Axial compressor

This article is licensed under theGNU Free Documentation License.It uses material from the Wikipedia article "Axial compressor" (click for full Wikipedia text)

 'Axial compressors ' are Gas compressor in which the fluid flows mainly parallel to the rotation axis. Axial flow compressors have large mass flow capacity and high efficiencies, but have a smaller pressure rise per stage than centrifugal compressors. Axial compressors are widely used in gas turbines, notably jet engines. Engines using an axial compressor are known as  'axial-flow '. Almost all modern engines are axial-flow, the notable exception being those used in helicopters, where the smaller size of the centrifugal compressor is useful.

Description

Axial compressors are essentially a steam turbine reversed; instead of high-pressure gas flowing into the turbine and forcing it to rotate to provide power, in the compressor role power is provided from an external source in order to spin the system and compress the gas. A typical axial compressor has a  rotor  which looks like a fan with contoured blades followed by a stationary set of blades, called a  stator . As the diagram illustrates, compressor blades/vanes are relatively flat in section. Turbine blades/vanes, on the other hand, have significant curvature. Each pair of rotors and stators is referred to as a  stage , and most axial compressors have a number of such stages placed in a row along a common power shaft in the center. The stator blades are required in order to ensure reasonable efficiency; without them the gas would rotate with the rotor blades resulting in a large drop in efficiency. Improvements can be made by replacing the stators with a second set of fans rotating in the opposite direction, but these designs have generally proven to be too complex to be worthwhile. Each stage is smaller than the last, as the volume of air is reduced by the compression of the preceding stage. Axial compressors therefore generally have a conical shape, widest at the inlet. Compressors typically have between 9 and 15 stages. In a jet engine the compressor is powered by a turbine placed in the hot exhaust, using up some of its energy. In such a system axial compressors typically use between 60% and 65% of the engine's power in order to run. This explains why jet engines are not used in cars; even standing still at a red light the engine would be running almost full out just to idle. In aircraft this is not an issue, as the engine is running almost full out for the entire trip.

Development

Early axial compressors offered poor efficiency, so poor that in the early 1920s a number of papers claimed that a practical jet engine would be impossible to construct. Things changed dramatically after Alan Arnold Griffith published a seminal paper in 1926, noting that the reason for the poor performance was that existing compressors used flat blades and were essentially "flying stalled". He showed that the use of airfoils instead of the flat blades would dramatically increase efficiency, to the point where a practical jet engine was a real possibility. He concluded the paper with a basic diagram of such an engine, which included a second turbine that was used to power a propeller. Although Griffith was well known due to his earlier work on metal fatigue and stress (physics) measurement, little work appears to have started as a direct result of his paper. The only obvious effort was a test-bed compressor built by Griffith's colleague at the RAE, Haine Constant. Other early jet efforts, notably those of Frank Whittle and Hans von Ohain, were based on the much better understood centrifugal compressor which was widely used in superchargers. Griffith had seen Whittle's work in 1929 and pooh-poohed it, noting an error in the math and going on to claim that the frontal size of the engine would make it useless on a high-speed aircraft. Real work on axial-flow engines started in the late 1930s, in several efforts that all started at about the same time. In England, Haine Constant reached an agreement with the steam turbine company Metropolitan Vickers (Metrovick) in 1937, starting their turboprop effort based on the Griffith design in 1938. In 1940, after the successful run of Whittle's centrifugal-flow design, their effort was re-designed as a pure jet, the Metrovick F.2. In Germany, von Ohain had produced several working centrifugal engines, some of which had flown, but all short-term development efforts had moved on to Junkers and BMW, who used axial-flow designs. In the United States, both Lockheed and General Electric were awarded contracts in 1941 to develop axial-flow engines, the former a pure jet, the later a turboprop. Northrop also started their own project to develop a turboprop, which the US Navy eventually contracted in 1943. Westinghouse Electric Corporation also entered the race in 1942, their project proving to be the only successful one of the US efforts, later becoming the Westinghouse J30. By the 1950s every major engine development had moved on to the axial-flow type. As Griffith had originally noted in 1929, the large frontal size of the centrifugal compressor caused it to have higher drag than the narrower axial-flow type. Additionally the axial-flow design could improve its compression ratio simply by adding additional stages and making the engine slightly longer. In the centrifugal-flow design the compressor itself had to be larger in diameter, which was much more difficult to "fit" properly on the aircraft. On the other hand, centrifugal-flow designs remained much less complex (the major reason they "won" in the race to flying examples) and therefore have a role in places where size and streamlining are not so important. For this reason they remain a major solution for helicopter engines, where the compressor lies flat and can be built to any needed size without upsetting the streamlining to any great degree.

