The Compressor: Gas Turbine Engine Keystone

Abstract

All engines involve the raising of pressure in the engine’s medium, be it liquid compression of water, piston-cylinder compression of air or compression by aerodynamic forces. The last is the hardest because it is required to be efficient and stable. An aerodynamic air compressor works by slowing velocities to raise pressure. The airfoil has a difficult time doing this reliably and well and so does the compressor. When it does, it allows for amazing performance of the engine in terms of power and efficiency.

Compression of air by dynamic means was a known technology. The supercharger impeller in the R-4360 shown in Fig. 9.1 is a radial flow compressor that can achieve pressure rise factor on the order of three or so, a level that is sufficient to make a jet engine practical. For such an engine, the compressor impeller wheel just had to be made sufficiently large in size. It functions by taking in air and hurling it out in a radial direction with speeds that approaches the sound speed. The outflow air is then gently slowed so that the air’s kinetic energy is converted to pressure. A flow tube with a gradually increasing area (called a diffuser) does that with an adverse pressure gradient as discussed in connection with flow separation from the upper surface of the wing. Since the diffuser has walls, boundary layers grow on them and they are subject to the same flow separation issue as we had on the upper surface of a wing.

The impeller of the radial flow compressor has problematic aspects that include a rather large surface area on the impeller and its guiding vanes. Thus, friction is experienced on those internal surfaces with an impact on efficiency. Another is the awkward geometry where the diameter of the compressor as a whole has to be about 3–4 times the size of the inlet duct. That limits its application into a sleek airplane or engine nacelle. Finally, the mass flow rate is rather modest but, in practice, that was partially addressed by having the impeller have two sides into which air flows. On the positive side, the compressor worked quite well. Frank Whittle and Hans von Ohain both used this kind of compressor on their early engines. Figure 10.1 shows a General Electric an I-16 and a later I-40 engine derived from the cooperation between Great Britain and the US during the war. Both engines have double-sided radial flow compressors.

Fig. 10.1

Top: A General Electric I-16 (military designation J31, ca. 1942) engine on display at the Smithsonian Air and Space Museum. Note the double-sided radial flow compressor. Most of the diffuser ducting has been cut away but the air passages can be discerned. They include a 90° turn into the combustor. A better view of the diffuser ducting can be seen on an earlier Whittle engine in Fig. 12.5. Further diffusion takes place at the combustor inlet with a large flow area increase. A single stage turbine driving the compressor exits flow to the left. The lower figure is of a more advanced (early 1943) jet engine of a similar design (jet flow to the right) but with an in-line rather than reverse flow combustor (GE I-40, also known as the (Allison produced) J33) (Courtesy General Electric)

The I-16 engine had a military designation of J31, with an airflow rate of 33 lbs/s (15 kg/s), a compressor pressure ratio of 3.8, a turbine inlet temperature of 1220 °F (660 °C), and produced 1600 lbs¹ of thrust. This was a small engine. The turbine shaft carried about 3400 hp to the compressor. The power in the jet at sea-level static conditions is about 2400 hp. These numbers are quite similar to those of the Jumo 004 engine cited earlier. To put these numbers into perspective and appreciate what a technology step that the gas turbine was, consider that the P&W R-4360 engine, the largest production ICE engine that ever saw service application, produced an output of over 4000 hp processing about 9 lbs/s (4 kg/s) of air (at sea level). These engines are very different even with the gas turbine in its infancy.

We could discuss the radial flow compressor type in greater detail, but the reality was that another type proved to be superior. Part of the reason is that an analysis of the engine as a propulsion device points in the direction of a need for higher pressure capability from the compressor. A multiplicity of radial impellers operated in a series would be possible but quite awkward in practice for it would mean processing the air radially in and out a number of times. That awkwardness does not mean, however, that this approach was not used. To wit, Rolls Royce built a very successful Dart engine based on this principle. Today, the radial flow compressor is used in a number of smaller turboshaft engines in combination with the more promising and now much more common axial flow compressor.

The history of Whittle’s efforts to build his first radial flow compressor engines with limited private and military resources is an interesting story. It speaks volumes of his tireless effort and his belief in the engine. It would indeed ultimately be accepted as potentially important for the military. As that was realized by the British military, his firm was incorporated into Rolls Royce, presumably for better management and access to the necessary financial and technical resources.

Von Ohain and Heinkel also tried to progress in developing better jet engines. While they enjoyed some success in garnering the interest of military funding early on, they eventually were beaten by other industrial concerns (Junkers and BMW) and succumbed to the needs of a war that was not going well.

