30079_57a.pdf

CHAPTER 57

GAS TURBINES

Harold Miller

GE Power Systems

Schenectady, New York

57.1

INTRODUCTION

1723

57.3 APPLICATIONS

1749

57.1.1 Basic Operating

Principles

57.3.1 Use of Exhaust Heat in

Industrial Gas Turbines

1723

1749

57. 1.2 A Brief History of Gas

Turbine Development

and Use

57.3.2 Integrated Gasification

Combined Cycle

1751

1727

57.3.3 Applications in Electricity

Generation

57.1.3 Components,

Characteristics and

Capabilities

1753

57.3.4 Engines for Aircraft

1755

1728

57.3.5 Engines for Surface

Transportation

57.1.4 Controls and Accessories

1737

1757

57.1.5 Gas Turbine Operation

1740

57.4 EVALUATIONAND

SELECTION

57.2 GAS TURBINE

PERFORMANCE

1759

1740

57.4.1 Maintenance Intervals,

Availability, and

Reliability

57.2.1 Gas Turbine

Configurations and

Cycle Characteristics

1759

1740

57.4.2 Selection of Engine and

System

57.2.2 Trends in Gas Turbine

Design and Performance

1761

1747

57.1 INTRODUCTION

57.1.1 Basic Operating Principles

Gas turbines are heat engines based on the Brayton thermodynamic cycle. This cycle is one of the

four that account for most of the heat engines in use. Other cycles are the Otto, Diesel and Rankine.

The Otto and Diesel cycles are cyclic in regard to energy content. Steady-flow, continuous energy

transfer cycles are the Brayton (gas turbine) and Rankine (steam turbine) cycles. The Rankine cycle

involves condensing and boiling of the working fluid, steam, and utilizes a boiler to transfer heat to

the working fluid. The working fluid in the other cycles is generally air, or air plus combustion

products. The Otto, Diesel and Brayton cycles are usually internal combustion cycles wherein the

fuel is burned in the working fluid. In summary, the Brayton cycle is differentiated from the Otto

and Diesel cycle in that it is continuous, and from the Rankine in that it relies on internal combustion,

and does not involve a phase change in the working fluid.

In all cycles, the working fluid experiences induction, compression, heating, expansion, and ex-

haust. In a non-steady cycle, these processes are performed in sequence in the same closed space,

This chapter was written as an update to chapter 72 of the Handbook's previous edition. Much of

the structure and significant portions of the text of the previous edition's chapter is retained. The new

edition has increased emphasis on the most significant current and future projected gas turbine con-

figurations and applications. Thermodynamic cycle variations are presented here in a consistent for-

mat, and the description of current cycles replaces the discussions of some interesting and historical,

but less significant, cycles described in the earlier edition. In addition, there is a new discussion of

economic and regulatory trends, of supporting technologies, and their interconnection with gas turbine

development. The author of the previous version had captured the history of the gas turbine's de-

velopment, and this history is repeated and supplemented here.

Mechanical Engineers' Handbook, 2nd ed., Edited by Myer Kutz.

formed by a piston and cylinder that operate on the working fluid one mass at a time. In contrast,

the working fluid flows through a steam turbine power plant or gas turbine engine, without interrup-

tion, passing continuously from one single-purpose device to the next.

Gas turbines are used to power aircraft and land vehicles, to drive generators (alternators) to

produce electric power, and to drive other devices, such as pumps and compressors. Gas turbines in

production range in output from below 50 kW to over 200 MW. Design philosophies and engine

configurations vary significantly across the industry. Aircraft engines are optimized for high power-

to-weight ratios, while heavy-duty, industrial, and utility gas turbines are heavier, being designed for

low cost and long life in severe environments.

The arrangement of a simple gas turbine engine is shown in Fig. 57.1a. The rotating compressor

acts to raise the pressure of the working fluid and force it into the combustor. The turbine rotation

is caused by the work produced by the fluid while expanding from the high pressure at the combustor

discharge to ambient air pressure at the turbine exhaust. The resulting mechanical work drives the

mechanically connected compressor and output load device.

The nomenclature of the gas turbine is not standardized. In this chapter, the term blading refers

to all rotating and stationary airfoils in the gas path. Turbine (expander) section rotating blades are

buckets, a term derived from steam turbine practice. Turbine section stationary blades are nozzles.

The combustion components in contact with the working fluid are called combustors; major com-

bustor components are fuel nozzles and combustion liners. Some combustors (Can-annular and silo-

types) have transition pieces that conduct hot gas from the combustion liners to the first-stage nozzles.