Axial-flow jet engines

In the jet engine application, the compressor faces a wide variety of operating conditions. On the ground at takeoff the inlet pressure is high, inlet speed zero, and the compressor spun at a variety of speeds as the power is applied. Once in flight the inlet pressure drops, but the inlet speed increases (due to the forward motion of the aircraft) to recover some of this pressure, and the compressor tends to run at a single speed for long periods of time. There is simply no "perfect" compressor for this wide range of operating conditions. Fixed geometry compressors, like those used on early jet engines, are limited to a design pressure ratio of about 4 or 5:1. As with any heat engine, fuel efficiency is strongly related to the compression ratio, so there is very strong financial need to improve the compressor stages beyond these sorts of ratios. Additionally the compressor may stall if the inlet conditions change abruptly, a common problem on early engines. In some cases, if the stall occurs near the front of the engine, all of the stages from that point on will stop compressing the air. In this situation the energy required to run the compressor drops suddenly, and the remaining hot air in the rear of the engine allows the turbine to speed up whole engine dramatically. This condition, known as  'surging ', was a major problem on early engines and often led to the turbine or compressor breaking and shedding blades. For all of these reasons, axial compressors on modern jet engines are considerably more complex than those on earlier designs.

Spools

All compressors have a sweet spot relating rotational speed and pressure, with higher compressions requiring higher speeds. Early engines were designed for simplicity, and used a single large compressor spinning at a single speed. Later designs added a second turbine and divided the compressor into "low pressure" and "high pressure" sections, the latter spinning faster. This  'two-spool ' design resulted in increased efficiency. Even more can be squeezed out by adding a third spool, but in practice this has proven to be too complex to make it generally worthwhile. That said, there are several three-spool engines in use, perhaps the most famous being the Rolls-Royce RB.211, used on a wide variety of commercial aircraft.

Bleed air, variable stators

As an aircraft changes speed or altitude, the pressure of the air at the inlet to the compressor will vary. In order to "tune" the compressor for these changing conditions, designs starting in the 1950s would "bleed" air out of the middle of the compressor in order to avoid trying to compress too much air in the final stages. This was also used to help start the engine, allowing it to be spun up without compressing much air by bleeding off as much as possible. Bleed systems were already commonly used anyway, to provide airflow into the turbine stage where it was used to cool the turbine blades, as well as provide pressurized air for the air conditioning systems inside the aircraft. A more advanced design, the  'variable stator ', used blades that can be individually rotated around their axis, as opposed to the power axis of the engine. For startup they are rotated to "open", reducing compression, and then are rotated back into the airflow as the external conditions require. The General Electric J79 was the first major example of a variable stator design, and today it is a common feature of most military engines. Closing the variable stators progressively, as compressor speed falls, reduces the slope of the surge (or stall) line on the operating characteristic (or map), improving the surge margin of the installed unit. By incorporating variable stators in the first five stages, General Electric Aircraft Engines has developed a ten-stage axial compressor capable of operating at a 23:1 design pressure ratio.

