Other Components of the Jet Engine

Abstract

The compressor may be the challenging component of a gas turbine engine, a combustor and a turbine also must function efficiently. The design aspects of these two components are described here together with how the engine is configured and works as a system.

12.1 The Turbine

The turbine is the source of all the mechanical power in the engine. It can also be configured in radial or axial inflow configurations. An example of the former is similar to a child’s pinwheel. Only Hans von Ohain used such a turbine in the first jet engine that flew in the Heinkel 178. Von Ohain and everybody else understood that the better way to go was to use an axial turbine, adapting the knowledge from steam turbines for power generation that had been in commercial service for over a decade. The radial turbine is now just a footnote, although one can find it used in automotive turbochargers where operating efficiency requirements are not as strict as they would be for an aircraft application (Fig. 12.1).

Fig. 12.1

A cutaway of an automotive turbocharger. Ambient air is drawn in from the left and exits in the spiral diffuser (blue). The red zone is the radial inflow (exhaust from the reciprocating engine) to the turbine wheel. It is relatively simple and performs adequately well (Photo: Quentin Schwinn (NASA))

The description of the function of an axial turbine parallels that of the compressor. Rotors and stators are involved. The first element is stationary and is called a nozzle. This nozzle gives the flow kinetic energy in the azimuthal (rotation) direction which the rotor that follows it can harvest. The gas exiting the nozzle may be supersonic. The relative velocity incoming to the rotor will be lower than the exit velocity from the nozzle because the turbine rotor blades are receding away from it.

In contrast to the compressor where the flow in its stages is adequately describable as locally incompressible, the turbine flow is definitely in the compressible regime. The blading operates with a favorable pressure gradient so that large flow turning angles can be withstood without concern for stall. This aspect of the difference between compressors and turbine explains the physical difference between these components. Figures 10.2 (and 12.3) shows that the power transmitted from the turbine to the compressor is generated with relatively few turbine stages while the compressor absorbs the same power as input but must process the air gently (against an adverse pressure) to avoid the stall issue. To be specific, the P&W JT8D engine in Fig. 10.2 has a 1-stage fan and 13 (6 + 7) stages of compression, all driven by 4 (1 + 3) stages in the turbine. The contrast between compressor and turbine blading is starkly illustrated when we note that a single turbine blade row in the high-pressure turbine can power six stages of the compressor.

The image of turbine nozzles and blades with their cooling air bleed holes (Fig. 12.2) suggests that the compressor feeds more than just the combustor. It also provides the cooling air for the turbine. The word “cooling” is a bit curious because the compressor air is, in fact, hot. It is, however, cooler than the combustor exit air and thus effective as a coolant. It must also be at the highest possible pressure to be able to enter the turbine flow through the many small holes. Identifying an optimum amount of cooling air from the compressor has to consider the positive aspect of operating at effectively higher turbine inlet temperature against the compressor power used to provide the cooling air. In modern engines, the cooling bleed is in the range of 5–15% of the air through the compressor.

Fig. 12.2

Two blades of an aircraft gas turbine nozzle (one blade is covered by yellow tape) and rotor in the shop. This display is set up for training purposes. The rotor blades on the wheel to the left of the nozzle block are covered by orange plastic to protect the students. To the right of the nozzle element, the rotor blades are naked and turn the flow through a rather large angle. The single block element of the nozzle is held in place with a plastic tie. Note the heavy use of cooling air holes to bathe the nozzle material in cool air and minimize contact with the hot combustion gas. The second turbine stage is also shown as two stator sections (4 blades) (Courtesy General Electric, photo by author)

Fig. 12.3

A P&W J57 (or JT3C) in sectioned model form showing clearly the flow area expansion at the end of the compressor. The model illustrates the late 1950s technology of combustor design and the two compressor segments: low- and high-pressure elements powered by separate sections of the turbine with a connection by two shafts, inside one another. The artifact is located at the Wings Over the Rockies Air and Space Museum, Denver CO (cropped photo by Ryan Frost, https://commons.wikimedia.org/wiki/File:Sectioned_Pratt_%26_Whitney_J57.jpg)

Looking back at the ICE that also employs cooling air to maintain the temperatures of the cylinders, we note that the fuel energy transferred to the cooling air in that engine is lost for propulsion purposes. The gas turbine, by contrast, loses very little fuel energy heat as a pure loss. The mass of cooling air stays in the flow so that both heat and mass end up in the propulsion jet and contribute to the thrust produced.