Looking at the history of Whittle and von Ohain, it is possible to conclude that great ideas may come from individuals, but they soon lose control over them as powerful industry interests are drawn to the challenges and opportunities.

10.1 Axial Flow Compressor: The Bedrock of Modern Engines

A new compressor design approach was developed in Germany as well as in Great Britain. In Germany, the approach to finding the best compressor configuration for a jet engine was more deliberate and entertained a variety of approaches. The military establishments on both sides of the war were reluctant to spend funding and manpower resources on a new engine development program whose benefit might not be realized in time to affect the outcome of the war, then raging. Nevertheless, in Germany, two individuals in the Luftwaffe establishment managed to foster a funded program for a jet engine. The largest question was: how does one best build a compressor with good performance and good growth potential—within the context of a practical engine?

The best place to start an investigation about the inner workings of a jet engine is to examine one. Figure 10.2 is an image of a 1960s jet engine that is technically a fanjet engine (or turbofan) and it illustrates most of the attributes of both types of engines quite well. A fanjet contrasts with a turbojet in that the latter is not equipped with a fan. The first generation of jet engines were the simpler turbojets.² In this later turbofan, the fan is the first rotating set of blades that supplies air to a separate exit nozzle as well as to the engine’s compressor (nearer the central rotor axis). The purpose of the fan is to increase the propulsive efficiency by increasing the air flow rate processed and reducing the average jet velocity. There were other advantages to the fanjet over the turbojet that led to their rapid adoption by the airline industry as soon as they became available: reduced noise and higher thrust at takeoff conditions when the necessary runway length was an issue.

Fig. 10.2

Cross-section image of a P&W JT8D turbofan engine. The first set of blades at the flow entry at left is an inlet guide vane (Courtesy R. Lindlauf, New England Air Museum)

We should note that the compressor and turbine in this particular engine are sectioned into two segments each, running on separate shafts rotating (one inside the other) at differing speeds. The components are referred to as high- and low-pressure compressor and with similar appellations for the turbine. The low-pressure shaft rotates inside the hollow shaft connecting the high-pressure components. The innovation of using two spools by Pratt & Whitney allowed the pressure ratio to be elevated significantly over a single shaft engine and it gave better off-design performance.

The speed(s) of the engine shaft(s) is one of the elements of engine performance displayed to the pilot of a jet airplane, usually expressed, not in terms of rotations per minute (RPM), but as a percentage of a nominal maximum value.

When this configuration was perfected,³ it allowed for the building of two airplanes that had a profound impact on the jet engine industry, namely the US Air Force Boeing B-52 and the Boeing 707. The pressure ratio for this J57 engine compressor was around 12 (see Fig. 12.3) and for the later turbofan engine that employed the dual spool compressor concept illustrated in Fig. 10.2, it was almost 20. This turbofan powered several airliners built in the 1960s and 70s. A more modern GE90-115B turbofan engine (see Fig. 11.2) powering the Boeing 777 bettered this compression ratio with a value of over 40. The latest engine offerings by manufacturers in 2020 are 50:1 at Rolls Royce (Trent XWB) and 60:1 in (GE9X) in very efficient, very high bypass ratio engines.

To get a sense of the power modern engines process, at takeoff conditions, the turbine of the GE90-115B turbine supplies the fan with an estimated 120,000 hp and the compressor with more than 70,000 to produce 52 tons of thrust at takeoff processing almost one and a half tons of air each second! The power involved in the jet itself as the airplane sits at full power on the runway is difficult to estimate without access to proprietary data, but it can be safely said that it easily exceeds 100,000 hp. It is hardly surprising that when the gas turbine age dawned, builders of large airplanes never considered using the ICE again. The gas turbine engines made large, long range airplane possible.

10.2 Compressor Pressure Ratio

To this point, we have been rather vague about the importance of the compressor pressure ratio as a parameter in the design of a gas turbine engine. It must obviously be greater than one (no compressor), but how large? This is where the distinction between the piston (ICE or Otto⁴ cycle) and the gas turbine (or Brayton cycle⁵) engines is starkly apparent. Fuel burning for the ICE is designed to burn out all the oxygen in the air to get the most heat and pressure from the process. In that engine, the compression pressure ratio (outlet to inlet) that the air is subjected to is strictly limited to what can be compressed without pre-ignition. In the compressor of the gas turbine air is not compressed with fuel as a fuel air-air mixture and there is no parallel limitation on compression. The compressor pressure ratio can be chosen to meet other criteria.

