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

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

the wall temperature in two ways: (1) by reecting a large part of the incident

gas radiation; and (2) by providing a layer of thermal insulation between the

hot gas and the wall, thereby reducing the temperature of the supporting base

metal. A further benet may be gained if an oxidation-resistant base coat is

applied because it reduces the oxidation constraint on the choice of liner-wall

material.

An ideal TBC would be chemically inert and have good mechanical

strength, resilience to thermal shock, and resistance to wear and erosion.

Above all, it would have a low thermal conductivity and a thermal expansion

coefcient that is similar to that of the base metal. A typical TBC comprises

a metallic base coat (such as 0.1 mm of Ni Cr AL Y), plus one or two layers

of ceramic (such as partially yttria stabilized zirconia). Recent developments

in the strain tolerance of TBCs have reduced the necessity for an intermedi-

ate coat, and two-layer coatings are now sometimes specied for improved

mechanical integrity.

Plasma ame spraying is generally used to apply the ceramic and base

coat layers because it is found to provide durable and reproducible coatings.

A typical overall coating thickness is around 0.4–0.5 mm, which gives metal

temperature reductions of the order of 40–70 K, depending on the heat ux

through the liner wall. In this context, it should be noted that for a TBC to be

fully effective, there must be adequate heat removal from the “cold” side of

the liner wall. Inevitably, this means that liner geometries will become more

complex as various features (such as ns, ribs, etc.) are added to augment

the J convective heat transfer from the outer wall in order to derive full benet

from the TBC on the inner wall.

The reduction in wall temperature obtained from using a TBC may be cal-

culated by adding another term to Equation 8.25 to give

RCRC KK

ii11 22 12

+=+= =

−−

(8.46)

where

KktTT

i11−

(

)

−

(

)

TBC

KktTT

ii−

(

)

−

(

)

where T

is the hot-side surface temperature of TBC, T

is the cold-side sur-

face temperature of liner wall, and T

is the temperature at interface between

TBC and liner wall.

From the previous example we have

794 460 000321=−,. ,

W/m

1926 1176=−

(

)

W/m

Heat Transfer 351

229100 13 715=

(

)

−./ ,,W/m

921 810 400=−,.W/m

Solution of these equations for no TBC gave

1283 1265==KK.

The value of thermal conductivity used in the previous example for the

Nimonic wall was 26 W/(m·K). The thermal conductivity of a TBC is typically

an order of magnitude lower. Thus, for a coating of thickness 0.5 mm on a

wall of 1.2 mm thickness we have

KTT

i11

2600005

−

=−(./. )( ),

and

KTT

i−

=−

26 00012(./. )( ).

Substitution of these expressions for R

, C

, R

, C

, K

1–i

, and K

i–2

into Equation

8.46 yields

TTT

1304 1236 1220===KKK

Thus, the TBC has reduced the peak metal temperature from 1283 to 1236

K. Note that the hot-side temperature of the TBC is 21 Κ higher than the

metal surface without TBC.

8.14 Materials

Continuing efforts to improve engine performance and reduce fuel con-

sumption rely heavily on the development of new combustor materials to

withstand the harsher environmental conditions associated with operating

at higher pressures and temperatures. During the past 30 years, materials

and manufacturing processes have improved appreciably in terms of higher

temperature capability and lower cost. With the continuing development of

new materials and new methods of construction, this progress seems likely to

continue for the foreseeable future. The material requirements for future aero

engines have been reviewed in papers by Kirk [27] and Rosen and Facey [28].

Current production combustors are typically fabricated from sheets

of nickel- or cobalt-based alloys. These conventional materials still have

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

considerable development potential and will continue to dominate the aero-

nautical scene for some time to come. In the longer term, ceramics and ceramic

composites offer major improvements provided a number of inherent draw-

backs can be overcome.

The basic requirements for combustor materials can be listed as follows:

High-temperature strength•

Resistance to oxidation and corrosion•

Low density•

Low thermal expansion•

Low Young’s modulus•

Resistance to thermal fatigue•

Low cost•

Easy to fabricate•

High thermal conductivity•

From inspection of this list, it is clear that the traditional requirements

of good mechanical strength and oxidation resistance at high temperatures

are by no means the only desirable attributes of combustor materials. The

metal alloys now in common use are satisfactory for long-term operation at

temperatures up to around 1100 K. Oxidation becomes rapid at temperatures

above around 1300 K. However, developments in TBCs that incorporate an

oxidation-resistant base coat have tended to diminish the importance of a

material’s oxidation resistance.