12.2 The Combustor

We have largely left the topic of vortices to describe how a compressor and turbine work. To first order, vorticity is not a major issue in describing the workings of a jet engine except to say that the bound vortices on the blades of both compressor and turbine must be in play. In the combustor, the vortex will be seen to play an important role. First, however, a look at the functionality of this component.

Combustion in a jet engine or better, a gas turbine, takes place at nearly constant pressure. The combustion process there takes place just as it does in a candle, the fireplace, or the burner of a hot air balloon. The consequence of burning fuel at constant pressure is that the air involved is heated and, being a gas, it expands proportionately in volume. The ideal gas law, pv = RT, is in play here. The resulting lower density (= 1/ν) causes the heated air to rise in the fireplace chimney or, when contained by it, causes the hot air balloon to experience a net buoyancy and rise in the cooler atmosphere.

In reality, the (total and static) pressure in the combustion chamber of a jet engine falls a small amount for two reasons: pressure losses associated with the flow through the hardware and the process of adding heat to a flowing medium always leads to a loss in pressure. Any pressure drop is certainly to be avoided because the compressor had to provide it. From a technical viewpoint, one can show that this loss is minimized if the Mach number of the flow to which heat is added is kept small. This is done in a jet engine design by forcing the flow to slow with an area increase as it enters the chamber (see Figs. 12.3 and 12.4).

Fig. 12.4

Simplified sketch of a gas turbine combustor showing the swirler and other features and functions. The purpose of the swirler is to form a rotating mass of burning gas with helical vorticity on its outer edges. Simultaneously, the two illustrative rotational cells serve to continuously ignite the fresh air charge by bringing burning fuel forward (https://en.wikipedia.org: file: combustor diagram airflow.png by SidewinderX). The lower sketch is a simplified illustration of the changes in gas flow properties (total temperature in red and total pressure in solid green; static pressure in dashed green) that occur in a typical combustor that, in reality, is quite three-dimensional in nature

In the chamber itself, three major challenges are presented to the designer. These are, first, reducing the volume of the space necessary to carry out combustion to a minimum. Specifically, the length of the combustor must also allow the combustion process to be complete. This is a problem because flame propagation speeds are very much lower than the flow speed through the chamber even when the Mach number there is held low. Here the vortex comes to the rescue. At the burner inlet one or more vortices are formed so that the swirling motion lengthens the residence time of the burning gas. This usually done by swirling the flow with an array that resembles an inlet guide vane (see Fig. 12.4a). The goal is to avoid forming soot, unburned hydrocarbons and simultaneously avoid forming pollutants such a carbon monoxide and nitrogen oxides. The second challenge is to produce a gas uniformly high enough in temperature acceptable to the turbine without itself being at risk of material failure due to overheating material surfaces or chemical oxidation. The uniform aspects have to do with the desire to avoid subjecting the blades to a variation in temperature that is hot and hotter at a high frequency. Lastly, the combustion must be stable in that it is, once lit, continuous and free of needing attention.

In practice, the burner of a gas turbine engine is operated as a combination of two processes: primary combustion where the chemistry is such that most oxygen is burned to achieve a stable and very hot flame and a secondary zone where additional air dilutes the combustion gas to a temperature that the turbine can tolerate. Figure 12.4 is a schematic representation of the process.

The lower part of the figure is a simplified sketch of the variation of the important flow parameters through the combustor. The red line shows a one-dimensional (averaged over the available flow area) variation of the total temperature, noting specifically the very hot primary combustion zone and the subsequent cooling by dilution with cooling air.

