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.

The present chapter will consider only the major components: the fan and compressor, the combustor, the turbine and the propulsive nozzle. Because of its simplicity it is convenient to begin by considering the nozzle, but prior to this the issue of gas properties will be addressed.

Gas properties in the aircraft gas turbine

In the treatment of the engines for the New Efficient Aircraft in Chapters 1–10 the specific heat capacity of the gas at constant pressure cp and ratio of specific heat capacities γ = cp∕cv were assumed to be equal for the air and for the products of combustion and to be constant regardless of temperature and pressure. This is a major over-simplification which will be corrected somewhat in the present and later chapters.

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. There are small losses associated with mixing of core exhaust and bypass stream, but these will be neglected here. As a result, if the fan pressure ratio is fixed then so too is the pressure at outlet from the turbine and the conditions through the core are also determined. Fixing the pressure ratio across the fan is a more direct way of specifying properties inside the engine than, for example, the bypass ratio; the pressure ratio is also the dominant term in the expression for jet velocity and therefore for gross and net thrust.

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.

This chapter looks at the operation of simple gas turbines and outlines the method of calculating the power output and efficiency. The treatment is simplified by treating the working fluid as a perfect gas with the properties of air, but later some examples are discussed to assess the effect of adopting more realistic assumptions.

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.

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.

More recent turbojet engines had two spools so that the compression and expansion were split into parts. For flight at sustained speeds well in excess of the speed of sound a turbojet engine remains an attractive option and a two-shaft example, the Rolls-Royce Olympus 593, is shown in Figure 5.2. Four of these engines were used to propel the Concorde at around twice the speed of sound. The low-pressure (LP) compressor and LP turbine are mounted on one shaft to form the LP spool. The LP shaft passes through the high-pressure (HP) shaft on which are mounted the HP compressor and the HP turbine. The compression process is split between two spools for reasons to do with operation at speeds below the design speed, including starting; this is discussed in some detail in Chapter 12.

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.

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. The situation with fuel and CO2 is quite unlike other areas in which ICAO is involved as setter of regulations, such as noise (discussed in the Appendix) and pollutants (like NOx, discussed in Chapter 11). Without regulations of some sort there would be no incentive for an airline to reduce these noise and NOx because there is no direct cost.

Introduction

When the engine for a new civil transport, the New Large Aircraft, was considered in Chapters 1 to 10 many assumptions were introduced to make the treatment as simple as possible. In the treatment of the engine for a New Fighter Aircraft in Chapters 13–18 the level of complexity was increased. The properties of the gas were allowed to be different before and after burning of the fuel and the effect of the mass flow of fuel added to the gas passing through the turbine was included. The effect of the cooling air supplied to the turbines was allowed for and the effect of the pressure loss in the combustor was accounted. It is appropriate to recalculate the performance of an engine for the civil aircraft with some of these effects included and that level of fidelity will apply to most of this chapter.

Another difference between the treatment for the civil engine in Chapters 1–10 and the treatment for the combat aircraft was the mixing of the core and bypass streams upstream of the final propulsive nozzle in the combat engine. Some large civil engines are mixed and this chapter therefore opens with a brief consideration of this option. Following this the consequences of different levels of fidelity in modelling will be addressed. A significant part of the chapter uses the most accurate model to look at the impact of cooling air, pressure drop in the combustor and component efficiency on the thrust and sfc of engines; this is done first for the engine on-design and then off-design. The chapter concludes with a brief consideration of propulsion for high-speed civil aircraft.

The benefit of mixing in large civil engines

It has been quite common for large civil engines to have a mixed exhaust, where the core and bypass streams are mixed before the final propulsive nozzle. Indeed, most of the Rolls-Royce RB211 engines were mixed and mixing is still common for smaller engines.

History and regulation

As jet air transport increased in the 1960s the annoyance to people living and working around major airports was becoming intense. Regulations affecting international air transport are governed by the International Civil Aviation Organisation (ICAO), but this body was moving so slowly that in 1969 the US Federal Aviation Agency (FAA) made proposals for maximum permitted noise levels. After extensive discussions in the USA these were formally approved as Federal Aviation Regulation (FAR) Part 36 in 1971, retroactive with effect from 1969, but only for new aircraft. Shortly afterwards the ICAO Committee on Aircraft Noise published similar recommendations, to be known as Annex 16, a formal addendum to the 1944 Chicago Convention on Civil Aviation; each member state had then to accept the rules in Annex 16 and write them into their legal framework. The underlying principle for the noise certification of aircraft under FAR Part 36 and Annex 16 are similar and has remained unchanged ever since, with the levels under the US and ICAO rules subsequently becoming virtually identical.

The certification for noise relies on measurements at three positions, two for take-off (referred to as lateral and flyover) and one for landing (referred to as approach). The levels are expressed in decibels (EPNdB) using effective perceived noise level (EPNL), described in outline below. The layout for testing is shown in Figure A1.

The noise at the lateral position is the highest noise measured along a line parallel to the runway whilst the aircraft is departing at full power and the maximum usually occurs when the aircraft has climbed to about 1000 feet. Flyover noise is measured directly under the flight path after take-off and at an altitude where it is normal to cut-back the power to reduce the noise whilst still maintaining a safe rate of climb. The approach noise is also measured directly under the flight path as the aircraft prepares to land, with the glide slope carefully controlled.