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

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

gain is sufciently large, the heat-release uctuations will drive the pressure

uctuations. The magnitude of the gain is determined by the value of n.

Similar time lag models have been applied to a variety of industrial burn-

ers (see the review by Putnam [31]), and to ramjets [46,47]. More recent exten-

sions of the τ–n model to include nonlinear acoustics in combustors have been

described by Culick [48]. To a large extent, the success of the τ–n approach

depends on predicting the time lag. The main uncertainties lie in the estima-

tion of mixing rates and, in liquid-fuelled systems, evaporation rates also.

Advances in computational uid dynamics (CFD) allow the time history

of complex reacting ow elds to be computed without the need to specify a

combustion time lag. For example, Menon et al. [49,50] have used CFD analy-

sis to describe combustion instabilities in ramjets. The main drawback to

this approach, in addition to the long computational time, is that the results

obtained tend to be combustor-specic, so that few general conclusions can

be drawn.

Janus and Richards [51] have developed a simple, nonlinear model in which

the combustion process is represented by a well-stirred reactor with nite

kinetics. The model was developed to provide explanations for specic exper-

imental observations on LPM gas turbine combustors, and to provide guid-

ance in developing active control strategies. Conservation equations for the

combustor and fuel injector provide a set of ordinary differential equations

that can be solved on a personal computer. Comparison with experimental

data shows good agreement with the predictions of the model that, accord-

ing to Janus and Richards, can be used to examine stability trends associated

with changes in fuel/air ratio, mass ow rate, geometry, ambient conditions,

and other relevant parameters.

Recently, Lieuwen and Yang [52] have provided an excellent overview

of the operational experience, fundamental mechanisms, and modeling of

combustion instabilities in gas turbine engines. They state that combustion

instabilities routinely constrain the operating envelope and power output of

gas turbines for power generation and damage combustor liners, transition

pieces, and fuel nozzles. For aero engines, instabilities appear in a few cases,

mostly in the development stage. Military engines, however, have experi-

enced major problems with low-frequency instabilities in augmentors. This

work describes research performed through 2005.

References

1. Mahan, J. R., and Karchmer, Α., “Combustion and Core Noise,” Aero Acoustics of

Flight Vehicles: Theory and Practice, Vol. 1; Noise Sources, WRDC Technical Report

90-3052, 1991.

Combustion Noise 311

2. Bragg, S. L., “Combustion Noise,” Journal of the Institute of Fuel, January,

pp. 12–16, 1963.

3. Thomas, Α., and Williams, G. T., “Flame Noise: Sound Emission From Spark-

Ignited Bubbles of Combustible Gas,” Proceedings of the Royal Society, London,

Series A, Vol. 294, No. 1439, pp. 449–66, 1966.

4. Ballal, D. R., and Lefebvre, A. H., “Turbulence Effects on Enclosed Flames,” Acta

Astronautica, Vol. 1, pp. 471–83, 1974.

5. Ballal, D. R., and Lefebvre, A. H., “The Structure and Propagation of Turbulent

Flames,” Proceedings of the Royal Society, London, Series A, Vol. 344, pp. 217–34,

1975.

6. Strahle, W. C., “Some Results in Combustion Generated Noise,” Journal of Sound

and Vibration, Vol. 23, No. 1, pp. 113–25, 1972.

7. Strahle, W. C., and Muthukrishnan, M., “Correlation of Combustor Rig Sound

Power Data and Theoretical Basis of Results,” AIAA Journal, Vol. 18, No. 3,

pp. 269–74, 1980.

8. Shivashankara, B. N., and Crouch, R. W., “Noise Characteristics of a Can-Type

Combustor,” Journal of Aircraft, Vol. 14, No. 8, pp. 751–6, 1977.

9. Kazin, S. B. and Emmerling, J. J., “Low Frequency Core Engine Noise,” ASME

Paper 74-WA/Aero-2, 1974.

10. Ho, P. Y., and Tedrick, R. N., “Combustion Noise Prediction Techniques

for Small Gas Turbine Engines,” in M. J. Crocker, ed., INTER-NOISE 72

Proceedings, Inst. of Noise Control Engineering, Poughkeepsie, New York, pp.

507–12, 1972.

11. Strahle, W. C., and Shivashankara, B. N., “Combustion Generated Noise in Gas

Turbine Combustors,” NASA CR-134843, 1974.

12. Ho, P. Y., and Doyle, V. L., “Combustion Noise Prediction Update,” AIAA Paper

79-0588, 1979.

13. Mathews, D. C., and Rekos, N. F., Jr., “Prediction and Measurement of Direct

Combustion Noise in Turbopropulsion Systems,” Journal of Aircraft, Vol. 14,

No. 9, pp. 850–9, 1977.