Also shown in the sketch is the variation of total and static pressures. The total pressure falls slightly during the process of heat addition to the air. The relative value of the static pressure is a reflection of the local Mach number. In diffusion, the total pressure stays constant while the static pressure rises because of the flow area increase. In the burner, the averaged flow behavior is for the Mach number to increase (toward thermal choking but not coming close to that) so that the net result is a decreasing (average) static pressure. The area decrease in the aft half of the combustor as shown in the upper sketch is an additional reason for the static pressure to decrease in the flow direction. While complicated, the design of the combustor mercifully involves subsonic, primarily low Mach number aerodynamics.

One cannot speak about the combustor without mentioning its pivotal role between the compressor and the turbine. The amount of fuel added to the flow must be just the right amount to raise the air temperature from that exiting the compressor to that acceptable by the turbine. The compressor air is hot from the compression work that went into it. Here is where the compression ratio of the compressor really matters.

In any functioning gas turbine engine, the amount of fuel added for combustion is limited and there is always more air processed than the amount necessary to burn the fuel. The excess oxygen flowing through the turbine has two consequences: the environment for the turbine is an oxidizing one that has the potential for setting the stage for material failure due to oxidation of the hardware in the burner and in the turbine. Good design learned from experience in the fields of metallurgy and cooling technique have largely put this concern to rest. The other consequence is that the left-over oxygen can be burned in an afterburner, i.e., in another combustor behind the turbine to obtain thrust for applications and situations where it is needed. An afterburner was used in the supersonic Concorde airliner to achieve cruising speed although not used during cruise. It is used primarily in military fighter jets.

Considering the progress made starting from the early engines of the 1940s to today’s reliable machines, the goals in the combustor have largely been brought to a high degree of perfection. Included in the list of desired goals is the need to minimize undesirable combustion product emissions into the atmosphere. The burners have been so reduced in volume that they are hardly discernable as a component of the engine. Failures in operating engines are seldom attributable to the combustor, except perhaps when maintenance is given a short shrift. The combustion chambers and the region around the turbine entry are, therefore, the subject of scrutiny by the people who maintain jet engines.

Figure 12.5 also illustrates the need to keep the combustion chamber short. A long chamber will involve a long rotating shaft connecting the turbine to the compressor. If the shaft is insufficiently stiff in bending, it will be unstable. In the J47 of Fig. 12.5, engineers dealt with this problem by installing a bearing in the middle of the shaft.

Fig. 12.5

Left: The voluminous combustion chamber on an early British engine, the Whittle W1 on display at the Smithsonian National Air and Space Museum. At right, the burner cans on a GE J47, an early (1947) single shaft, axial flow compressor engine (its turbine is at right under the plastic), on display at General Electric

In practice, the combustion chambers of a gas turbine are configured in one of two arrangements: straight-through flow or reverse flow. In most modern engines, the ability to achieve combustion in a small volume has led combustors to the use of the straight variety. The various engine cross-sections in this text illustrate this arrangement. Examples of the reverse flow configuration are shown in Figs. 10.2 (I-16) and 12.5 (Whittle W.1A). The reverse flow combustor allows for a short shaft between turbine and compressor.

12.3 Putting It Together into an Engine

How do the jet engine and its components work to produce thrust? We mentioned that the compressor and turbine function very efficiently. If the component efficiencies were an impossible 100%, then a compressor and turbine hooked together and could function without any consequence to the universe. They aren’t, of course, so we have to pay for what they do. How could we do this? … and the answer is: by giving the turbine a higher volume flow rate to process. The only practical way to do that is to heat the gas prior to turbine entry so that it expands in volume—again, like the flue gas in the chimney.