The requirement of low cost is just as important for the combustor as for

other engine components, especially as most high-temperature materials

tend to be high-cost materials also. Low density (i.e., low weight) is clearly of

special importance for aero-engine combustors.

The steep temperature gradients that exist around features such as dilution

holes and cooling slots result in high thermal stresses. Thus, good thermal

fatigue strength is an important prerequisite for satisfactory liner life.

A high thermal conductivity is a desirable material property because it

facilitates the dissipation of heat from local “hot spots” in the liner wall. It is

especially benecial in pseudoporous wall constructions, such as Lamilloy,

Transply, and effusion cooling, which rely for their effectiveness on achiev-

ing a high rate of heat transfer from the liner wall to the cooling air owing

through it.

8.14.1 Metal Alloys

During the past 40 years, the nickel-based alloys Nimonic 75 and Hastelloy

X have been widely used in the UK and the United States, respectively, as sheet

materials for combustor liners. The success of these alloys is closely linked to

Heat Transfer 353

their ease of fabrication by forming and welding. They are satisfactory for

long-term operation at temperatures up to around 1100 K. Above this tem-

perature, the strength of these materials falls to unacceptable levels.

Modern combustor liners make extensive use of Nimonic 263 and the cobalt-

based Haynes 188 (HS 188), which is generally regarded as a replacement

for Hastelloy X because of its exceptionally high strength at temperatures

above 1070 K. Nimonic 263 has superior strength-temperature characteristics

to Nimonic 75, is easier to fabricate, and is relatively cheap. Nimonic 86 has

been developed as an alternative to Nimonic 263. It is highly oxidation resis-

tant and is readily fabricated, but its mechanical strength is inferior to that of

Nimonic 263 over most of its temperature range.

8.14.2 Ceramics

Although metal materials will continue to be of prime importance for the

foreseeable future, with developments taking place both in the materials

themselves and in manufacturing processes, for the long term, only ceramic

materials have the capability of meeting future engine requirements.

Ceramics have good mechanical strength at high temperatures, low

density, and are oxidatively stable at temperatures well beyond the capa-

bility of unprotected metals and alloys. These are clearly attractive quali-

ties where the reduction of liner-wall cooling air is an important design

requirement.

Silicon compounds are considered to be most promising and foremost

among these are silicon carbide and silicon nitride. Monolithic silicon nitride

and silicon carbide exhibit high strength and stiffness up to about 1680 and

1880 K, respectively [27]. The main drawback to these and other ceramics is

that, although they are strong at high temperatures, they do not possess the

toughness and ruggedness that engineers have become accustomed to with

ductile metals. To some extent, this problem can be alleviated by the incorpo-

ration of particles or whiskers to deect and arrest cracks [28]. An important

asset of this ceramic matrix construction is that when failure occurs, it does so

in a gradual and progressive manner instead of the catastrophic fracture that

is normally associated with monolithic materials. Also, continuous ceramic

laments allow ceramic components to tolerate minor aws and, generally, to

mimic metallic behavior, but with a higher temperature capability. Ceramic

composites that utilize continuous silicon carbon bers to reinforce silicon

carbide have been commercially available for some time [27].

A series of full-scale engine tests have been carried out at Solar on a Centaur

50 industrial gas turbine tted with silicon carbide composite combustion

liners. These liners functioned well over long periods of engine operation

with no apparent problems [29]. This progress encourages the notion that

developments in material properties, design methods, and manufacturing

processes are now reaching a stage where the introduction of ceramic lin-

ers in certain engine applications merits serious consideration. However, a

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

number of concerns still exist that are related to the inherent brittleness of

ceramics such as, for example, damage arising due to the ingestion of foreign

objects.

8.14.3 Mechanical integrity

In designing a combustor, the mechanical integrity of all the components

in the module is of vital importance. Predicting component life is an essen-

tial part of reliability and service warranties. There are several key elements

involved in ensuring the mechanical integrity of casing: pressure contain-

ment, life, fan-blade-off, and shock loads.

The casing must neither buckle nor rupture under the most extreme pres-

sure loadings seen by the engine, and pressure-vessel tests are performed

to assess this. The casing must last the life of the modern engine, which is

approximately: 15,000 hours for naval marine engine, 25,000 ights for a large

civil aero engine, or 100,000 hours for an industrial engine. If a fan blade is

lost during engine running, engine surges, or a bird is ingested, the casing

must not buckle and ange integrity must be maintained despite the large

vibration caused by the out-of-balance engine. In the event of a ame-out,

there is still high compressor-air pressure outside the combustor wall, while

the pressure inside the combustor rapidly collapses causing a buckling load

on the outer combustor wall. The combustor should be able to preserve its

mechanical integrity under these conditions. Finally, the casing should pass

various high-cycle fatigue tests.