14. Rayleigh, Lord, The Theory of Sound, Vol. 2., Dover, New York, p. 226, 1945.

15. Seto, S. P., private communication, 1989.

16. Kenworthy, M. J., Bahr, D. W., Mungur, P., Burrus, D. L., Mehta, J. Μ., and

Cifone, A. J., “Dynamic Instability Characteristics of Aircraft Turbine Engine

Combustors,” AGARD CP-450, ISBN 92-835-0503-4, 1989.

17. Janus, M. C., Richards, G. Α., Yip, M. J., and Robey, E. H., “Effects of Ambient

Conditions and Fuel Composition On Combustion Stability,” ASME Paper

97-GT-266, 1997.

18. Keller, J. Ο., Bramlette, Τ. Τ., Westbrook, C. Κ., and Dec, J. Ε., “Pulse

Combustion: The Importance of Characteristic Times,” Combustion and Flame,

Vol. 75, pp. 33–44, 1989.

19. Vandsburger, U., McManus, K., and Bowman, C., “Effects of Fuel Spray

Vaporization on the Stability Characteristics of a Dump Combustor,” AIAA

Paper 89-2436. Presented at the 25th Joint Propulsion Conference, July 10–12,

Monterey, CA, 1989.

20. Mehta, J., Mungur, P., Dodds, W., Bahr, D., and Clouser, S., “Fuel Effects on

Gas Turbine Combustor Dynamics,” AIAA Paper 90-1957. Presented at the 26th

Joint Propulsion Conference, Orlando, FL, 1990.

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

21. Janus, M. C., and Richards, G. A., “Results of a Model for Premixed Combustion

Oscillations,” Presented at the American Flame Research Committee Meeting,

September 30–October 1, Baltimore, MD, 1996.

22. McManus, K. R., Poinsott, T., and Candel, S. M., “A Review of Active Control

of Combustion Instabilities,” Progress in Energy and Combustion Science, Vol. 19,

pp. 1–29, 1993.

23. Scalzo, A. J., Sharkey, W. T., and Emmerling, W. C., “Solution of Combustor Noise

in a Coal Gasication Cogeneration Application,” Turbomachinery International,

September/October, pp. 22–7, 1989.

24. Hudson, R. H., “Pressure Oscillations in a Combustor with Dual-Orice Fuel

Injection,” ASME Paper 67-GT-23, 1967.

25. Ibrahim, M., Darling, D., Sanders, T., and Zaller, M., “Dynamic Response of Fuel

Nozzles for Liquid-Fueled Gas Turbine Combustors,” ASME Paper 96-GT-54,

1996.

26. Richards, G. A., “Gas Turbine Combustion Instability.” Proceedings of the Spring

Technical Meeting of the Central States Section of the Combustion Institute,

May 5–7, St. Louis, MO, 1996.

27. Richards, G. A., and Janus, M. C., “Characterization of Oscillations During

Premix Gas Turbine Combustion,” ASME Paper 97-GT-244, 1997.

28. Keller, J. J., “Thermoacoustic Oscillations in Combustion Chambers of Gas

Turbines,” AIAA Journal, Vol. 33, No. 12, pp. 2280–7, 1995.

29. Richards, G. A., Gemmen, R. S., and Yip, M. J., “A Test Device for Premixed Gas

Turbine Combustion Oscillations,” DOE/METC-96/1027 (DE96004367), 1996.

30. Candel, S. M., “Combustion Instabilities Coupled by Pressure Waves and Their

Active Control,” Twenty-Fourth (International) Symposium on Combustion, The

Combustion Institute, Pittsburgh, PA, pp. 1277–96, 1992.

31. Putnam, A. A., Combustion Driven Oscillations in Industry, American Elsevier,

New York, 1971.

32. Harrje, D. T., and Reardon, F. H., “Liquid Propellant Rocket Combustion

Instability,” NASA SP-194, 1972.

33. Yang, V., and Anderson, W., eds., Liquid Rocket Combustion Instability, AIAA,

Cambridge, MA, 1995.

34. Gutmark, E., Wilson, K. J., Schadow, K. C., Stalnaker, R. A., and Smith, R.

A., “Combustion Characteristics and Passive Control of an Annular Dump

Combustor,” AIAA Paper 93-1772, 1993.

35. Richards, G. A., Yip, M. J., Robey, E., Cowell, L., and Rawlins, D., “Combustion

Oscillation Control by Cyclic Fuel Injection,” Journal of Engineering for Gas

Turbines and Power, Vol. 119, pp. 340–3, 1997.