To be specific about this, we deal with two important temperatures that arise in connection with discussion of the gas turbine. These are the turbine inlet temperature and the compressor exit temperature. For brevity, these total temperatures will be called TIT and CET in this paragraph only. Heaven forbid we start speaking in engineering jargon! The CET is determined by the compression ratio as the compressor is close to isentropic. For example, the compressor outlet temperature from a compressor with pressure ratio of 40, operating in the stratosphere (- 60 °F, - 51 °C) is more than 800 °F (425 °C). This temperature will be even hotter if the compressor is designed to have a higher value of the pressure ratio. The ultimate limit is reached when the CET reaches the maximum allowable TIT. At this design point little or no fuel can be injected and the oxygen available is plentiful. In such a design, the stoichiometry (the relative amounts of chemical reactants involved: fuel and oxygen) is fuel-lean, quite lean, in fact. However: no heat, no work. Evidently, there must be a middle ground between compression ratio one and the value that leads to the CET raised to equal the TIT. That is indeed the case. The virtues of an engine are two thermodynamic ones: work delivered per unit mass flow air processed (specific work) and work delivered per unit of heat provided (thermal efficiency). An engine with a high specific work will be compact and light in weight because the machinery involved will be appropriately modest while an engine design with high efficiency will be heavier but sparing of fuel use.

These two design conditions do not occur at the same compressor pressure ratio. In fact, as the pressure ratio is increased (in a design sense) the maximum specific work configuration is reached first, followed at a higher pressure by the maximum efficiency point. Thus, a choice for a preference must be made. The engineer would illustrate this with a cool graph, but we are not doing this here! (He would also have jargonized compressor pressure ratio as CPR!). The optimum pressure ratio has to be examined in the context of the question: how is the engine going to be used? Long distance flights require high efficiency for low fuel consumption while short flights benefit from not having to drag a heavy engine up to flight altitude and then back down, relatively often.

10.3 Compressor Aerodynamics

So how does the blading in these compressors work? The compressor is a machine that exploits the aerodynamics of wings. After all, the propeller is a compressor of sorts, why not build on that? Actually, the propeller is not a compressor, but it does accelerate a flow and the flow so produced could be slowed to raise pressure. A great advantage of flows in a compressor over similar flows in a freestream is that it has to be an internal flow with boundaries. That means that bound vorticity on the rotating and any stationary blading will largely avoid the creation of the trailing vortices that propeller blades must live with. A second advantage is that the flow behind such blading can be made relatively uniform, a virtue the propeller blade does not enjoy to the same degree. Finally, the turning that such blading imparts to the flow can be undone by stationary blades behind rotating ones. In order to do that, the compressor would have to be built with two kinds of blades, some that rotate and others that do not. Such a blading combination is called a stage and is illustrated in Fig. 10.2, when one looks closely. This arrangement is not unlike the two elements of a radial compressor with its a rotating impeller and a stationary diffuser. The same can be said about the blading combination of a counterrotating propeller.

Unfortunately, in order to fully understand how an axial flow compressor works, we have to draw diagrams that show the changes in speed and direction of the flow. According to Bernoulli, speed changes that slow the flow result in the pressure increase we seek. A matter that is slightly complicating is that we will have to look at the situation in the reference frame of the blades themselves. In other words, looking at the rotor blades (the rotating ones) we have to look from a rotating viewpoint whereas the goings-on around the stator blades (the stationary ones) can be safely observed in the laboratory.

The flow through the compressor annulus (the space between the inner and outer cases) is helical in nature for reasons that will become clear further on. Commonly made representations of flow through such a compressor as straight-through parallel to the rotational axis involve an excessive use of artistic license and are inaccurate. Air that enters the compressor at 12 o’clock on the face may exit the compression process at 8 o’clock or thereabouts, depending on the design, specifically the compressor pressure ratio. For that reason, the descriptive word “axial” flow compressor is somewhat misleading, but we will stick to it because it is the accepted convention. To set up a helical flow pattern, the incoming axial flow will have to be turned by what is called an inlet guide vane (IGV). The first set of (stationary) blades on the engine in Fig. 10.2 are IGVs. This blade set turns the flow and necessarily accelerates it. The pressure will fall through this set of blades. That drop will be made up at the compressor exit where the last stator redirects the flow in the axial direction (in most design cases). Figure 10.3 shows the inlet guide vane flow velocities⁶ where the axial direction speed is more or less conserved because mass flow rate has to be conserved. The “more or less” words have to do with flow area and density changes that may also be involved in such designs. In some compressor designs, the inlet guide vane is omitted. This omission will be clearer when we look at the fan in that family of engines.