A stage of the compressor consists of a row of rotor blades, all at one axial position in the gas

turbine, and the stationary blade row downstream of it. A turbine stage consists of a set of nozzles

occupying one axial location and the set of buckets immediately downstream. Rotating blading is

attached either to a monolithic rotor structure or to individual discs or wheels designed to support

the blading against centrifugal force and the aerodynamic loads of the working fluid. The terms discs

and wheels are used interchangeably.

Gas turbine performance is established by three basic parameters: mass flow, pressure ratio, and

firing temperature. Compressor, combustor, and turbine efficiency have significant, but secondary,

effects on performance, as do inlet and exhaust systems, turbine gas path and rotor cooling, and heat

loss through turbine and combustor casings.

In gas turbine catalogues and other descriptive literature, mass flow is usually quoted as com-

pressor inlet flow, although turbine exit flow is sometimes quoted. Output is proportional to mass

flow.

Pressure ratio is quoted as the compressor pressure ratio. Aircraft engine practice is to define the

ratio as the total pressure at the exit of the compressor blading divided by the total pressure at the

inlet of the compressor blading. Industrial/utility turbine manufacturers generally refer to the static

pressure in the plenum downstream of the compressor discharge diffuser (upstream of the combustor)

divided by the total pressure downstream of the inlet filter and upstream of the inlet of the gas

turbine. Similarly, there are various possibilities for defining turbine pressure ratio. All definitions

yield values within 1 or 2% of one another. Pressure ratio is the primary determinant of simple cycle

gas turbine efficiency. High pressure results in high simple cycle efficiency.

Firing temperature is defined differently by each manufacturer, and the differences are significant.

Heavy-duty gas turbine manufacturers use three definitions. There is an ISO definition of firing

temperature, which is a calculated temperature. The compressor discharge temperature is increased

by a calculated enthalpy rise based on the compressor inlet air flow and the fuel flow. This definition

is valuable in that it can be used to compare gas turbines or monitor changes in performance through

calculations made on the basis of field measurements. To determine ISO firing temperature, one does

not require knowledge of the secondary flows within the gas turbine. A widely used definition of

Fig. 57.1 Simple engine type: (a) open cycle; (b) closed cycle (diagrammatic). 1

firing temperature is the average total temperature in the exit plane of the first stage nozzle. This

definition is used by General Electric for its industrial engines. Westinghouse refers to "turbine inlet

temperature," the temperature of the gas entering the first stage nozzle. Turbine inlet temperature is

approximately 10O 0 C above nozzle exit firing temperature, which is in turn approximately 10O 0 C

above ISO firing temperature. Since firing temperature is commonly used to compare the technology

level of competing gas turbines, it is important to compare on one definition of this parameter.

Aircraft engines and aircraft-derivative industrial gas turbines have other definitions. One nomen-

clature establishes numerical stations—here, station 3.9 is combustor exit and station 4.0 is first-stage

nozzle exit. Thus, T 3 9 is very close to "turbine inlet temperature" and T 4 0 is approximately equal

to GE's "firing temperature." There are some subtle differences relating to the treatment of the

leakage flows near the first-stage nozzle. This nomenclature is based on SAE ARP 755A, a recom-

mended practice for turbine engine notation.

Firing temperature is a primary determiner of power density (specific work) and combined cycle

(Brayton-Rankine) efficiency. High firing temperature increases the power produced by a gas turbine

of a given physical size and mass flow. The pursuit of higher firing temperatures by all manufacturers

of large, heavy-duty gas turbines used for electrical power generation is driven by the economics of

high combined cycle efficiency.

Pressures and temperatures used in the following descriptions of gas turbine performance will be

total pressures and temperatures. Absolute, stagnation, or total values are those measured by instru-

ments that face into the approaching flow to give an indication of the energy in the fluid at any point.

The work done in compression or expansion is proportional to the change of stagnation temperature

in the working fluid, in the form of heating during a compression process or cooling during an

expansion process. The temperature ratio, between the temperatures before and after the process, is

related to the pressure ratio across the process by the expression T b IT a = (P b /P a } (y ~ l)/y , where y is

the ratio of working fluid specific heats at constant pressure and volume. The temperature and pressure

are stagnation values. It is the interaction between the temperature change and ratio, at different

starting temperature levels, that permits the engine to generate a useful work output.

