Lefebvre A.H., Ballal D.R. Gas Turbine Combustion: Alternative Fuels and Emissions

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Basic Considerations

1.1 Introduction

The primary purpose of this introductory chapter is to discuss the main

requirements of gas turbine combustors and to describe, in general terms,

the various types and congurations of combustors employed in aircraft and

industrial engines. The principal geometric and aerodynamic features that

are common to most types of gas turbine combustors are briey reviewed,

with special attention being given to fuel preparation and liner-wall cooling

to reect the important role these topics continue to play in combustor devel-

opment. Reference is made to most of the key issues involved in combustor

design and development, but the descriptive material is necessarily brief

because these and other important aspects of combustor performance are

described more fully in subsequent chapters.

Bearing in mind the pressures and exigencies of wartime Britain and

Germany, and the lack of knowledge and experience available to the designer,

it is perhaps hardly surprising that the rst generation of gas turbine com-

bustors were characterized by wide variations in size, geometry, and the

mode of fuel injection. With the passage of time and the post-war lifting of

information exchange, some commonalities in design philosophy began to

emerge. By around 1950, most of the basic features of conventional gas tur-

bine combustors, as we know them today, were rmly established.

Since that time, combustor technology has developed gradually and con-

tinuously, rather than through dramatic change, which is why most of the

aero-engine combustors now in service tend to resemble each other in size,

shape, and general appearance. This close family resemblance stems from

the fact that the basic geometry of a combustor is dictated largely by the need

for its length and frontal area to remain within the limits set by other engine

components, by the necessity for a diffuser to minimize pressure loss, and

by the requirement of a liner (ame tube) to provide stable operation over a

wide range of air/fuel ratios. During the past half century, combustion pres-

sures have risen from 5 to 50 atmospheres, inlet air temperatures from 450 to

900 K, and outlet temperatures from 1100 to 1850 K. Despite the continually

increasing severity of operating conditions, which are greatly exacerbated

2 Gas Turbine Combustion: Alternative Fuels and Emissions, Third Edition

by the concomitant increases in compressor outlet velocity, today’s combus-

tors exhibit close to 100% combustion efciency over their normal operating

range, including idling, and demonstrate substantial reductions in pollutant

emissions. Furthermore, the life expectancy of aero-engine liners has risen

from just a few hundred hours to many tens of thousands of hours.

Although many formidable problems have been overcome, the challenge of

ingenuity in design still remains. New concepts and technology are needed to

further reduce pollutant emissions and to respond to the growing requirement

of many industrial engines for multifuel capability. Gas turbines are “omniv-

orous” machines, capable of operating efciently on a wide variety of cheap

fuels, solid, liquid, and gaseous, with the exception of aircraft engines. Today,

the ever-rising cost of petroleum fuel is prompting research into developing

alternative liquid fuels and this is posing new combustor design challenges.

Another problem of increasing importance is that of acoustic resonance, which

occurs when combustion instabilities become coupled with the acoustics of the

combustor. This problem could be crucial to the future development of lean

premixed combustors.

It is clearly important that combustor developments should keep pace with

improvements in other key engine components. Thus, reduction of combus-

tor size and weight will remain an important requirement for aero engines,

whereas the continuing trend toward higher turbine inlet temperatures will

call for a closer adherence to the design temperature prole at the turbine

inlet. Simultaneously, the demand for greater reliability, increased durability,

and lower manufacturing, development, and maintenance costs seems likely

to assume added importance in the future. To meet these challenges, the

search goes on for new materials and new methods of fabrication to simplify

basic combustor design and reduce cost. The search has already led to the

development of advanced wall-cooling techniques and the widespread use

of thermal barrier refractory coatings within the combustion liner.

1.2 Early Combustor Developments

The material contained in this book is largely a chronicle of developments

in gas turbine combustion during the last half century. For both British

and German engineers, the development of a workable combustor was an

obstacle that had to be overcome in their independent and concurrent efforts

to achieve a practical turbojet engine. It proved to be a formidable task for

both groups and, in Whittle’s case, combustion problems dominated the rst

three years of engine development. The following abridged account of the

early history of gas turbine combustion in Britain, Germany, and the United

States is intended to cover the period from the start of World War II until

around 1950, by which time it was generally accepted that the piston engine

Basic Considerations 3

had reached its limit as a propulsion system for high-speed ight and the

gas turbine was rmly established as the powerplant of choice for aircraft

applications.