36. McManus, K. R., Vandsburger, U., and Bowman, C. T., “Combustor Performance

Enhancement Through Direct Shear Layer Excitation,” Combustion and Flame,

Vol. 82, pp. 75–92, 1990.

37. Choudhury, P. R., Gerstein, M., and Mojaradi, R., “A Novel Feedback Concept

for Combustion Instability in Ramjets,” 22nd JANNAF Combustion Meeting,

1985.

38. Bloxsidge, G. J., Dowling, A. P., Hooper, N., and Langhorne, P. J., “Active Control

of Reheat Buzz,” AIAA Journal, Vol. 26, pp. 783–90, 1988.

39. Langhorne, P. J., Dowling, A. P., and Hooper, N., “Practical Active Control

System for Combustion Oscillations,” Journal of Propulsion and Power, Vol. 6,

No. 3, pp. 324–33, 1990.

Combustion Noise 313

40. Sivasegaram, S., and Whitelaw, J. Η., “Active Control of Combustors with

Several Frequency Modes,” American Society of Mechanical Engineers DSC, Vol. 38,

pp. 69–74, 1992.

41. Schadow, K. C., Hendricks, E. W., and Hansen, R. J., “Recent Progress in the

Implementation of Active Combustion Control,” Proceedings of the 18th ICAS

Congress, Beijing, China, September 20–25, Vol. 1, Paper No. ICAS-92-2.5.3, pp.

942–52, 1992.

42. Schadow, K. C., Wilson, K. J., Gutmark, E., Yu, K., and Smith, R. Α., “Periodic

Chemical Energy Release for Active Control,” 11th ISABE-International Symposium

on Air Breathing Engines, Vol. 1, AIAA, Washington, DC, pp. 479–85, 1993.

43. Brouwer, J., Ault, Β. Α., Bobrow, J. Ε., and Samuelsen, G. S., “Active Control

for Gas Turbine Combustors,” The Twenty-Third (International) Symposium on

Combustion, The Combustion Institute, Pittsburgh, PA, pp. 1087–92, 1990.

44. Gemmen, R. S., Richards, G. Α., Yip, M. J., and Norton, T. S., “Combustion

Oscillation Chemical Control Showing Mechanistic Link to Recirculation Zone

Purge Time,” Eastern States Section Meeting of the Combustion Institute, Worcester,

MA, 1995.

45. Crocco, L., and Cheng, S. I., “Theory of Combustion Instability in Liquid

Propellant Rocket Motors,” AGARD Monograph, No. 8, Butterworths, London,

1956.

46. Reardon, F. H., “Very Low-Frequency Oscillations in Liquid Fuel Ramjets,”

NATO AGARD Conference Proceedings No. 450, 1989.

47. Yu, K. H., Trouve, Α., and Daily, J. W., “Low-Frequency Pressure Oscillations in a

Model Ramjet Combustor,” Journal of Fluid Mechanics, Vol. 232, pp. 47–72, 1991.

48. Culick, F. E. C., “Some Recent Results for Nonlinear Acoustics in Combustion

Chambers,” AIAA Journal, Vol. 32, No. 1, pp. 146–69, 1994.

49. Menon, S., and Jou, W. H., “Large-Eddy Simulations of Combustion Instabilities

in an Axisymmetric Ramjet Combustor,” Combustion Science and Technology,

Vol. 84, pp. 51–79, 1991.

50. Menon, S., “Secondary Fuel Injection Control of Combustion Instability in a

Ramjet,” Combustion Science and Technology, Vol. 100, pp. 385–93, 1994.

51. Janus, M. C., and Richards, G. Α., “A Model for Premixed Combustion

Oscillations,” Technical Note DOE/METC-96/1026 (DE96004366), 1996.

52. Lieuwen, T.C. and Yang, V., Combustion Instabilities in Gas Turbine Engines,

Progress in Astronautics and Aeronautics, Vol. 210, AIAA, Reston, VA, 2005.

Bibliography

Throughout this chapter, frequent reference has been made to the paper on combus-

tion and core noise by Mahan and Karchmer and to the articles on the active control of

combustion instabilities by Candel and McManus et al. These publications are listed

above as References [1], [30], and [22], respectively. They are strongly recommended

as additional reading for anyone seeking a comprehensive and in-depth review of

these important topics. For further information on current research activities relating

to combustion oscillations in modern premix gas turbine combustors, refer to [52]

which summarizes the most up to date research in this area through 2005.