We could start by adding just the right amount of heat to the air entering the turbine so that the losses in both components (compressor and turbine) are overcome and the engine operation is self-sustaining. That achievement probably brought the first reason for celebration to the pioneers Whittle and von Ohain! We could then add more heat. As the engine (of a fixed geometry) runs, the air mass flow rate through the engine must be conserved. It depends only on the product of flow velocity and air density. That means for the turbine that as the density is lowered by heating, the velocity must rise at every locale. In turn, the forces acting on the airfoils must increase because they depend on the dynamic pressure which depends on density and velocity squared. Heating the gas increases the forces, torque, and power from the turbine!

Whereas our ideal or self-sustaining operation left the outlet pressure behind the turbine at an atmospheric level, a greater heat supplied increases of the pressure at the turbine outlet because not all the pressure drop through it is needed to run the compressor. There is pressure left over. The engine becomes a pump: it processes air and delivers it at a higher pressure! With the addition of a nozzle, we can create a jet with (more) momentum to get thrust! Hot dog! More thrust is obtained by adding more heat up to the temperature limit that the turbine can tolerate.

The same physical argument about component entry temperature can be used (in reverse) to talk about the compressor where the work (i.e., the power) required for compression is also temperature dependent. Thus, if the compressor inlet air temperature is lowered, the power required is also lowered. One may conclude that the engine would best operated between temperature extremes that are as wide as possible. This is one reason that airplanes with jet engines operate so well in the cold stratosphere, where the temperature is around - 60 °F. Conversely there might be difficulties associated with operating an engine on very hot days when performance will be suboptimal.

The idea that wide temperature extremes are best for extracting mechanical work from a heat engine is a conclusion that the Frenchman Sadi Carnot (1796–1832) reached when he investigated the convertibility of heat to work generally. His examination of the performance of what we call heat engines laid the foundations for the Second Law of Thermodynamics that involved the identification of energy (through the First Law) and entropy as properties, the concept of reversibility, the notion of temperature, the role of the temperature extremes that apply to heat engines and limit their performance. His studies launched a great deal of understanding of the workings of nature and that knowledge fueled an important second period of the Industrial Revolution. The first having been centered on the use of the steam engine.

A truth that follows from the work of Carnot and others is that a perfect heat engine (a Carnot engine) can be imagined, although building it is not practical and may not be possible. The gas turbine engine, like all engines, is limited by the temperature extremes available. Specifically, the successful generation of mechanical power by means of any heat engine depends on heat from a high temperature source (or reservoir, as thermodynamicists like to say) intercepted and converted as it flows to a low temperature reservoir (a heat sink). Without both of these reservoirs, nothing useful can be made to happen. Consider, for example, that the heat engine that is planet earth and the life on it operates between input from the sun (~ 5800 K or some large number of degrees F) and radiates the waste energy to the near zero absolute temperature of space (~ 3 K). Consider also that a uniformly hot environment is not a place where much can happen. Is that where the notion of ‘hell’ being hot originated?

Enough of philosophical wanderings and let’s get back to our gas turbine engine! The arguments made above concerning the working of a gas turbine engine can be made much more precise using an expanded version of the conservation of energy statement we have used so far to include mechanical work (compressor and turbine) and heat (combustor). The books referenced in the bibliography should be helpful in the quantification of the performance of the engine. The analysis is rather simple in that the relations are algebraic and don’t involve much higher mathematics. The takeaway from such analysis is that the temperature extremes involving the maximum (turbine inlet) and minimum (compressor inlet) play a dominant role in the performance.

Finally, is this new engine any good? It is, apparently, a very successful engine to propel airplanes. As a postscript to our description of the gas turbine, it merits callout to some of the features that make it different from the ICE we know so well. The jet engine is a thrust, rather than a power, producing engine. Both engine types use liquid fuels, albeit slightly different kinds. For flight, using a liquid fuel is extremely important because of the ease of storing it onboard and handling it. Just contemplate for a moment the need for shoveling a solid fuel like coal on the ships of yesteryear or the steam locomotive. The difficulties involved in storing gaseous fuel under pressure in an airplane where payload volume is very dear have been considered and are not attractive. Aviation had to wait for the invention of the ICE, in part, for that reason.