1.2.1 Britain

One method of preparing a liquid fuel for combustion is to heat it above

the boiling point of its heaviest hydrocarbon ingredient, so that it is entirely

converted into vapor before combustion. This was the method adopted by

Whittle for his rst turbojet engine. This engine employed 10 separate tubular

combustors in a reverse-ow arrangement to permit a short engine shaft.

Whittle tried several vaporizer tube congurations, more than 30 in all, one

of which is illustrated in Figure 1.1. This gure shows that fuel was heated in

tubes located in the ame zone. The fuel was maintained at high pressure so

that vaporization could not occur until it had been injected through a nozzle

and its pressure reduced to that of the combustion zone. Whittle experienced

considerable difculties with this system, due mainly to problems of thermal

cracking and coking up of the vaporizer tubes, as well as difculties in con-

trolling the fuel ow rate.

After many trials and setbacks, Whittle adopted a combustor whose main

attraction was the replacement of vaporizer tubes by a pressure-swirl atom-

izer having a wide spray cone angle. Another interesting feature of the new

combustor was that most of the primary-zone airow entered the combus-

tion zone through a large air swirler located at the upstream end of the liner

around the fuel nozzle, as shown in Figure 1.2. This swirler served to create

a toroidal ow reversal that entrained and recirculated a portion of the hot

combustion products to mix with the incoming air and fuel. This arrange-

ment not only anchored the ame, but also provided the rapid mixing of

fuel vapor, air, and combustion products needed to achieve high heat-release

rates. The additional air required to complete combustion and reduce the gas

temperature to a value acceptable to the turbine was supplied through stub

Efflux gases

Pilot starting jet

Vaporizing tubes

Air entry

Figure 1.1

Early Whittle vaporizer combustor.

4 Gas Turbine Combustion: Alternative Fuels and Emissions, Third Edition

pipes that projected radially inward and through holes pierced in the liner

walls. After suitable development, this combustor was adopted for the Power

Jets W1 engine, which made the rst British turbojet-powered ight on the

evening of May 15, 1941.

Another early British engine was the De Havilland Goblin, which was the

rst engine to power the Lockheed P-40. (It was later replaced by General

Electric’s I-40 engine, which provided 33% more thrust.) The Goblin is of his-

torical interest because it was the rst British engine to use “straight-through”

combustors, as opposed to the “reverse-ow” type employed on all previous

engines. The rst British annular combustor appeared on the Metropolitan

Vickers Beryl engine. A noteworthy feature of this combustor was the use of

upstream fuel injection. This system was also used in other engines, the ear-

liest example being the German Jumo 004. The main advantage claimed for

upstream fuel injection was a longer residence time of the fuel droplets in the

combustion zone, which provided more time for fuel evaporation. Its main

drawback stemmed from the immersion of key components in the ame.

Cooling arrangements for the atomizer feed arm could be provided, but it

was difcult to eliminate entirely the problem of carbon deposition on the

atomizer face. For this reason, upstream fuel injection is no longer regarded

as a practical option.

Another interesting feature of the Metrovick combustor is the manner in

which dilution air was introduced into the combustion gases downstream

of the primary combustion zone. This corresponds closely to the method

employed in the Jumo 004. Figure 1.3 shows two rows of narrow scoops that

interleave cold airstreams between co-owing streams of hot combustion

products. The rst row of scoops provided air for the completion of combus-

tion, with any excess serving as dilution air. The air owing through the sec-

ond row of scoops was solely for dilution purposes. This type of “sandwich”

mixer has useful advantages in terms of low pressure loss and low pattern

factor. However, it carries a high weight penalty, which is clearly a serious

drawback for aircraft engines, and the scoops are also prone to burnout

Stub pipes

Swirl vanes

Atomizer

Efflux gases

Air inlet

Figure 1.2

Early Whittle atomizer combustor.

Basic Considerations 5

because of their exposure to the high-velocity combustion gases. Sandwich

mixers are no longer used, except in a highly abbreviated form where their

main function is to strengthen the liner and raise the ow discharge coef-

cient (typically to around 0.8, as opposed to around 0.6 for a plain hole).