This relationship between temperature and pressure can be demonstrated by a simple numerical

example using the Kelvin scale for temperature. For a starting temperature of 30O 0 K (27 0 C), a tem-

perature ratio of 1.5 yields a final temperature of 45O 0 K and a change of 15O 0 C. Starting instead at

40O 0 K, the same ratio would yield a change of 20O 0 C and a final temperature of 60O 0 K. The equivalent

pressure ratio would ideally be 4.13, as calculated from solving the preceding equation for P b IP a \

P b /P a = T b /Tl /y ~ l = 1.5 1 - 4 ' 0 - 4 = 4.13. These numbers show that, working over the same temperature

ratio, the temperature change and, therefore, the work involved in the process vary in proportion to

the starting temperature level. 2

This conclusion can be depicted graphically. If the temperature changes are drawn as vertical

lines a-b and c-d, and are separated horizontally to avoid overlap, the resultant is Fig. 57.2a. As-

suming the starting and finishing pressures to be the same for the two processes, the thin lines through

a-d and b-c depict two of a family of lines of constant pressure, which diverge as shown. In this

ideal case, expansion processes could be represented by the same diagram, simply by proceeding

down the lines b-a and c-d. Alternatively, if a-b is taken as a compression process, b-c as heat

addition, c-d as an expansion process, and d-a as a heat rejection process, then the figure a-b-c-d-a

represents the ideal cycle to which the working fluid of the engine is subjected.

Over the small temperature range of this example, the assumption of constant gas properties is

justified. In practice, the 327 0 C (60O 0 K) level at point d is too low a temperature from which to start

Fig. 57.2 Temperature changes and temperature-entropy diagram for ideal

simple gas turbine cycles.

the expansion. Figure 57. 2b is more realistic. Here, the lines of constant pressure have been con-

structed for ideal gas-air properties that are dependent upon temperature. Expansion begins from a

temperature of 125O 0 C. With a pressure ratio of 16:1, the end point of the expansion is approximately

48O 0 C. Now a-b represents the work input required by the compressor. Of the expansion work

capacity c-d, only the fraction c-d' is required to drive the compressor. An optical illusion makes

it appear otherwise, but line a-d' is displaced vertically from line b-c by the same distance every-

where. The remaining 435 0 C, line d'-d, is energy that can be used to perform useful external work,

by further expansion through the turbine or by blowing through a nozzle to provide jet thrust.

Now consider line b-c. The length of its vertical projection is proportional to the heat added. The

ability of the engine to generate a useful output arises from its use of the energy in the input fuel

flow, but not all of the fuel energy can be recovered usefully. In this example, the heat input pro-

portional to 1250-350 = 90O 0 C compares with the excess output proportional to 435 0 C (line d'-d}

to represent an efficiency of (435/900), or 48%. If more fuel could be used, raising the maximum

temperature level at the same pressure, then more useful work could be obtained at nearly the same

efficiency.

The line d-a represents heat rejection. This could involve passing the exhaust gas through a

cooler before returning it to the compressor, and this would be a closed cycle. But, almost universally,

d-a reflects discharge to the ambient conditions and intake of ambient air (Fig. 57.1Z?). Figure 57. Ia

shows an open-cycle engine, which takes air from the atmosphere and exhausts back to the atmos-

phere. In this case, line d-a still represents heat rejection, but the path from d to a involves the whole

atmosphere and very little of the gas finds its way immediately from e to a. It is fundamental to this

cycle that the remaining 465 0 C, the vertical projection of line d-a, is wasted heat because point d is

at atmospheric pressure. The gas is therefore unable to expand further and so can do no more work.

Designers of simple cycle gas turbines—including aircraft engines—have pursued a course of

reducing exhaust temperature through increasing cycle pressure ratio, which improves the overall

efficiency. Figure 57.3 is identical to Fig. 51.2b except for the pressure ratio, which has been increased

from 16:1 to 24:1. The efficiency is calculated in the same manner. The total turbine work is

proportional to the temperature difference across the turbine, 1250-410 = 84O 0 C. The compressor

work, proportional to 430-15 = 415 0 C, is subtracted from the turbine temperature drop 840-415 =

425 0 C. The heat added to the cycle is proportional to 1250-430 = 82O 0 C. The ratio of the net work

to the heat added is 425/820 = 52%. The approximately 8% improvement in efficiency is accom-

panied by a 7O 0 C drop in exhaust temperature. When no use is made of the exhaust heat, the 8%

efficiency may justify the mechanical complexity associated with higher pressure ratios. Where there

is value to the exhaust heat, as there is in combined Brayton-Rankine cycle power plants, the lower

pressure ratio may be superior. Manufacturers forecast their customer requirements and understand

Fig. 57.3 Simple cycle gas turbine temperature-entropy diagram for high (24:1) pressure ratio

and 125O 0 C firing temperature.

the costs associated with cycle changes and endeavor to produce gas turbines featuring the most

economical thermodynamic designs.