315

Heat Transfer

8.1 Introduction

In a modern combustor, the temperature of the gases released by the com-

bustion process may peak over 2100°C; this is much higher than the melt-

ing point of the combustor ame tube and turbine blades. Therefore, the

designer must adequately cool all the metal surfaces exposed to the hot gases

and improve structural integrity and durability. Furthermore, the amount of

cooling air must be minimized to maximize the air available for emissions

control.

The mechanical stresses experienced by the combustor liner are small in

comparison with those to which many other engine components are sub-

jected. However, it is called on to withstand high temperatures and steep

temperature gradients that threaten its structural integrity. To ensure a satis-

factory liner life, it is important to keep temperatures and temperature gra-

dients down to an acceptable level. Just what these levels are has never been

clearly dened, but for the nickel- or cobalt-based alloys in common use,

such as Nimonic 75, Hastelloy X, and HS188, the maximum operating tem-

perature should not exceed 1100 K. The mechanical strength of these materials

declines rapidly at temperatures above this level. The practical implication

of this limit is that some means must be provided to supplement the removal

of heat from the liner walls, which normally occurs through radiation to the

combustor casing and convection to the annulus air. The traditional method

has been to provide a lm of cool air along the inner surface of the liner.

Since the early 1960s, the need to develop more efcient methods of liner-

wall cooling has become increasingly important. The reasons for this are

fourfold.

1. Substantial reductions in engine fuel consumption continue to be

achieved by the use of higher pressure ratios, higher turbine entry

temperatures and, on aircraft engines, by higher bypass ratios.

Unfortunately, increases in pressure ratio raise the amount of heat

transferred to the liner walls by radiation. Moreover, the accompany-

ing increase in combustor inlet temperature impairs the ability of the

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

annulus air to cool the walls by convection. Thus, as pressure ratios

rise, the problem of wall cooling becomes more severe, to the extent

that on many engines more than one-third of the total combustor

airow is used in lm cooling the liner.

2. During the past 20 years, as the regulations governing pollutant

emissions have become increasingly stringent, the requirement

for lower levels of nitric oxides has resulted in more air being allo-

cated to combustion. This trend has led to a continuing decline in

the amount of air available for liner-wall cooling. Furthering this

trend was the knowledge that reductions in lm-cooling air are

highly benecial in lowering the emissions of carbon monoxide and

unburned hydrocarbons (see Chapter 9).

3. Increases in turbine inlet temperature call for improvements in com-

bustor pattern factor in order to maintain the integrity of the hot

sections downstream. Since lm-cooling air ows along the liner

wall, it makes no contribution to mixing. Thus, reductions in wall-

cooling requirements make a direct contribution to pattern factor

improvement by releasing more air for dilution zone mixing.

4. As noted by Dodds and Bahr [1], as combustor operating temperatures

have increased, component durability expectations have also risen.

Life expectations for combustor liners of early engines were only a

few hundred hours between repairs. Today, the demand from cus-

tomers is for many thousands of operating hours before combustor

maintenance is required.

For all these reasons, it is important to improve the effectiveness of the vari-

ous cooling devices that are now in widespread use, and to develop new

cooling schemes that are even more economical in their use of cooling air.

Also, heat transfer modeling of metal temperatures is necessary to deter-

mine the displacement, thermal stresses, and component life. To calculate

metal temperatures, it is necessary to input material property data, engine

performance data, air system data, and experimentally validated heat trans-

fer coefcients.

In this chapter, the heat-transfer processes that govern liner-wall temper-

atures are discussed in some detail, along with the methods employed to

combat the heat ux to the wall, thereby achieving an acceptable liner life.

8.2 Heat-Transfer Processes

For the purpose of analysis, a liner may be regarded as a container of hot

owing gases surrounded by a casing, with air owing between the container

Heat Transfer 317

and the casing. Broadly, the liner is heated by radiation and convection from

the hot gases inside it; it is cooled by radiation to the outer casing and by

convection to the annulus air. The relative magnitudes of the radiation and

convection components depend on the geometry and operating conditions of

the system. Under equilibrium conditions, the liner temperature is such that

the internal and external heat uxes at any point are just equal. Loss of heat

by conduction along the liner wall is comparatively small and, usually, may be

neglected.

The heat-transfer model shown in Figure 8.1 includes only the axial varia-

tion of properties. All properties are assumed to be constant around the cir-

cumference at any axial location.

