Basic Modules of Gas Turbine High Bypass Engine
Air Inlet Section: As shown in figure 5, the air inlet section is the most forward section of the engine. This section forms the air inlet for the engine and houses the inlet guide vanes. The purpose of the vanes is to direct the incoming air at the proper angle to the first fan stage. The vane at the bottom of the air inlet is thicker than the others in order to accommodate engine tubing. As indicated in figure 6, the no. 1 bearing support assembly is mounted in the center of the compressor inlet case. Behind the front support is the no. 1 bearing rear support. Mounted on the front of the inlet case, in the center is the front accessory drive support .
Air Inlet Case
In the above figure of inlet case, 1 is no. 1 bearing front support, 2 is fan-inlet case, 3 and 4 are boss inlet case, and 5 is connection assembly. As reflected in the drawing, the compressor inlet case comprises of 19 inlet guide vanes. Eighteen of which are identical, but the bottom vane is thicker to accommodate engine tubing. The vanes are brazed between the inner and outer shroud cases which are made of titanium (Walsh and Fletcher). The outer shroud has a double wall and the air passes through the hollow vanes to the center of the assembly, where it discharges forward through the front of the inner shroud case. A pressure oil passage in the support carried oil from the rear of the outer flange into the center, then rearward through the no. 1 bearing oil nozzle. Oil scavenge passage carries oil from a pump boss on the lower rear face of the support cavity back toward the outside of the support, then to another opening in the rear of the outer flange.
Compressor Module: the front compressor section and fan section of the engine are usually part of the same rotating assembly. The fan is actually the outer ends of the first two compressor blades. These two blades are enclosed by the front and rear fan case. The low pressure compressor partially compresses the air entering the engine before the air is delivered to high pressure compressor. While primary air flows through the core of the engine, secondary air from the fan is passing through the fan duct surrounding the engine. In this engine, the primary and secondary air is almost equally divided by weight. The first compressor is driven by the second, third, and fourth turbine stages through the inner or front compressor drive shaft. Thus, the first compressor rotates independently of the rear compressor.
The vanes at the first stage reduce the speed of air and direct it to the compressor rotor blades of second stage with the process continuing through all the stages of compressor rotor. From here, air enters the section of diffusion (Mattingly, Heiser and Pratt). Diffusers’ inlet has the total air velocity at the highest and maximum pressure of compressor. Air circulates in the diffuser wherein it meets with higher cross-sectional area, thereby decreasing the velocity of air and increasing the static pressure. Diffuser outlet has the maximum of static pressure. The altitude of the location of gas turbine alters the power developed by the gas turbine. The reason is the decreasing air density with altitude, hence on a dry or humid day; the gas turbine at high altitude will swallow air of less weight. With the increase in the forward speed, increases the ram air pressure along with pressure and air temperature. Ram air pressure is the air pressure of free stream created by the aircraft engine’s forward motion (Hill and Peterson). The impact of power produced by gas turbine on temperature of air intake is demonstrated in figure 7.
Variation in shaft power with air temperature of inlet
There are three types of compressors; axial, centrifugal and combination of axial and centrifugal; axial-centrifugal compressor.
Figure 7: (A) Axial and (B) Centrifugal flow compressor
Compressor of Axial flow: in the parallel direction to the engine’s longitudinal axis, air is compressed. This type of design of compressor comprises of several stages to create highly efficient compression ratios. This compressor is streamlined in shape and is optimum for ram jets or aircrafts with high speed. However, this design is less sturdy in comparison to centrifugal compressor which makes it susceptible to FOD (foreign object damage).
Axial Flow Compressor
Centrifugal flow compressor: with the turn of the rotor, air is sucked into blades next to the front rotor’s center. The air is accelerated by the centrifugal force with the outward movement from axis of rotation to the rotor’s edge. From there, it is forced via the section of diffuser with high kinetic energy. This results in increasing pressure due to reducing speed of air in diffuser and the conversion of velocity energy to pressure energy. This type of compressor is capable of high compression ratio however is not feasible for large engines due to its weight and size in comparison to axial flow compressor (Bloch and Soares).
Axial-Centrifugal flow compressor: Dual compressor or axial centrifugal flow compressor is the combination of two types of compressors. It possesses benefits and characteristics of the compressors. This type of compressor is axial flow compressor five to seven stages and centrifugal flow compressor one stage. The compressors are on the same shaft, therefore turning at the same speed and direction. This compressor is specific application designs like Army helicopters for US.
