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Chapter 2

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  • 1. GAS TURBINES Chapter BASIC 2 1. Introduction 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
  • 2. 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 some instances. 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.
  • 3. 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 plant. 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 service. 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 machines. 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
  • 4. 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.
  • 5. Figure 1. Basic Gas Turbine Figure 2. Brayton Cycle Efficiency VS Pressure Ratio
  • 6. Figure 3. Two Shaft Arrangement Figure 4. Combined Cycle
  • 7. Figure 5. Gas Turbine Designations
  • 8. Figure 6
  • 9. Figure 7 Figure 8
  • 10. 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 construction. 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. Figure 9
  • 11. Figure 10 Figure 11 2.2 Accessory Section
  • 12. 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 configuration. 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 pressure increases. 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.
  • 13. 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 compressor. Figure 12 A single-sided impeller does benefit from ram intake and its less turbulent air entry makes it well suited for aircraft installation. Figure 13
  • 14. 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. Figure 14
  • 15. 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 weight  Low starting power requirements. Figure 15 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 turbofan engines. 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
  • 16. intermediate compressor and the innermost compressor is the high pressure, or N3 compressor. Figure 16 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.
  • 17. Figure 17 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.
  • 18.  Low pressure rise per stage, approximately 1.27:1.  Compression ratio. 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 computed as: 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
  • 19. 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. Figure 18
  • 20. Figure 19 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. Figure 20
  • 21. 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 helicopters. 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.
  • 22. Figure 21 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.
  • 23. Figure 22 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. Figure 23 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.
  • 24. Figure 24 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 effect. Fig. 24. shows the high-pressure area of the first stage blade being pulled into
  • 25. 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 excessive. Figure 25 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
  • 26. 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 vector. 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 point.
  • 27. Figure 26 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 operation. 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.
  • 28. 7.2 Comparison of axial flow with centrifugal flow compressors. 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 pressure increases. 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 27 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.
  • 29. 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.
  • 30. 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
  • 31. Figure 30 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
  • 32. 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
  • 33. 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 combustor designs. 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 for starting 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
  • 34. 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 combustor.
  • 35. 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.
  • 36. Figure 35 shows an impulse turbine is driven by the impulse of hot gases on the blade. Figure 36 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
  • 37. 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
  • 38. combination of the impulse pressure and the reaction force of the flowing gases. 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, and filter. 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
  • 39. 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. 11.4. Governor/Control 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
  • 40. 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 reduce efficiency. 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.