Fig. 10.3

Velocities and geometry of an inlet guide vane. The velocity in the direction of the engine axis is roughly conserved. In this figure and the several to follow, the leading edges are drawn as sharp. In practice they are rounded to allow for a variation in the in-flow angle of attack

The fan flow is the outer annulus part of the flow that is directed around most of the engine. The fan as a topic will be addressed again after we explore the basic elements of the engine. In the example engine shown in Fig. 10.2, the fan has a single rotor with an exit guide vane (visible at the entrance to the bypass duct) to align the flow direction to the rear without any rotating flow component.

The flow situation for the compressor stages that follow the inlet guide vane is illustrated in Fig. 10.4. The rotor turning as indicated accelerates the flow away from the axial direction as shown. The rotor exit velocity in the laboratory reference frame at (2) is greater than at (1). The stator’s function is to restore the angle of the flow (3) to roughly the orientation it had before entering the rotor (1). In its absence, the flow would be turned even further away from the axial by a following rotor, an impossible situation if more than a few rotors were employed. Without stators the result would be unacceptably high flow velocities and a limitation of the air throughput. As with a stator, the lengths of axial velocity vectors would also be conserved as with the inlet guide vane. The desired rise in pressure in the stator is realized because the incoming velocity (2) is larger than the outgoing (3).

Fig. 10.4

Flow velocity vectors in the laboratory reference frame through a compressor stage. The axial components of all three vectors (not shown) are the mass flow determining flow velocities which are necessarily conserved. The rotor increases the flow velocity (from 1 to 2). The stator restores the flow direction (3) to approximately that of (1). The rotor details are illustrated in Fig. 10.5

How does the rotor increase the velocity as indicated? To visualize the velocities in the rotating reference frame, we must add the rotational velocity (red vectors in Fig. 10.5) and examine them in the rotating frame of reference. The green vectors (L1 and L2) in the laboratory reference frame are the same as those shown in Fig. 10.4 (1 and 2). The solid black arrows (R1 and R2) are the velocities in the rotor’s frame. They must align with the geometry of the rotor blade at the leading and trailing edges. Proceeding through the rotor blades reduces the relative velocities and the pressure rises. This statement requires proof and fortunately one is available, but not developed here to keep the narrative simple. In sum, pressure rises in both the rotor and the stator.

Fig. 10.5

Velocity vectors in the blade row of the rotor in its rotating reference frame. R1 is the relative velocity into the blade row and R2 is relative out. Green vectors (L1 and L2) are the laboratory reference frame velocities shown in Fig. 10.4 and the red lines are the rotational speeds for the change of reference

10.4 How Well?—Efficiency

The calculation of the pressure rise through a compressor blade row or stage is a simple application of the same energy equation used for determining the pressure on an airfoil, namely Bernoulli or its compressible equivalent. The velocities in play are those relative to the blades. That statement applies to both rotor and stator blading. The air state variables (temperature and density) will also rise as one proceeds from one stage to the next. The degree to which they change is easily related to the changing pressure through what is called a stage adiabatic efficiency. Such an efficiency relates the actual property changes to changes associated with an isentropic process, the accomplishment of which is always a design goal. While no one has ever been successful in devising a flow that perfect, the stage efficiency percentage is in practice quite high, reaching and exceeding 90%. It quantifies the modest frictional losses that occur. Its exact value for the compressor as a whole (the so-called adiabatic compressor efficiency) is of great interest to the engine performance evaluator. Its value can be gleaned from experimental testing or from detailed examination of the friction losses associated with the flow through the blading.

There is not much to talk about vortices in the flow through an axial flow compressor except that the boundary layers do contribute to rotational flow that ends up as turbulence. Rest assured, however, that the lifting airfoil behaves very much like the wing with a distributed vorticity distribution along the chords of the blades. The Kutta-Joukowski condition applies to fix the flow orientation at the trailing edge and the leading edges of the blades are rounded⁷ to accommodate a variation in the in-flow conditions.