Axial centrifugal flow compressor
Combustor Module: Three main types of combustion chambers are present, namely can combustor chamber, annular combustor chamber and can-annular combustor chamber. Combustor module comprises of combustion chambers, fuel nozzles and igniter plugs. Mixture of fuel and sir is burned in combustor and the product is transferred to turbine in the engine. Fuel is injected at the burner’s upstream end in the form of atomized spray. Simplex of dual fuel nozzles are variants of fuel nozzles which delivers gaseous or liquid fuel or a combination of both respectively. Combustion air mixes with the fuel by the function of swirler fans. This forms primary air, representing almost 25 % of total ingested air of engine. 1 part of fuel to 15 parts of air is the air-fuel mixture and the residual 75 % of air forms the blanket to the burning gases and lowers the burning temperature. The flow of cooling air reduces the temperature of almost 3000° F to the optimum temperature of inlet guide vanes of turbine. Secondary air is the cooled down air which is directed and controlled by the louvers and holes in the liner of combustion chamber. In high bypass ratio engines, the diameter of fan is much larger than low pressure compressor and the air sucked by the fan is moved through annular sleeve shaped casing which fits round compressor. The bypass air increases the cooling effect and also elevates the performance of gas turbines like development of power, mass flow and ingestion capabilities of FOD’s.
Annular Combustion Chamber: This enhances the compact design and the compressed air is directed to the annular space created through chamber liner around the assembly of turbine. Annular space between the outer wall and chamber directs the secondary air flow. This combustion chamber provides the increased volume of combustion per unit of meta area exposed.
Can Combustion Chamber: This has individual chambers of combustion. The compressor air enters into the chambers through the transition section. Each can have two tubes, cylindrical in shape and concentric the combustion chamber and chamber liner. Combustion carries out in the chamber liner and the holes and louvers controls airflow in area of combustion. Continuous airflow is important to prevent formation of carbon on liner inside. The deposition of carbon reduces the life of burner.
Can-Annular Combustion Chamber: this is a combination of annular and can type of combustion chamber and comprises of annular outer shell and can type mounted on the axis of engine. Combustion chamber is cooled through the air entering liner by louvers and holes.
Nitrogen oxides are produced by the combustor and the amount of their production increases with the increase in the temperature of flame and the corresponding inlet temperature of turbine. Hence, the inlet temperature of turbine should be kept low in the optimum range.
Turbine Module: The gases entering the inlet of turbine have kinetic energy which is transformed into horsepower and is used to run the compressor and support systems. The airfoil design used these days is the axial flow turbine design, however radial flow turbine design is also used. The radial turbine design is less complex, rugged, easy to manufacture and less expensive in comparison to axial flow design of turbine. Radial flow is the altered version of centrifugal compressor which makes them increasingly efficient in the small engines they are used. Alternatively, axial inflow design comprises of stages with each stage comprising of stationary vanes and rotating blades row on the disc. The blades of turbine are reaction of impulsive type and the modern airplanes have both types of sections of blades.
Turbine Blade with Twisted Contour
High pressure turbine blades are cooled with a combination of internal cooling and film cooling. Film cooling involves the injection of compressed air into the blade boundary layer to provide a thin layer of relatively cooler air that protects material surfaces. The automatic turbine rotor clearance control system, also known as the turbine case cooling system, controls and distributes fan air to cool and shrink the HPT and LPT cases. This process increases efficiency by reducing turbine tip clearance for takeoff, climb and cruise operations.
Exhaust Module: The exhaust duct takes the relatively high pressure, low velocity gases leaving the turbine wheel and accelerates this gas flow to sonic or supersonic speeds through the nozzle at its rear. When the turbine operates against a noticeable back pressure, the nozzle converts the remaining pressure energy into a high velocity exhaust. The exhaust duct is constructed of two stainless steel cones.
The outer cone is usually bolted to the turbine casing with the inner cone supported from the outer cone. The vanes used to support the inner core straighten the swirling gas flow. Two types of nozzles are in use these days, convergent type and the convergent-divergent type. The convergent nozzle has a fixed area whereas the convergent-divergent area would be variable.
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J. Mattingly, W. Heiser and D. Pratt. Aircraft Engine Design. AIAA, 2003.
Kerrebrock, Jack L. Aircraft Engines and Gas Turbines. The MIT Press, 1992.
P. Hill and C. Peterson. Mechanics and Thermodynamics of Propulsion. Prentice Hall, 1991.
P. P. Walsh and P. Fletcher. Gas turbine performance. Carlton: John Wiley & Sons, 2004.
Treager. Aircraft: Gas Turbine Engine Technology. McGraw-Hill, 2002.