9.0 Introduction

The compressor, which raises the pressure of the air before combustion, and the turbine, which extracts work from the hot high-pressure combustion products, are at the very heart of the engine. Up to nowwe have assumed that it is possible to construct a suitable compressor and turbine without giving any attention to how this might be done. In this chapter an elementary treatment is given with the emphasis being to find the overall features of the compressors and turbines, including the number of stages, the suitable rotational speeds, the diameters and some indication of the flowpath. The details of blade shape will not be addressed. Further information is obtainable in a recent book by Dixon and Hall (2013) and at a more specialised level for compressors in Cumpsty (2004).

The description of turbomachinery is based on the fan, compressor and turbine for the engines of the NEA. To avoid unnecessary duplication, the design is restricted to the case with a fan pressure ratio of 1.5 at cruise. At this condition the stagnation pressure and temperature entering the engine are p02 = 35.6kPa and T02 = 245.4K. From Exercise 7.4 the mass flow through the engine to give the required thrust is 392 kg/s and of this 32.1 kg/s goes through the core (bpr = 11.2). The turbine inlet temperature is 1500 K, the overall pressure ratio is 45 and the HP compressor pressure ratio is 18. From Exercise 7.6 the fan diameter is 2.66 m.

8.0 Introduction

It would be possible to calculate the performance of an engine in the manner of Exercises 7.2 to 7.7 for every conceivable operating condition, e.g. for each altitude, forward speed and rotational speed of the components. This is not an attractive way of considering variations and it does not bring out the trends as clearly as it might. An alternative is to predict the variations by using the appropriate dynamic scaling – apart from its usefulness in the context of engine prediction, the application of what is conventionally called dimensional analysis is illuminating. The creation of groups which are actually non-dimensional is less important than obtaining groups with the correct quantities in them. The reasoning behind these ideas is discussed in Chapter 1 of Cumpsty (2004). For compatibility with the usual terminology, however, the phrases ‘dimensional analysis’ and ‘non-dimensional operating conditions’ will be retained here.

Using the ideas developed on the basis of dynamic scaling it is possible to estimate the engine performance at different altitudes and flight Mach numbers when the engine is operating at the same non-dimensional condition. From this it is possible to assess the consequences of losing thrust from an engine at cruise and, in some cases, to estimate how the engine will behave on a static test bed.

8.1 Engine variables and dependence

Figure 8.1 shows a schematic engine installed under a wing. For simplicity this is a mixed-flow engine, where the core and bypass streams are mixed before entering a single propulsive nozzle. Later the treatment can be generalized by considering the two streams separately. Alternatively, with bypass ratios to 10 or more, the errors may not be large if the analysis concentrates on the bypass stream and assumes the ratio of core jet velocity to bypass jet velocity stays approximately constant.

13.0 Introduction

This part of the book begins the consideration of the engine requirements of a new combat aircraft. In parallel with the treatment in earlier chapters for the engines of the New Efficient Aircraft, the approach chosen is to address the design of engines for a possible new aircraft so that the text and exercises can be numerically based with realistic values. The specifications for the New Fighter Aircraft used here have a marked similarity to those available for the new Eurofighter or Typhoon.

The topic of the present chapter is the nature of the combat missions and the type of aircraft involved. Figure 13.1 shows the different regions in which aircraft operate in terms of altitude versus Mach number, with the lines of constant inlet stagnation temperature overlaid. We are concerned here with what are referred to in the figure as fighters, a major class of combat aircraft. Figure 13.1 shows the various boundaries for normal operation. Even high-speed planes do not normally fly at more than M = 1.2 at sea level because in the high density air the structural loads on the aircraft and the physiological effects on the crew become too large. At high altitude high-speed aircraft do not normally exceed M ≈ 2.3, largely because the very high stagnation temperatures preclude the use of aluminium alloys without cooling. The boundary to the left in Figure 13.1 corresponds to the aircraft having insufficient speed to create the necessary lift; this is the stalling boundary and can be expressed in terms of the equivalent airspeed, that is the airspeed at sea level having the same value of 1/2ρV2.

11.0 Introduction

Up to this point consideration has been given only to the design point of the engine. This is clearly not adequate for a variety of reasons. Engines sometimes have to give less than their maximum thrust to make the aircraft controllable and to maintain an adequate life for the components. Furthermore all engines have to be started, and this requires the engine to accelerate from very low speeds achieved by the starter motor. The inlet temperature and pressure vary with altitude, climate, weather and forward speed and these need to be allowed for.

To be able to predict the off-design performance it is necessary to have some understanding of the way the various components behave and this forms the topic of the present chapter. It is fortunate that to understand off-design operation and to make reasonably accurate predictions of trends it is possible to approximate some aspects of component performance. The most useful of these approximations is that the turbines and the final propulsive nozzle are perceived by the flow upstream of them as choked. Another useful approximation is that turbine blades operate well over a wide range of incidence so that it is possible to assume a constant value of turbine efficiency independent of operating point. These approximations make it possible to consider the matching of a gas turbine jet engine and to see how the various components operate together at the conditions for which they are designed (the design point) and at off-design conditions. Off-design performancewill form the topic of Chapter 12.

