combustion chamber as a function of compressor speed.
TURBINE
All gas turbine engines have a turbine located downstream of the combustor to extract energy from the hot flow and turn the
compressor. Work is done on the turbine by the hot exhaust flow from the combustor.
Since the turbine extracts energy from the flow, the pressure decreases across the turbine. The pressure gradient helps
keep the boundary layer flow attached to the surface of the turbine blades. Since the boundary layer is less likely to
separate on a turbine blade than on a compressor blade, the pressure drop across a single turbine stage can be much
greater than the pressure increase across a corresponding compressor stage. A single turbine stage can be used to drive
multiple compressor stages. Because of the high pressure change across the turbine, the flow tends to leak around the tips
of the blades. The tips of turbine blades are often connected by a thin metal band to keep the flow from leaking.
Turbine blades exist in a much more hostile environment than compressor blades. Sitting just downstream of the
combustor, the blades experience flow temperatures of more than a thousand degrees Fahrenheit. Turbine blades must be
made of special materials that can withstand the heat, or they must be actively cooled. In active cooling, the nozzles and
blades are hollow and cooled by air which is bled off the compressor. The cooling air flows through the blade and out
through the small holes on the surface to keep the surface cool.
All gas turbine engines have a turbine located downstream of the combustor to extract energy from the hot flow and turn the
compressor. Work is done on the turbine by the hot exhaust flow from the combustor.
Since the turbine extracts energy from the flow, the pressure decreases across the turbine. The pressure gradient helps
keep the boundary layer flow attached to the surface of the turbine blades. Since the boundary layer is less likely to
separate on a turbine blade than on a compressor blade, the pressure drop across a single turbine stage can be much
greater than the pressure increase across a corresponding compressor stage. A single turbine stage can be used to drive
multiple compressor stages. Because of the high pressure change across the turbine, the flow tends to leak around the tips
of the blades. The tips of turbine blades are often connected by a thin metal band to keep the flow from leaking.
Turbine blades exist in a much more hostile environment than compressor blades. Sitting just downstream of the
combustor, the blades experience flow temperatures of more than a thousand degrees Fahrenheit. Turbine blades must be
made of special materials that can withstand the heat, or they must be actively cooled. In active cooling, the nozzles and
blades are hollow and cooled by air which is bled off the compressor. The cooling air flows through the blade and out
through the small holes on the surface to keep the surface cool.
TURBINE (Continued)
The compressor drive turbine is an “impulse reaction”-type designed for maximum efficiency in converting hot-gas flow into
rotational mechanical energy. A first-stage fixed nozzle directs flow into the first-stage of rotating blades. The impulse of
expanding hot gas upon the lower surface of each rotating blade propels motion in the upward direction.
Hot gas flow above the following blade creates a lower pressure above the blade as above an aircraft wing, causing
additional rotational force. Subsequent stages operate identically, multiplying the rotational force. Compressor and loaddriving
turbines consist of a varying number of stages, depending upon the load being driven and other design
considerations.
The compressor drive turbine is an “impulse reaction”-type designed for maximum efficiency in converting hot-gas flow into
rotational mechanical energy. A first-stage fixed nozzle directs flow into the first-stage of rotating blades. The impulse of
expanding hot gas upon the lower surface of each rotating blade propels motion in the upward direction.
Hot gas flow above the following blade creates a lower pressure above the blade as above an aircraft wing, causing
additional rotational force. Subsequent stages operate identically, multiplying the rotational force. Compressor and loaddriving
turbines consist of a varying number of stages, depending upon the load being driven and other design
considerations.
TURBINE SHAFTS
The figure above shows the standard gas turbine shaft arrangements. Single shaft illustration is the traditional single shaft
assembly. It consists of the axial flow compressor; Turbine and Power Turbine are all mechanically linked. If we add to this
shaft the generator and gearbox, we have a shaft system with a high moment of inertia. This is the favored configuration for
electrical generation because this provides additional speed (Frequency) stability of the electrical current during large load
fluctuations. This configuration is typical of heavy-duty industrial “frame” turbines, such as the MS7001.
The twin shaft illustration shows the standard two shaft arrangement with the compressor and turbine only connected, and an
unconnected power turbine and output shaft that will rotate independently. This configuration is favored for variable speeddrive
packages, such as pumps and compressors, because the gas generator or gas producer can run at its own optimum
speed for a given load. The LM2500 utilizes this configuration and has been applied to both electric power generation and a
variety of mechanical drive applications.
Aircraft jet engines have for many years been adapted for industrial use as shown in the diagrams above. The concentric
shaft illustration, above left, shows a more complicated aero-derivative industrial turbine arrangement. This, too, is still
essentially a two shaft configuration but the gas generator core (an original jet-engine) was designed with two spools, a Low
Pressure Shaft and a High Pressure Shaft. This engine configuration allows the load to be driven from either the exhaust end
or the compressor air intake end. This is the configuration used by the LM6000
The concentric shaft with power turbine illustration is essentially a two shaft arrangement with a gas generator originally
designed for propulsion. An independently rotating Power Turbine, manufactured especially to match the flow of the jet
engine, is added to the gas path as the power/torque producer. This configuration is found in the LM1600 and the LMS100.
The figure above shows the standard gas turbine shaft arrangements. Single shaft illustration is the traditional single shaft
assembly. It consists of the axial flow compressor; Turbine and Power Turbine are all mechanically linked. If we add to this
shaft the generator and gearbox, we have a shaft system with a high moment of inertia. This is the favored configuration for
electrical generation because this provides additional speed (Frequency) stability of the electrical current during large load
fluctuations. This configuration is typical of heavy-duty industrial “frame” turbines, such as the MS7001.
