GAS TURBINE THEORY
This article reviews the basic gas turbine cycle and major turbine components and their functions.
The earliest example of jet propulsion can be traced as far back as 150 B.C.to an Egyptian namedHero. Hero invented a toy that rotated on top of a boiling pot due to the reaction effect of hot air or steam exiting several nozzles arranged radially around a wheel. He called his invention anaeolipile (Figure 1).
Figure 1: Heros Aeolipile from 150 BC
In 1232, the Chinese used rockets to frighten enemy soldiers. Around 1500, Leonardo da Vinci drew a sketch of a device that rotated due to the effect of hot gasses flowing up a chimney. The device was intended to be used to rotate meat being roasted. In 1629, another Italian named Giovanni Branca actually developed a device that used jets of steam to rotate a turbine that in turn was used to operate machinery. This was the first practical application of a steam turbine.
The first patent for a turbine engine was granted in 1791 to an Englishman named John Barber. It incorporated many of the same elements of a modern gas turbine but used a reciprocating compressor. There are many more early examples of turbine engines designed by various inventors, but none were considered true gas turbines because they incorporated steam at some point in the process.
In 1872, a man by the name of Stolze designed the first true gas turbine. His engine incorporated a multistage turbine section and a multistage axial flow compressor. He tested working models in the early 1900s. Charles Curtis, the inventor of the Curtis steam engine, filed the first patent application in the US for a gas turbine engine. His patent was granted in 1914.
The General Electric Company started their gas turbine division in 1903. An engineer named Stanford Moss lead most of the projects. His most outstanding development was the General Electric turbo-supercharger during World War I. It used hot exhaust gasses from a reciprocating engine to drive a turbine wheel that, in turn, drove a centrifugal compressor used for supercharging. The evolutionary process of turbo-supercharger design and construction made it possible to construct the first reliable gas turbine engines. General Electric placed its first gas turbine into commercial operation in 1949.
Useful work or propulsive thrust can be obtained from a gas turbine engine. It may drive a generator, pump, or propeller or, in the case of a pure jet aircraft engine, develop thrust by accelerating the turbine exhaust flow through a nozzle.
Large amounts of power can be produced by such an engine that, for the same output, is much smaller and lighter than a reciprocating internal combustion engine. Reciprocating engines depend on the up and-down motion of a piston, which must then be converted to rotary motion by a crankshaft arrangement, whereas a gas turbine delivers rotary shaft power directly.
Although conceptually the gas turbine engine is a simple device, the components for an efficient unit must be carefully designed and manufactured from costly materials because of the high temperatures and stresses encountered during operation. Thus, gas turbine engine installations are usually limited to large units where they become cost effective. Figure 2 shows a simple gas turbine arrangement.
Figure 2: Simple Gas Turbine
A gas turbine operates by doing the following:
1. Continuously drawing in fresh air.
2. Compressing the air to a higher pressure.
3. Adding and burning fuel in the compressed air to increase its energy.
4. Directing the high-pressure, high-temperature air to an expansion turbine that converts the gas energy to mechanical energy of a rotating shaft. The resulting low-pressure, lower temperature gases are discharged to atmosphere.
Most of the turbine output is required to operate the compressor; only the remainder is available to supply shaft work to a generator, pump, or other device.
Figure 3 compares a simple gas turbine cycle to that of a four-cycle reciprocating engine. A major difference is that the gas turbine produces continuous power, while the reciprocating engine cycle is intermittent.
Figure 3: Gas Turbine to Reciprocating Engine Comparison
In a four-cycle engine, a loss occurs due to the pressure drop involved in the exhaust stroke. This loss is avoided by creating a cycle in which the exhaust stroke is longer than the compression stroke, thus allowing the working fluid to be expanded to atmospheric pressure. Such a cycle has been devised and is called the Brayton cycle (Figure 4). It is also called the constant pressure cycle since combustion and exhaust both take place at a constant pressure. When the Brayton cycle is worked out for a steady-flow process, the result is a simple gas turbine cycle.
Figure 4: Brayton Cycle
In a simple gas turbine cycle, combustion and exhaust occur at constant pressure and compression and expansion occur continuously, rather than intermittently as previously stated. This means that gas turbine power is continuously available, whereas in a reciprocating engine, power takeoff is available only on the exhaust stroke.
The points on Figure 3 and Figure 4 are consistent. At point 1, air enters the compressor. The high-pressure discharge air at point 2 is mixed with fuel in the combustor. The product of this continuous combustion at point 3 enters the turbine and is expanded to atmospheric pressure (point 4). The turbine provides the horsepower by converting the kinetic energy of the gases to mechanical energy of the rotating shaft. This provides the horsepower to drive the compressor and the load, usually a generator.
An idealized gas turbine engine operating without any losses on this simple Brayton cycle is considered first. If, for example, air enters the compressor at 59F and atmospheric pressure is compressed to 15 times atmospheric pressure, it then absorbs heat from the fuel at a constant pressure until the temperature reaches 2,012F prior to expansion through the turbine back to atmospheric pressure. This idealized unit would require a turbine output of 1.68 kilowatts for each kilowatt of useful power with 0.68 kilowatt absorbed to drive the compressor. The thermal efficiency of the unit (net work produced divided by energy added through the fuel) would be 48%.