Under steady-state conditions, the rate of heat transfer into a wall element

must be balanced by the rate of heat transfer out. Therefore, for an element

with inside surface area

∆A

RCKA RCAKA

11 22 12

121

(

)

∆=+

(

)

∆=∆

−www

(8.1)

K, the heat conduction along the liner wall, is always negligibly small com-

pared to the radiation and convection terms. Also, the liner wall is usually so

thin that

∆≅∆AA

. Thus, Equation 8.1 can be simplied to

RCRC K

11 22 12

+=+=

−

(8.2)

where K

1–2

is the conduction heat transfer through a solid liner wall due to a

temperature gradient within the wall, i.e.,

−

=−

(

)

(8.3)

= convection heat flux

to annulus air

= radiation heat flux

to casing

1–2

= conduction heat flux

through wall

= convection heat

flux from gas

= radiation heat

flux from gas

K = conduction along liner wall

Figure 8.1

Basic heat-transfer processes.

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

In the following sections, expressions for R

, C

, R

, and C

are derived.

When inserted into Equation 8.2, they allow liner-wall temperatures to be

calculated for any stipulated combustor inlet conditions of pressure, tempera-

ture, air mass ow rate, and air/fuel ratio.

8.3 Internal Radiation

In most gas turbine combustors, a sizeable portion of the heat transferred

from the hot gases contained within the liner to the liner wall is by radiation.

In fact, in those regions of the combustor where the cooling air injected at the

walls forms an effective barrier between the hot gas and the wall, radiation

represents the only mechanism by which heat can be transferred from the

gas to the wall. For the combustion gases generated by gas turbine fuels, the

total emitted radiation has two components: (1) the “nonluminous” radia-

tion that emanates from certain heteropolar gases, notably carbon dioxide and

water vapor, and (2) the “luminous” radiation that depends on the number

and size of the solid particles (mainly soot) in the ame. Both modes of

radiant heat transfer enter into calculations of the liner-wall temperature

and the amount of air to be employed in liner-wall cooling.

8.3.1 radiation from Nonluminous gases

The rate of heat transfer by nonluminous radiation from a gas to its enclosure

can be calculated from knowledge of the size and shape of the gas volume

and its mean or “bulk” conditions of pressure, temperature, and chemical

composition. Let us consider the radiation heat exchange between a gas

at temperature T

and the surface of a black-body container at temperature

. While the black surface emits and absorbs heat at all wavelengths, the

gas emits only a few narrow bands of wavelengths and absorbs only those

wavelengths included in its emission bands. The net radiant heat transfer is

given by

RTT

=−

(

)

σε α

gg g

(8.4)

where σ is the Stefan–Boltzmann constant = 5.67 × 10

−8

W/( m

), ε

is the

gas emissivity at temperature T

, and α

is the gas absorptivity at tempera-

ture

Both ε

and α

are functions of gas composition. However, ε

relates to

the emission of radiation from the gas to the wall and depends on T

, but

applies to the absorption by the gas of radiation from the wall, and hence

depends on

Heat Transfer 319

In practice, the surface exposed to the ame is not black, but has an effective

absorptivity that is less than unity. For most practical purposes, this effect

may be accounted for by introducing the factor 0.5(1 + ε

) to obtain

RTT

05 1

(

)

−

(

)

.,σεε

wggg

(8.5)

in which ε

is dependent on the material, temperature, and degree of oxida-

tion of the wall. Approximate mean values of ε

at typical liner-wall tem-

peratures for Nimonic and stainless steel are 0.7 and 0.8, respectively.

Investigation over a wide range of values has shown [2] that to a suf-

ciently close approximation













15.

(8.6)

Hence, Equation 8.5 may be rewritten as

RTTT

15 25

05 1

(

)

−

(

)

σεε

wggg

(8.7)

In practice, the composition and temperature of the hot gases are far from

homogeneous. Thus, the variables in the above expression must be repre-

sented by mean or “effective” values. Improvements in accuracy could be

achieved by the method of zoning, as described by Hottel [3], but this would

demand a more exact knowledge of the distribution of fuel and temperature

in the combustion zone than is available for most current chamber designs.

The bulk or mean gas temperature, T

, is obtained as the sum of the chamber

entry temperature, T

, and the temperature rise due to combustion ΔT

comb

TT T

gcomb

=+∆

(8.8)

ΔT

comb

may be obtained from standard temperature-rise curves. When these

curves are used, the appropriate value for the fuel/air ratio is the product of

the local fuel/air ratio and the local level of combustion efciency. Most heat-

transfer calculations are carried out at high pressures, for which it is reason-

able to assume a combustion efciency of 100%.

Values of ε

for nonluminous ames may be obtained from the following

approximate formula due to Reeves [4].

gbg

=− −

(

)









−

2290

exp,

PqlT

(8.9)

where Ρ is the gas pressure (kPa), T

is the gas temperature (Κ), l

is the beam

length (m), and q is the fuel/air ratio by mass.