As with the wing, one can ask only so much of a blade row before the turning or angles of attack are too large for air to follow your geometric imposition on flow direction. The blading of a compressor can stall. Naturally, the compressor would want to be designed to be operated as close to stall as practical to realize as much compression per stage and to keep the number of stages to a minimum. As a matter of record, we note that early on in the history of axial compressor development, the pressure ratio per stage has grown from 1.15 (J47, 1947) to an estimated average of 1.28 (GE 90, 1995). This is no small feat because it dramatically reduced the size and weight of the compressor.

In practice, the number of stages in any compressor is significantly larger than the number of turbine stages because the compressor air moves into a region of higher pressure. This adverse pressure gradient leads to the same flow separation issues associated with the wing aerodynamics. The turning that can be asked of each compressor blade is rather small and contrasts dramatically with the large turning angles in a turbine where the pressure falls. Considering that the power produced by the turbine is almost identical to that absorbed by the compressor, blading aerodynamics obviously plays a large role in determining the configuration of a gas turbine engine.

10.5 The Control Problem

The axial flow compressor operates with blades at an angle of attack to the air streaming through the compressor. The airstream speed and direction relative to the blades are controlled by the speed of rotation of the rotor blades and the engine’s air throughput, i.e., the air axial velocity. If that combination of speeds leads to blade stall, the compressor becomes partially or totally nonfunctional. This possibility can occur when the pilot demands rapid changes in thrust. The pilot might therefore be confronted with the delicate operation of the compressor as well as that of the airplane. That is an untenable situation, especially in the military setting of the mid 1940s. An automatic control mechanism was and is necessary to shield the pilot from concern about the compressor operation.

Axial flow compressor development in Great Britain (by Metrovick) initially failed to manage such control and, in flight testing, the axial flow compressor was judged, at the time, to be unworthy of further development because of the excessive workload on the pilot. On the other side, a control system was devised by the German companies Junkers and BMW.⁸ It was crude but effective. It consisted of an airflow rate control mechanism in the exhaust nozzle that, by measurements linked to it, always held the flow rate through the engine so that angles of attack on the compressor blades were always in a functional position and the compressor avoided stall.

It may be interesting to note that the German development effort initiated to find and produce a good jet engine involved a number of firms that include the two mentioned above as well as Heinkel (where Hans von Ohain was active) and Daimler Benz. The compressors considered included both radial and axial elements as well as hybrid radial/axial devices that spun the outflow, not in the radial direction, but rearward at roughly 45°. Most interesting is that Daimler Benz proposed a configuration that was to be a fanjet. This part of the history is described in greater detail in the bibliography reference entitled “Turbulent Journey” by the present author. When the war tide shifted against Germany, the scope of the jet engine development program was reduced to the projects by Junkers (Jumo 004) and BMW (BMW 003) with only the Junkers effort leading to a production program and the engine’s use in primarily two airplanes, the Me-262 fighter and the Ar-234 bomber. The Daimler recognition of the advantages of the turbofan engine was not singular. Back in the mid 1930s, Frank Whittle in Great Britain had recognized the advantages and patented the idea. The difficulties of developing the simpler turbojet were sufficiently challenging that he let the patent lapse.

That German control technology was shared with Allied victors who quickly recognized the value of the axial flow compressor and its advantages. Rather quickly, engineers learned to adapt more sophisticated control systems on the compressor. These successful steps ultimately led to the dominant use of the axial compressor in a large majority of engines. There were exceptions to be sure, but they were few and for special reasons.

Modern turbojet and fan engines employ electronic engine control systems to avoid stall. These systems manage the rate of power lever changes demanded by the pilot so that they are responsive and safe for compressor operation. A mechanical part of such systems may include variable stator angle control on the compressor so that flow angles are always close to optimal. In Fig. 10.4, the flow angle 2–3 is variable to suit such needs. Additional controls are necessary for starting gas turbine engines and these include bleed valves midway in the compressor.

The most common summary of performance of a compressor is on a so-called compressor map. It is somewhat akin to a drag polar of a wing and presents most of the important information regarding the characteristics of a compressor. It consists of a plot of pressure ratio versus air weight flow rate. Lines of constant rotational speed and constant efficiency are noted. The efficiency will have a maximum value at some point. Also shown is the line along which the compressor stalls. Stall is encountered at low air flow rates. The control system’s function is to keep the point of operation in the region of highest efficiency and away from stall. Air weight flow rate and rotational speed (usually in RPM) are corrected to reflect inlet conditions to make the map as compact as possible. The corrections reflect inlet conditions that differ from those of a standard day, typically defined for aeronautical purposes as 29.92 in/Hg (14.7 psia or 101 kPa) and 15 °C (59 °F).