If, for a unit operating between the same pressure and temperature limits, the compressor and the turbine are only 80% efficient (i.e., the work of an ideal compressor equals 0.8 times the actual work, while the actual turbine output is 0.8 times the ideal output), the situation changes drastically even if all other components remain ideal. For every kilowatt of net power produced, the turbine must now produce 2.71 kilowatts while the compressor work becomes 1.71 kilowatts.
The thermal efficiency drops to 25.9%. This illustrates the importance of highly efficient compressors and turbines. Historically, it was the difficulty of designing efficient compressors, even more than efficient turbines, that delayed the development of the gas turbine engine. Modern units can have compressor efficiencies of 86-88% and turbine efficiencies of 88-90 percent at design conditions.
Raising the turbine-inlet temperature can increase efficiency and power output. All materials lose strength at very high temperatures, however, and since turbine blades travel at high speeds and are subject to severe centrifugal stresses, turbine inlet temperatures above 1,100C require special blade cooling. It can be shown that for every maximum turbine inlet temperature there is also an optimum pressure ratio. Modern aircraft gas turbines with blade cooling operate at turbine inlet temperatures above 1,370C and at pressure ratios of about 30:1.
The gas path is the path by which gases flow through the gas turbine from the air inlet through the compressor, combustion section, and turbine, to the turbine exhaust. When the turbine starting system is actuated and the clutch is engaged, ambient air is drawn through the air inlet plenum assembly, filtered and compressed in a multistage, axial flow compressor. For pulsation protection during startup, compressor bleed valves are open and the variable inlet guide vanes are closed. When the high speed relay actuates, the bleed valves begin operation automatically and the variable inlet guide vane actuator energizes to position the inlet guide vanes for normal turbine operation. Compressed air from the compressor flows into the annular space surrounding the combustion chambers, it then flows into the spaces between the outer combustion casings and the combustion liners, and enters the combustion zone through metering holes in each of the combustion liners.
Fuel from an off base source is provided to flow lines, each terminating at the primary and secondary fuel nozzles in the end cover of the separate combustion chambers.
On liquid-fueled machines, the fuel is controlled prior to being distributed to the nozzles to provide an equal flow into each liquid fuel distributor valve mounted on each end cover and each liquid fuel line on each secondary nozzle assembly.
On gas-fueled machines, the fuel nozzles are the metering orifices that provide the proper flow into the combustion zones in the chambers. The nozzles introduce the fuel into the combustion zone within each chamber where it mixes with the combustion air and is ignited by one or more of the spark plugs. At the instant fuel is ignited in one combustion chamber, flame is propagated through connecting crossfire tubes, to all other combustion chambers where it is detected by four primary flame detectors, each mounted on a flange on the combustion casings.
The hot gases from the combustion chambers flow into separate transition pieces attached to the aft end of the combustion chamber liners and flow from there to the three stage turbine section. Each stage consists of a row of fixed nozzles and a row of turbine buckets. In each nozzle row, the kinetic energy of the jet is increased, with an associated pressure drop, which is absorbed as useful work by the turbine rotor buckets, resulting in shaft rotation used to turn the generator rotor to generate electrical power. After passing through the third stage buckets, the gases are directed into the exhaust diffuser. The gases then pass into the exhaust plenum and are introduced to atmosphere through the exhaust stack.
This section briefly describes the following major components of the GE turbine:
The air inlet system is specifically designed to modify the quality of air under various temperature, humidity, and contamination situations and make it more suitable for use.
The gas turbine compressor is typically an axial flow design that efficiently compresses a large volume of air. The compressor can consist of up to 18 individual stages operating in series.
The combustion system consists of liners into which fuel is added and burned with a portion of the compressed air. The excess compressed air is used to cool the products of combustion.
Fuel is injected into each liner by fuel nozzles that atomize the fuel for good burning.
The fuel is ignited initially by electric igniters. Once the fire is started, the combustion process is self-sustaining as long as fuel and air are available.
The turbine consists of typically three to four stages. Each stage consists of a stationary row of nozzles where the high-energy gases are increased in velocity and directed toward a rotating row of buckets (airfoils) attached to the turbine shaft. The high velocity gases push against the buckets, converting the gases kinetic energy into shaft power.
Varying the amount of fuel injected into the combustion liners changes the energy from the combustion system, available to drive the turbine.
The exhaust system is an internally insulated diffuser duct that directs the gas turbine exhaust flow from the power turbine exit to the HRSG or into a bypass stack, for a simple cycle turbine.
The exhaust system gradually diffuses the exhaust flow for maximum pressure recovery of the exhaust flow thereby enhancing turbine performance.
Support systems consisting of lube oil, cooling water, ignition and fuel system, hydraulic oil, fire protection, and wash water are covered in more detail later.
The base that supports the gas turbine is a structural steel fabrication of welded steel beams and plate. Its prime function is to provide a support upon which to mount the gas turbine.
Lifting trunnions and supports are provided, two on each side of the base in line with the two structural cross members of the base frame. Machined pads on each side on the bottom of the base facilitate its mounting to the site foundation. Two machined pads atop the base frame are provided for mounting the aft turbine supports.
The typical GT has rigid leg type supports at the compressor end and supports with top and bottom pivots at the turbine end.
Figure 5 shows the major components discussed above. This figure is for a GE 7FA gas turbine but is typical of most gas turbines.
Figure 5: Typical Gas Turbine Major Sections