5.0 Introduction

This chapter looks at the layout of some jet engines, using cross-sectional drawings. This begins with relatively simple engines and leads to engines for a recent large aircraft, the Boeing 787 and an engine for the smaller Bombardier C-series. Two concepts are introduced in the chapter. One is the multi-shaft engine with separate low-pressure and high-pressure spools. The other is the bypass engine in which some, very often most, of the air compressed by the fan bypasses the combustor and turbines.

Any consideration of practical engines must address the temperature limitations on the turbine. The chapter ends with some discussion of cooling technology and of the concept of cooling effectiveness.

5.1 The turbojet and the turbofan

Figure 5.1 shows a cut-away drawing of a Rolls-Royce Viper engine. This is typical of the simplest form of turbojet engine, which was the norm in the 1950s when it entered service, with an axial compressor coupled to an axial turbine, all on the same shaft. (The shaft, the compressor on one end and turbine on the other are sometimes referred to together as a spool.) Even for this very simple engine, which was originally designed to be expendable as a power source for target drones, the drawing is complicated. For more advanced engines such drawings become unhelpful at this small scale and simplified cross-sections are therefore more satisfactory and will be shown. A simplified cross-section is also shown for the Viper in Figure 5.1, as well as a cartoon showing the major components.

Preface to the third edition

The civil aircraft, the Airbus A380, which resembles the New Large Aircraft of Part 1 of earlier editions, has been in service for several years. Attention in the aircraft industry has now shifted to two-engine aircraft with a greater emphasis on reduction of fuel burn, so the model created for Part 1 is the New Efficient Aircraft, a twin aimed at high efficiency. There is another change to highlight, which is the switch to using fan pressure ratio as the independent design variable rather than bypass ratio. In the time since the first edition, the typical fan pressure ratios have been reduced, and this has necessarily led to a considerable increase in complexity. The changes relating to military combat engines are relatively small.

Another major change is the inclusion of a co-author. Andy Heyes had used the second edition in teaching a course in Imperial College and was well placed in terms of knowledge and experience to work on the third edition.

For the third edition, we would like to acknowledge additional help from friends and colleagues. From Rolls-Royce, we would mention Conrad Banks, John Bolger, Simon Gallimore, Peter Hopkins, Glen Knight, Paul Madden, Alan Newby, Ian Rainbow and Joe Walsh; from Pratt and Whitney, Yuan Dong, Alan Epstein and Jayant Sabnis; from Stanford University, Juan Alonso and Anil Variyar; from the University of Cambridge, Chez Hall and John Young; from Imperial College, Aaron Costall, Ricardo Martinez-Botas and Peter Newton; from Ohio State, Mike Dunn, retired from NASA, Tony Strazisar; retired from General Electric, Meyer Benzakein; and retired from Airbus, Jeff Jupp.

2.0 Introduction

The engine requirements for an aircraft depend upon the size, range and speed selected, but they also depend on the aerodynamic behaviour of the aircraft and the way in which it is operated. In this chapter some very elementary aspects of civil aircraft aerodynamic performance are described (if further explanation is needed the reader is referred to Anderson (2011)). These lead to a brief description of the conditions which are most critical for the engine: take-off, climb and cruise. It is possible to see why cruising fast and high is desirable, and to calculate the range. We can also develop the New Efficient Aircraft, the NEA, to form the basis of the exercises in Part 1 of the book. Knowing the ratio of lift to drag it is possible from this stipulation of the NEA to estimate the total thrust requirement.

2.1 Payload versus range

It is common to show aircraft performance by plotting maximum allowable payload against range capability, and an example corresponding to the Airbus A330-300 is shown in Figure 2.1. The presumption is that only enough fuel is carried to enable the mission, though reserves are also carried to enable diversions or other contingencies to be handled safely – it is the intention that these reserves will not be used. As the range is increased the weight of fuel at take-off must increase. The maximum payload allowed by the structure is marked by line OA and, as the range is increased along the line OA, the payload can remain constant at its maximum value, despite the increased weight of fuel, until the point A is reached. At point A the total weight of the aircraft has reached the maximum take-off weight and the range for this is denoted by R1.

18.0 Introduction

This chapter will look only briefly at the design of the compressors and turbines for combat engines, following on from Chapter 9 for the civil transport engine. The flow Mach numbers inside the compressors tend to be higher than for the subsonic transport and this makes the treatment of each blade row rather special; the design rules must take account of the presence of strong shock waves. A very important design consideration is how the compressor will behave at off-design conditions since combat engines are off-design for so much of their operation. The off-design problem arises because of the large density ratio between inlet and outlet of the compressor, and the reduction in this ratio when is decreased. The turbine stages do not suffer from this off-design problem. The turbines are required to produce large work output in relation to the blade speed, that is must be high, but at off-design conditions for the engine the turbine condition is essentially unaltered from the condition at design point. This, it may be recalled from Chapters 12 and 17, is because the turbines and the propulsive nozzle are effectively choked, so the turbines are forced to operate at the same non-dimensional condition.