Bypass

For jet engine applications, the "whole idea" of the engine is to move air to provide thrust. In most cases the engine can actually provide much more energy than it can air; the inlet into the compressor is simply too small to move the amount of air that the engine could, in theory, heat and use. A number of engine designs had experimented with using some of the turbine power to drive a secondary "fan" for added air flow, starting with the Metrovick F.3, which placed a fan at the rear of a late-model F.2 engine. A much more practical solution was created by Rolls-Royce in their early 1950's Rolls-Royce Conway engine, which enlarged the first compressor stage to be larger than the engine itself. This allowed the compressor to blow cold air past the interior of the engine, somewhat similar to a propeller. This technique allows the engine to be designed to produce the amount of energy needed, and any air that cannot be blown through the engine due to its size is simply blown around it. Since that air is not compressed to any large degree, it is being moved without using up much energy from the turbine, allowing a smaller core to provide the same mass flow, and thrust, as a much larger "pure jet" engine. The resultant engine is called a "turbofan." This technique also has the added benefit of mixing the cold bypass air with the hot engine exhaust, greatly lowering the exhaust temperature. Since the sound of a jet engine is strongly related to the exhaust temperature, bypass also dramatically reduces the sound of the engine. Early jetliners from the 1960s were famous for their "screaming" sound, whereas modern engines of greatly higher power generally give off a much less annoying "whoosh" or even buzzing. Mitigating this savings is the fact that Drag (physics) increases exponentially at high speeds, so while the engine is able to operate far more efficiently, this typically translates into a smaller real-world effect. For instance, the latest Boeing 737's with high-bypass CFM56 engines operates at an overall efficiency about 30% better than the earlier models. Military turbofans, on the other hand, especially those used on combat aircraft, tend to have so low a bypass ratio that they are sometimes referred to as "leaky turbojets."

Turbine cooling

The limiting factor in jet engine design is not the compressor, but the temperature at the turbine. It is fairly easy to build an engine that can provide enough compressed air that when burnt will melt the turbine; this was a major cause of failure in early German engines. Improvements in air cooling and materials have dramatically improved the temperature performance of turbines, allowing the compression ratio of jet engines to increase dramatically. Early test engines offered perhaps 3:1 and production engines like the Jumo 004 were about 4:1, about the same as contemporary piston engines. Improvements started immediately and have not stopped; the latest Rolls-Royce Trent operates at about 40:1, far in excess of any piston engine. Since compression ratio is strongly related to fuel economy, this eightfold increase in compression ratio really does result in an eightfold increase in fuel economy for any given amount of power, which is the reason there is strong pressure in the airline industry to use only the latest designs.

Design notes

Energy Exchange between rotor and fluid

The relative motion of the blades relative to the fluid adds velocity or pressure or both to the fluid as it passes through the rotor. The fluid velocity is increased through the rotor, and the stator converts kinetic energy to pressure energy. Some diffusion also occurs in the rotor in most practical designs. The increase in velocity of the fluid is primarily in the tangential direction (swirl) and the stator removes this angular momentum. The pressure rise results in a stagnation temperature rise. For a given geometry the temperature rise depends on the square of the tangential Mach number of the rotor row. Current turbofan engines have fans that operate at Mach 1.7 or more, and require significant containment and noise suppression structures to reduce blade loss damage and noise.

Velocity diagrams

The blade rows are designed at the first level using velocity diagrams. The velocity diagram shows the relative velocities of the blade rows and the fluid. The axial flow through the compressor is kept as close as possible to Mach 1 to maximize the thrust for a given compressor size. The tangential Mach number determines the attainable pressure rise. The blade rows turn the flow through and angle ß and larger turning allows a higher temperature ratio, but requires higher solidity. Modern blades rows have lower aspect ratios and higher solidity.

Compressor maps

A map shows the performance of a compressor and allows determination of optimal operating conditions. It shows the mass flow along the horizontal axis, typically as a percentage of the design mass flow rate, or in actual units. The pressure rise is indicated on the vertical axis as a ratio between inlet and exit stagnation pressures. A surge or stall line identifies the boundary to the left of which the compressor performance rapidly degrades and identifies the maximum pressure ratio that can be achieved for a given mass flow. Contours of efficiency are drawn as well as performance lines for operation at particular rotational speeds.

Compression stability

Operating efficiency is highest close to the stall line. If the downstream pressure is increased beyond the maximum possible the compressor will stall and become unstable. Typically the instability will be at the Helmholtz frequency of the system, taking the downstream plenum into account.

This article is licensed under the GNU Free Documentation License.It uses material from the Wikipedia article "Axial compressor" (click for full Wikipedia text) Hide