There are important technical differences between the engines that say a lot about where and how they are used. The ICE works well and efficiently at power levels less than maximum. The gas turbine does not. The gas turbine works most efficiently at full thrust (or power) which is just what is needed for an airplane in cruise. The ICE works very well in automobiles where it hardly ever runs at full power. The demands of traffic insist on that.

In considering the turboshaft gas turbine for small applications, say under 500 hp, (about 400 kW) the boundary layers on the walls of the various flow components become a larger proportion of the total flow, meaning that friction is proportionally more important. The associated reduction in component efficiencies and the total pressure losses become important contributors to increased fuel consumption (per horsepower). The gas turbine is not easily well adapted in small sizes with high efficiency as a goal.

Finally, an important operational difference is that the turbomachinery, the compressor and turbine, has a large inertia invested in rotation. That will result in a slower response to power demand changes when compared to a similar requirement made of the ICE. The slower temporal response has consequences in the military setting or when emergencies are encountered.

12.4 How Do You Start This Thing?

When we examined the ideal engine consisting of perfect (thermodynamically reversible) compressor and turbine, we implied that all we had to do is spin it up to operating speed and then let it run on its own. That is indeed what has to be done to a real engine. Gas turbine engines typically have connections to a device that has enough power to raise the speed to the point where the engine operation is self-sustaining. The power for an electric motor may be provided by a battery, an auxiliary power unit (an APU), or from a power cart on the ground. Typically, commercial airliner engines are started by an air turbine supplied by an APU. Another source could be air from a storage source or, as it has been done, generated by burning a charge of solid rocket propellant-like material to generate the gas flow. There are interesting circumstances under which such means were or are used but they are quite uncommon. In flight, an engine may “flame out” or stop working and needs to be restarted. In some circumstances, the airplane’s speed and altitude may be exploited to have the flight airstream turn the engine over. It is hard to imagine that the circumstances that call for such action are benign.

12.5 Bleed Valves and Variable Stator Geometry

There is another aspect of starting a jet engine that is interesting. Consider the operation of an axial compressor. For a running situation, the blade height and flow passage widths are very much smaller at the outlet end of the compressor than at its inlet. The reason is that the increasing density associated increasing pressure as we look further down the compressor path is such that a smaller flow area is required at the high-pressure end to pass the mass flow being processed. During starting, that high density is not yet achieved but the mass flow forced in by the inlet must exit at the outlet. That implies that, initially, the flow velocities near the rear are very high, potentially high enough to “choke” the flow by reaching sonic speeds. If and when that happens the increasing flow demanded by the front of the compressor cannot be accommodated and the process is stuck; unless…

There are two aspects of the design of a modern gas turbine engine that are used to prevent this situation from being problematic. The first is to allow a bleed from the midsection of a compressor to remove partially compressed air from that location so that flow to the rear blockage is reduced. These bleed valves shut when operation reaches a normal operating state.

A second aspect that deals with starting is that in some compressor designs, the stators can be made to change the angles they present to the flow. Thus, they can be oriented to minimize blockage during startup. These two solutions can be employed in concert when they are designed to do so. The pilot sitting in the cockpit’s left seat is ignorant of all this because there is an engine control system that takes the necessary measurements and, with actuators of various kinds, does all that is necessary without burdening the pilot with its doings (Fig. 12.6).

Fig. 12.6

External view of a small gas turbine compressor with variable geometry stators. The five rings are attached to levers that change the orientation of the stators mounted on shafts through the case. The hydraulic lines and the associated cylinder are used to fix the stator orientation

Finally, the starting process is eased by the configuration of the engine when it has more than one shaft. Specifically, the starting process is initiated with the high-pressure components, the high-pressure compressor, burner and turbine, while the low-pressure components wait for things to happen on the inner portion of the engine. After all, the high-pressure section of a multi-spool gas turbine is an engine within an engine! When the high-pressure section is running in a sustained way, the low-pressure components follow suit and the entirety springs to life.