Axial compressors are rotating, airfoil-based compressors in which the working fluid principally flows parallel to the axis of rotation. This is in contrast with other rotating compressors such as centrifugal, axi-centrifugal and mixed-flow compressors where the air may enter axially but will have a significant radial component on exit.
Axial flow compressors produce a continuous flow of compressed gas, and have the benefits of high efficiencies and large mass flow capacity, particularly in relation to their cross-section. They do, however, require several rows of airfoils to achieve large pressure rises making them complex and expensive relative to other designs (e.g. centrifugal compressor).
Axial compressors are widely used in gas turbines, such as jet engines, high speed ship engines, and small scale power stations. They are also used in industrial applications such as large volume air separation plants, blast furnace air, fluid catalytic cracking air, and propane dehydrogenation. Axial compressors, known as superchargers, have also been used to boost the power of automotive reciprocating engines by compressing the intake air, though these are very rare.
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Axial compressors consist of rotating and stationary components. A shaft drives a central drum, retained by bearings, which has a number of annular airfoil rows attached. These rotate between a similar number of stationary airfoil rows attached to a stationary tubular casing. The rows alternate between the rotating airfoils (rotors) and stationary airfoils (stators), with the rotors imparting energy into the fluid, and the stators converting the increased rotational kinetic energy into static pressure through diffusion. A pair of rotating and stationary airfoils is called a stage. The cross-sectional area between rotor drum and casing is reduced in the flow direction to maintain axial velocity as the fluid is compressed.
A compressor in which the fluid enters and leaves in the axial direction is called axial flow compressor. So, the centrifugal component in the energy equation does not come into play. Here the compression is fully based on diffusing action of the passages. The main parts include a stationary (stator) part and a moving (rotor) part. The diffusing action in stator converts absolute kinetic head of the fluid into rise in pressure. The relative kinetic head in the energy equation is a term that exists only because of the rotation of the rotor. The rotor reduces the relative kinetic head of the fluid and adds it to the absolute kinetic head of the fluid i.e., the impact of the rotor on the fluid particles increases its velocity (absolute) and thereby reduces the relative velocity of the between the fluid and the rotor. In short, the rotor increases the absolute velocity of the fluid and the stator converts this into pressure rise. Designing the rotor passage with a diffusing capability can produce a pressure rise in addition to its normal functioning. This produces greater pressure rise per stage which constitutes a stator and a rotor together. This is the reaction principle in turbomachines. If 50% of the pressure rise in a stage is obtained at the rotor section, it is said to have a 50% reaction.
The increase in pressure produced by a single stage is limited by the relative velocity between the rotor and the fluid, and the turning and diffusion capabilities of the airfoils. A typical stage in a commercial compressor will produce a pressure increase of between 15% and 60% (pressure ratios of 1.15-1.6) at design conditions with a polytropic efficiency in the region of 90-95%. To achieve different pressure ratios, axial compressors are designed with different numbers of stages and rotational speeds. As a general rule-of-thumb we can assume that each stage in a given compressor has the same temperature rise (deltaT). Therefore, as the entry temperature (Tstage entry) to each stage must increase progressively through the compressor, the ratio deltaT/Tstage entry must decrease, thus implying a progressive reduction in stage pressure ratio through the unit. Hence the rear stage develops a significantly lower pressure ratio than the first stage.
Higher stage pressure ratios are also possible if the relative velocity between fluid and rotors is supersonic, but this is achieved at the expense of efficiency and operability. Such compressors, with stage pressure ratios of over 2, are only used where minimising the compressor size, weight or complexity is critical, such as in military jets.
The airfoil profiles are optimised and matched for specific velocities and turning. Although compressors can be run at other conditions with different flows, speeds, or pressure ratios, this can result in an efficiency penalty or even a partial or complete breakdown in flow (known as compressor stall and pressure surge respectively). Thus, a practical limit on the number of stages, and the overall pressure ratio, comes from the interaction of the different stages when required to work away from the design conditions. These “off-design” conditions can be mitigated to a certain extent by providing some flexibility in the compressor. This is achieved normally through the use of adjustable stators or with valves that can bleed fluid from the main flow between stages (inter-stage bleed).
Modern jet engines use a series of compressors, running at different speeds; to supply air at around 40:1 pressure ratio for combustion with sufficient flexibility for all flight conditions.
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 A. A. 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 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 Royal Aircraft Establishment, Hayne Constant. Other early jet efforts, notably those of Frank Whittle and Hans von Ohain, were based on the more robust and better understood centrifugal compressor which was widely used in superchargers. Griffith had seen Whittle's work in 1929 and dismissed it, noting a mathematical error, 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, Hayne 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 including the world's first jet aircraft (He 178), but development efforts had moved on to Junkers (Jumo 004) and BMW (BMW 003), which used axial-flow designs in the world's first jet fighter (Messerschmitt Me 262) and jet bomber (Arado Ar 234). 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 latter a turboprop. Northrop also started their own project to develop a turboprop, which the US Navy eventually contracted in 1943. Westinghouse also entered the race in 1942, their project proving to be the only successful one of the US efforts, later becoming the 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.
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 the 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.
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, pinoeered on the Bristol Olympus, 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 as there is a trade off between higher fuel efficiency and the higher maintenance involved pushing up total cost of ownership compared to a two spool design. 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.
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.
For jet engine applications, the whole idea of the engine is to move air to provide thrust. In most cases (the exception being high-altitude, high-speed flight), the stoichiometry of the fuel-air mixture dictates an engine which produces far more shaft horsepower than is needed by the compressor stage to provide compressed air for combustion. In an engine with zero bypass all the air passing through the engine is combusted resulting in a very hot, high velocity, exhaust. Consider that thrust is calculated from the mass flow (in Kg/sec) passing through the engine multiplied by the difference in velocity (in m/sec) between the inlet and exhaust velocities. Note that this a linear function. However, the kinetic energy in the exhaust is derived from the mass flow multiplied by one half the square of the difference between the inlet and exhaust velocities. Therefore, a doubling of the velocity difference (with a fixed mass flow rate) results in four times the energy imparted to the exhaust while only a linear doubling of actual thrust. In contrast, by doubling the mass flow rate (while keeping the same inlet/exhaust velocity differential) results in doubling of thrust with only a linear doubling of energy imparted to the exhaust. Note that energy imparted to the exhaust is not available to provide thrust. By using the excess of shaft horsepower to accelerate air (which is not combusted) through the engine increases the mass-flow rate while reducing the overall inlet and exhaust velocity differential, increasing the specific impulse of the system. The limiting factors for increasing the bypass ratio are the available shaft horsepower (over and above that which is required by the compressor stage for a steady-state power output), and the increased weight of a larger fan mechanism, which on an aircraft subtracts from either the payload or fuel carrying capacity.
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 1950s Conway engine, which enlarged the first compressor stage to be larger than the engine itself. This "bypass engine" 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. This 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.
Negatively affecting this saving is the fact that drag increases dramatically 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."
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 which were hampered by the unavailability of high temperature metals. 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. Improvements started immediately and have not stopped; the latest Rolls-Royce Trent operates at about 40:1, twice the ratio of diesel piston engines.
Since compression ratio is strongly related to fuel economy, this eightfold increase in compression ratio results in an 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.
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.
The blade rows are designed at the first level using velocity diagrams. A 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 an angle β; larger turning allows a higher temperature ratio, but requires higher solidity.
Modern blade rows have low aspect ratios and high solidity.
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.
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.
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