1.2.2 germany

1.2.2.1 Jumo 004

This engine is of great historical interest because it was the world’s rst mass-

produced turbojet and one that saw extensive service in World War II. It was

among the rst engines to employ axial ow turbomachinery and straight-

through combustors. Each of the six tubular combustors was supplied with

fuel at pressures up to 5.2 MPa (750 psi) from a pressure-swirl atomizer, which

sprayed the fuel upstream into the primary combustion zone. Figure 1.4

shows schematically the basic combustor design. The primary air owed into

the liner through six swirl vanes, the amount of air being sufcient to achieve

near-stoichiometric combustion at the engine design point. Mixing between

combustion products and dilution air was achieved using an assembly of

stub pipes that were welded to a ring at their upstream end and to the outer

perimeter of a 10-cm diameter dished bafe at their downstream end. The hot

Sandwich

scoops

Efflux

gases

Fuel

Air

Figure 1.3

Metrovick annular combustor.

Stub pipes

Efflux gases

Upstream fuel injection

Air inlet

Swirl vanes

Figure 1.4

Jumo 004 tubular combustor.

6 Gas Turbine Combustion: Alternative Fuels and Emissions, Third Edition

combustion products owed radially outward through the gaps between the

stub pipes to meet and mix with part of the cold secondary air. The remaining

secondary air owed through the stub pipes, incidentally serving to protect

them from burnout because of their immersion in the hot combustion gases,

to provide further mixing of hot and cold gases in the recirculation zone

created by the presence of the bafe.

1.2.2.2 BMW 003

The only other German turbojet engine to be developed to the produc-

tion stage during World War II was the BMW 003. This engine employed

an annular combustor tted with 16 equispaced, downstream-spraying,

pressure atomizers. Each fuel nozzle was surrounded by a bafe and the

primary combustion air owed both through and around it. The method

used to inject the dilution air was a sandwich mixer arrangement, which

interleaved streams of cold secondary air between parallel streams of hot

combustion products. Dilution air owed through 40 scoops attached to

the outer liner, alternating in circumferential locations with 40 similar

scoops attached to the inner liner. The end result, as shown in Figure 1.5,

was a combustor having a relatively low pressure loss, but also a fairly

high length/height ratio.

1.2.3 The united States

During the development of the W1 engine, the decision was made to build a

larger Whittle engine of 1600 lb thrust to be designated as the W2B. In 1941, a

W2B engine, complete with drawings, was delivered to the General Electric

Company (GE); within six months this company had built two more engines

with the same design. In 1947, Pratt and Whitney (P&W), having been fully

preoccupied with piston engine production throughout the war, made its

rst entry into the turbojet arena by licensing the Nene engine from Rolls

Royce. Having established a foothold in the turbojet engine business, GE and

P&W lost no time in producing their own independent combustor designs.

For example, GE’s Whittle-derived J31 engine employed a reverse-ow

Sandwhich scoops

Efflux gases

Inlet air

Figure 1.5

BMW 003 annular combustor.

Basic Considerations 7

combustor, but a straight-through version (see Figure 1.6) was adopted for

the J33 and for subsequent engines such as the J35 and J47. For its J57 engine,

shown in Figure 1.7, P&W employed eight tubular liners located within an

annular casing. Each liner had a perforated tube along its central axis that

extended about halfway down the liner. In effect, this central tube converted

the tubular liner into a small annular combustor, supplied with fuel from six

equispaced, pressure-swirl nozzles.

By 1943, Westinghouse had developed successful axial-ow turbojet

engines without any European input. An annular combustor was selected

for its J30 engine, whereas a dual-annular conguration was adopted for the

J34. This dual-annular concept was ahead of its time, and interest in it lapsed

until it was resurrected by GE in the 1970s to serve as a low-emissions com-

bustor for their CFM56-B engine.

By the end of the 1940s, the development work carried out in the UK,

Germany, and the United States had established the basic design features of

aero-engine combustors that have remained largely unchanged. The main

components are a diffuser for reducing the compressor outlet air velocity

to avoid high-pressure losses in combustion, a liner (or ame tube) that is

arranged to be concentric within the outer combustor casing, means for sup-

plying the combustion zone with atomized or vaporized fuel and, with tubu-

lar liners, interconnectors (or cross-re tubes) through which hot gases can

ow from a lighted liner to an adjacent unlighted liner.