In this chapter the consideration will be based on the case 1 engine of Chapter 17, with fan pressure ratio 4.5 and HP compressor ratio 6.66 at design point, sea-level static. This was calculated as Exercise 16.6 with the technology standard given in Table 16.1. The polytropic efficiency chosen for the fan was 0.85, for the HP compressor it was 0.90 and for the turbine 0.875.

4.0 Introduction

The gas turbine has many important applications but it is most widely used as the jet engine. Many of the gas turbines used in land-based and ship-based applications are derived directly from aircraft engines. The gas turbines designed specifically for land use, most often as part of a combined cycle plant, are rather different, but the concepts for the cycle is similar to the jet engine and some of the technology is related to that for aircraft propulsion.

The attraction of the gas turbine for aircraft propulsion is the large power output in relation to the engine weight and size – it was this which led the pre-Second World War pioneers to work on the gas turbine. Most of the pioneers then had in mind a gas turbine driving a propeller, but Whittle and later von Ohain realised that the exhaust from the turbine could be accelerated to form the propulsive jet. Sometimes the gas turbine is still used to drive a propeller to form an efficient engine for relatively low-speed flight, the turbo-prop. However the gas turbine is used for aircraft propulsion, it is the high power output for a given weight is that makes it attractive.

Purists will object to this description of the gas turbine as a cycle. Strictly speaking a cycle uses a fixed parcel of fluid which in a gas turbine would be compressed, heated in a heat exchanger, expanded in a turbine and then cooled in a heat exchanger. The ideal gas turbine is sometimes called a Joule or Brayton cycle. The gas turbine ‘cycle’ we consider here takes in fresh air, burns fuel in it and then discharges it after the turbine: in other words it does not cycle the air. Here we are nevertheless adopting the standard terminology of the industry.

20.0 Introduction

This chapter returns to some general issues related to both civil and military engines. These are topics which can be more satisfactorily addressed with the background of earlier chapters.

20.1 Civil aviation and the environment

The choice of aircraft here, the New Efficient Aircraft, was predicated on the need to reduce fuel burn. The cruise Mach number was lower than current new large aircraft and the full payload maximum range, R1, was distinctly smaller than recent designs of aircraft offered for sale. The small reduction in speed would have little effect on the passenger and the vast majority of flights currently offered would be accommodated with the reduced range. The industry as a whole does not seem willing to embrace this approach to reducing fuel burn, even though fuel is now a major cost, perhaps 40% of the direct costs. No aircraft flying in 2015 has been designed with fuel cost approaching the values currently prevailing and expected in the future. Despite the relatively high cruise Mach numbers and long R1 ranges being offered, most people nevertheless agree that global warming mitigation demands a reduction in CO2 emissions.

ICAO have come forward with a metric for fuel burn but, as described in Chapter 1, this is not well conceived if the object is to display the relative efficiency of different aircraft and aircraft types in moving people or freight. Now ICAO has to decide how to turn this metric into regulations for fuel burn. The earliest these regulations could begin to affect new aircraft types is probably 2020 with existing types several years later. It seems unlikely, however, that this will matter very much, for there is already major economic pressure to reduce mission fuel burn.

16.0 Introduction

In this chapter we will consider three separate engine designs corresponding to distinct operating conditions. For convenience here the three design points are at the tropopause (altitude 11 km; standard atmosphere temperature 216.65 K and pressure 22.7 kPa) for Mach numbers of 0.9, 1.5 and 2.0. The thrusts required for these conditions were determined in Exercise 14.4a. At each condition a separate engine is designed – this is quite different from designing the engine for a single condition and then considering the same engine operation at different conditions, referred to off-design, which is the topic of Chapter 17.

For this exercise all design points will correspond to the engine being required to produce maximum thrust, even though the ultimate suitability of an engine for its mission may depend on performance, particularly fuel consumption, at conditions for which the thrust is very much less than maximum. The designs will first be for engines without an afterburner (operation ‘dry’) and then with an afterburner; the afterburner will be assumed to raise the temperature of the exhaust without altering the operating condition of the remainder of the engine so the stagnation pressure entering the propulsive nozzle is unchanged.

The engines considered will all be of the mixed turbofan type – such an engine was shown in Figure 15.1 with a sketch showing the station numbering system adopted. Note that the numbering shows station 13 downstream of the fan in the bypass and station 23 downstream of the fan for the core flow; in the present simplified treatment it will be assumed that p023 = p013 and that T023 = T013.