The twin shaft illustration shows the standard two shaft arrangement with the compressor and turbine only connected, and an
unconnected power turbine and output shaft that will rotate independently. This configuration is favored for variable speeddrive
packages, such as pumps and compressors, because the gas generator or gas producer can run at its own optimum
speed for a given load. The LM2500 utilizes this configuration and has been applied to both electric power generation and a
variety of mechanical drive applications.
Aircraft jet engines have for many years been adapted for industrial use as shown in the diagrams above. The concentric
shaft illustration, above left, shows a more complicated aero-derivative industrial turbine arrangement. This, too, is still
essentially a two shaft configuration but the gas generator core (an original jet-engine) was designed with two spools, a Low
Pressure Shaft and a High Pressure Shaft. This engine configuration allows the load to be driven from either the exhaust end
or the compressor air intake end. This is the configuration used by the LM6000
The concentric shaft with power turbine illustration is essentially a two shaft arrangement with a gas generator originally
designed for propulsion. An independently rotating Power Turbine, manufactured especially to match the flow of the jet
engine, is added to the gas path as the power/torque producer. This configuration is found in the LM1600 and the LMS100.
NOx CONTROL
Oxides of Nitrogen result from the thermal fixation of molecular nitrogen and oxygen in the combustion air. Its rate of
formation is extremely sensitive to local flame temperature and, to a lesser extent, to local oxygen concentrations. Virtually
all thermal NOx is formed in the region of the flame at the highest temperature. Maximum thermal NOx production occurs
at a slightly lean fuel-to-air ratio due to the excess availability of oxygen for reaction within the hot flame zone. Control of
local flame fuel-to-air ratio is critical in achieving reductions in thermal NOx.
Combustion Controls
Reduction of Nox emissions are accomplished by:
• Injection of water or steam at the fuel nozzle in order to reduce combustion temperature
• Specially designed Dry Low Emissions (DLE) combustors and fuel systems
The injection of water or steam into the flame area of a turbine combustor provides a heat sink, which lowers the flame
temperature and thereby reduces thermal NOx formation. Water or steam injection, also referred to as "wet controls," have
been applied effectively to both aeroderivative and heavy duty gas turbines, and to all configurations. Reduction
efficiencies of 70 to 85+ percent can be achieved with properly controlled water or steam injection, with NOx emissions
generally higher for oil-fired turbines than for natural gas-fired units. The most important factor affecting reduction efficiency
is the water-to-fuel ratio. In general, NOx reduction increases as the water-to-fuel ratio increases; however, increasing the
ratio increases carbon monoxide and, to a lesser extent, hydrocarbon emissions at water-to-fuel ratios less than one.
Further, energy efficiency of the turbine decreases with increasing water-to-fuel ratio.
Post-Combustion Controls
The major type of post-combustion control used in gas turbines is Selective Catalytic Reduction (SCR). Applications use
SCR to supplement reductions from steam or water injection, or combustion modifications. Carefully designed SCR
systems can achieve NOx reduction efficiencies as high as 90 percent. The Selective Catalytic Reduction (SCR) process
reduces NOx emissions by using ammonia in the presence of a catalyst. Vaporized ammonia is injected into the flue gas
at the appropriate temperature. The ammonia functions, in the presence of the NOx removal catalyst, as a reducing agent
to decompose nitrous oxides NOx in the flue gas into nitrogen gas and water vapor.
Oxides of Nitrogen result from the thermal fixation of molecular nitrogen and oxygen in the combustion air. Its rate of
formation is extremely sensitive to local flame temperature and, to a lesser extent, to local oxygen concentrations. Virtually
all thermal NOx is formed in the region of the flame at the highest temperature. Maximum thermal NOx production occurs
at a slightly lean fuel-to-air ratio due to the excess availability of oxygen for reaction within the hot flame zone. Control of
local flame fuel-to-air ratio is critical in achieving reductions in thermal NOx.
Combustion Controls
Reduction of Nox emissions are accomplished by:
• Injection of water or steam at the fuel nozzle in order to reduce combustion temperature
• Specially designed Dry Low Emissions (DLE) combustors and fuel systems
The injection of water or steam into the flame area of a turbine combustor provides a heat sink, which lowers the flame
temperature and thereby reduces thermal NOx formation. Water or steam injection, also referred to as "wet controls," have
been applied effectively to both aeroderivative and heavy duty gas turbines, and to all configurations. Reduction
efficiencies of 70 to 85+ percent can be achieved with properly controlled water or steam injection, with NOx emissions
generally higher for oil-fired turbines than for natural gas-fired units. The most important factor affecting reduction efficiency
is the water-to-fuel ratio. In general, NOx reduction increases as the water-to-fuel ratio increases; however, increasing the
ratio increases carbon monoxide and, to a lesser extent, hydrocarbon emissions at water-to-fuel ratios less than one.
Further, energy efficiency of the turbine decreases with increasing water-to-fuel ratio.
Post-Combustion Controls
The major type of post-combustion control used in gas turbines is Selective Catalytic Reduction (SCR). Applications use
SCR to supplement reductions from steam or water injection, or combustion modifications. Carefully designed SCR
systems can achieve NOx reduction efficiencies as high as 90 percent. The Selective Catalytic Reduction (SCR) process
reduces NOx emissions by using ammonia in the presence of a catalyst. Vaporized ammonia is injected into the flue gas
at the appropriate temperature. The ammonia functions, in the presence of the NOx removal catalyst, as a reducing agent
to decompose nitrous oxides NOx in the flue gas into nitrogen gas and water vapor.