Within the liner itself, the distribution of air is arranged to ensure that the pri-

mary combustion zone operates at a much higher fuel/air ratio than the overall

Graduated air admission

Efflux gases

Inlet air

Figure 1.6

General Electric J33 tubular combustor.

Six fuel nozzles

Central air tube

Efflux gases

Inlet air

Figure 1.7

Pratt & Whitney J57 tuboannular combustor.

8 Gas Turbine Combustion: Alternative Fuels and Emissions, Third Edition

combustor fuel/air ratio. More air is admitted downstream of the primary zone

to complete the combustion process and to dilute the combustion products to a

temperature acceptable to the turbine.

1.3 Basic Design Features

It is of interest to examine briey the considerations that dictate the basic

geometry of the “conventional” gas turbine combustor. It is also instructive

because it helps to dene the essential components needed to carry out the

primary functions of a combustion chamber.

Figure 1.8a shows the simplest possible form of combustor—a straight-

walled duct connecting the compressor to the turbine. Unfortunately, this

simple arrangement is impractical because the pressure loss incurred would

be excessive. The fundamental pressure loss because of combustion is pro-

portional to the square of the air velocity and, for compressor outlet veloci-

ties of the order of 170 m/s, this loss could amount to almost one-third of

the pressure rise achieved in the compressor. To reduce this pressure loss to

an acceptable level, a diffuser is used to lower the air velocity by a factor of

about 5, as shown in Figure 1.8b. Having tted a diffuser, a ow reversal

must then be created to provide a low-velocity region in which to anchor the

ame. Figure 1.8c shows how this may be accomplished with a plain bafe.

The only remaining defect in this arrangement is that to produce the desired

temperature rise, the overall chamber air/fuel ratio must normally be around

30–40, which is well outside the limits of ammability for hydrocarbon–air

mixtures. Ideally, the air/fuel ratio in the primary combustion zone should

Fuel

Air

(d)

Fuel

(c)

Air

Fuel(b)

Air

Fuel(a)

Air

Figure 1.8

Derivation of conventional combustor conguration.

Basic Considerations 9

be around 18, although higher values (around 24) are sometimes preferred if

low emissions of nitric oxides (NO

) is a prime consideration. To deal with

this problem, combustion is sustained by a recirculatory ow of burned

products that provide a continuous source of ignition for the incoming fuel–

air mixture. The air not required for combustion is admitted downstream of

the combustion zone to mix with the hot burned products, thereby reducing

their temperature to a value that is acceptable to the turbine.

Figure 1.8 thus illustrates the logical development of the conventional gas

turbine combustion chamber in its most widely used form. As would be

expected, there are many variations on the basic pattern, shown in Figure 1.8d,

but, in general, all chambers incorporate an air casing, diffuser, liner, and fuel

injector as key components.

The choice of a particular type and layout of combustion chamber is deter-

mined largely by engine specications, but it is also strongly inuenced by

the desirability of using the available space as effectively as possible. On large

aircraft engines, the chamber is almost invariably of the straight-through

type, in which the air ows in a direction essentially parallel to the axis of the

chamber. For smaller engines, the reverse-ow annular combustor provides

a more compact unit and allows close coupling between the compressor and

turbine. In most combustors, the fuel is injected into the burning zone in the

form of a well-atomized spray, obtained either by forcing it through a ne

orice under pressure, or by utilizing the pressure differential across the liner

wall to create a stream of high-velocity air that shatters the fuel into ne drop-

lets before transporting it into the primary combustion zone.

1.4 Combustor Requirements

A gas turbine combustor must satisfy a wide range of requirements whose

relative importance varies among engine types. However, the basic require-

ments of all combustors may be listed as follows:

1. High-combustion efciency (i.e., the fuel should be completely

burned so that all its chemical energy is liberated as heat)

2. Reliable and smooth ignition, both on the ground (especially at very

low ambient temperatures) and, in the case of aircraft engines, after

a ameout at high altitude

3. Wide stability limits (i.e., the ame should stay alight over wide

ranges of pressure and air/fuel ratio)

4. Low pressure loss

5. An outlet temperature distribution (pattern factor) that is tailored to

maximize the lives of the turbine blades and nozzle guide vanes