The efficiency levels calculated in the preceding example are very high because many factors

have been ignored for the sake of simplicity. Inefficiency of the compressor increases the compressor

work demand, while turbine inefficiency reduces turbine work output, thereby reducing the useful

work output and efficiency. The effect of inefficiency is that, for a given temperature change, the

compressor generates less than the ideal pressure level while the turbine expands to a higher tem-

perature for the same pressure ratio. There are also pressure losses in the heat addition and heat

rejection processes. There may be variations in the fluid mass flow rate and its specific heat (energy

input divided by consequent temperature rise) around the cycle. These factors can easily combine to

reduce the overall efficiency.

57.1.2 A Brief History of Gas Turbine Development and Use

The use of a turbine driven by the rising flue gases above a fire dates back to Hero of Alexandria in

150 BC. It was not until AD 1791 that John Barber patented the forerunner of the gas turbine,

proposing the use of a reciprocating compressor, a combustion system, and an impulse turbine. Even

then, he foresaw the need to cool the turbine blades, for which he proposed water injection.

The year 1808 saw the introduction of the first explosion type of gas turbine, which in later forms

used valves at entry and exit from the combustion chamber to provide intermittent combustion in a

closed space. The pressure thus generated blew the gas through a nozzle to drive an impulse turbine.

These operated successfully but inefficiently for Karavodine and Holzwarth from 1906 onward, and

the type died out after a Brown, Boveri model was designed in 1939. 3

Developments of the continuous flow machine suffered from lack of knowledge, as different

configurations were tried. Stolze in 1872 designed an engine with a seven-stage axial flow compressor,

heat addition through a heat exchanger by external combustion, and a 10-stage reaction turbine. It

was tested from 1900 to 1904 but did not work because of its very inefficient compressor. Parsons

was equally unsuccessful in 1884, when he tried to run a reaction turbine in reverse as a compressor.

These failures resulted from the lack of understanding of aerodynamics prior to the advent of aircraft.

As a comparison, in typical modern practice, a single-stage turbine drives about six or seven stages

of axial compressor with the same mass flow.

The first successful dynamic compressor was Rateau's centrifugal type in 1905. Three assemblies

of these, with a total of 25 impellers in series giving an overall pressure ratio of 4, were made by

Brown, Boveri and used in the first working gas turbine engine, built by Armengaud and Lemale in

the same year. The exhaust gas heated a boiler behind the turbine to generate low-pressure steam,

which was directed through turbines to cool the blades and augment the power. Low component

efficiencies and flame temperature (828 0 K) resulted in low work output and an overall efficiency of

3%. By 1939, the use of industrial gas turbines had become well established: experience with the

Velox boiler led Brown, Boveri into diverging applications; a Hungarian engine (Jendrassik) with

axial flow compressor and turbine used regeneration to achieve an efficiency of 0.21; and the Sun

Oil Co. in the United States was using a gas turbine engine to improve a chemical process. 2

The history of gas turbine engines for aircraft propulsion dates from 1930, when Frank Whittle

saw that its exhaust gas conditions ideally matched the requirements for jet propulsion and took out

a patent. 4 His first model was built by British Thomson-Houston and ran as the Power Jets Type U

in 1937, with a double-sided centrifugal compressor, a long combustion chamber that was curled

round the outside of the turbine and an exhaust nozzle just behind the turbine. Problems of low

compressor and turbine efficiency were matched by hardware problems and the struggle to control

the combustion in a very small space. Reverse-flow, can-annular combustors were introduced in 1938,

the aim still being to keep the compressor and turbine as close together as possible to avoid shaft

whirl problems (Fig. 57.4). Whittle's first flying engine was the Wl, with 850 Ib thrust, in 1941. It

was made by Rover, whose gas turbine establishment was taken over by Rolls-Royce in 1943. A

General Electric version of the Wl flew in 1941. A parallel effort at General Electric led to the

development of a successful axial-flow compressor. This was incorporated in the first turboprop

engine, the TGlOO, later designated the T31. This engine, first tested in May of 1943, produced 1200

horsepower from an engine weighing under 400 kg. Flight testing followed in 1949. An axial-

compressor turbojet version was also constructed, designated the J35. It flew in 1946. The compressor

of this engine evolved to the compressor of the GE MS3002 industrial engine, which was introduced

in 1950 and is still in production. 5

A Heinkel experimental engine flew in Germany in 1939. Several jet engines were operational

by the end of the Second World War, but the first commercial engine did not enter service until 1953,

the Rolls-Royce Dart turboprop in the Viscount, followed by the turbojet de Havilland Ghost in the

Comet of 1954. The subsequent growth of the use of jet engines has been visible to most of the

world, and has forced the growth of design and manufacturing technology. 6 By 1970, a range of

standard configurations for different tasks had become established, and some aircraft engines were

established in industrial applications and in ships.

Gas turbines entered the surface transportation fields also during their early stages of development.

The first railway locomotive application was in Switzerland in 1941, with a 2200-hp Brown, Boveri

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