8.15 Liner Failure Modes

Inspection of combustion-chamber components during engine overhaul

often reveals buckling and cracking of the liner. When such failure occurs

after a relatively short period of operation, the cause is usually found to be

errors in manufacture or design. Hot spots on the liner wall may be created

by igniter plugs, liner support struts, or any other object whose presence in

the annulus airow causes a breakdown in convective cooling in the wake

region immediately downstream. Even if these problems do not arise, after

long periods of engine operation it is customary to observe signs of dis-

tress in the form of liner buckling and cracking. Usually, the cracks origi-

nate at geometrical discontinuities, such as cooling rings and air-admission

holes, or at other points where residual stresses may be induced during

manufacture.

High thermal stresses are usually due to a combination of temperature

distribution and liner stiffness [1]. Cooling rings are not only stiff, but they

Heat Transfer 355

also operate close to the cooling-air temperature. Downstream of the cooling

ring, the metal temperature rises as the lm-cooling air loses its effective-

ness and a thermal gradient is created between the stiff ring and the relatively

weak metal downstream. As the hot areas try to expand, they are subjected

to compression by the surrounding cold metal, which causes them to yield.

After engine shutdown, these same areas are forced into tension by the sur-

rounding unyielded metal. After many cycles of engine startup and shut-

down, distortion and/or cracking can occur.

In general, buckling results from long periods of operation under a combi-

nation of high temperature and high-temperature gradients. Loss of material

thickness, or pitting of the surface due to oxidation or high-temperature cor-

rosion, clearly facilitates buckling and must be circumvented by using materi-

als with good oxidation resistance or by the application of TBCs. With aircraft

engines, buckling loads are highest when ame blowout occurs at maximum

power and low altitude. For marine and industrial engines running offshore,

corrosion caused by salt ingestion and by the high sulfur content of the fuel

are additional factors to consider in assessing the durability and failure of

combustor liners.

Nomenclature

convection heat ux from combustion gas to liner, W/m

convection heat ux from liner to annulus air, W/m

C/H carbon/hydrogen mass ratio of fuel

liner diameter (can) or height (annular), m

hydraulic mean diameter (4 × ow area/wetted perimeter), m

h heat-transfer coefcient, W(m

K conduction heat transfer along liner wall, W/m

1–2

conduction heat ux through liner wall, W/m

k thermal conductivity, W/(m·K)

L luminosity factor

mean beam length of radiation path, m

m mass velocity ratio [(ρU)

/(ρU)

]



mass ow rate, kg/s

Nu Nusselt number

P total pressure, kPa

Q heat ux, W/m

q fuel/air ratio by mass

radiation heat ux from combustion gas to liner, W/m

radiation heat ux from liner to casing, W/m

Re Reynolds number

slot Reynolds number (U

S/µ

)

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

S (x/ms)

025

s slot height, m

T absolute temperature, Κ

w,ad

adiabatic wall temperature, Κ

ΔT

comb

temperature rise due to combustion, Κ

t slot lip thickness, m

liner wall thickness, m

U velocity, m/s

x distance downstream of slot, m

α absorptivity

ε emissivity

η lm-cooling effectiveness

local combustion efciency

µ dynamic viscosity, kg/(ms)

ρ density, kg/m

σ Stefan–Boltzmann constant, 5.67 × 10

−8

W/(m

)

Subscripts

a air

an annulus

g gas

L liner

pz primary zone

ref reference value

s slot

w liner wall

1 ame side of liner wall

2 coolant side of liner wall

3 combustor inlet condition

References

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ed., Design of Modern Gas Turbine Combustors, pp. 343–476, Academic Press,

New York, NY, 1990.

2. Lefebvre, A. H., and Herbert, M. V., “Heat-Transfer Processes in Gas Turbine

Combustion Chambers,” Proceedings of the Institution of Mechanical Engineers,

Vol. 174, No. 12, pp. 463–73, 1960.

3. Hottel, H. C., “Some Problems in Radiative Transport,” in D. B. Olfe and

S. S. Penner, eds., International Developments in Heat Transfer, ASME, New York,

pp. 126–33, 1960.

4. Reeves, D., “Flame Radiation in an Industrial Gas Turbine Combustion

Chamber,” National Gas Turbine Establishment, UK, NGTE Memo M285, 1956.

Heat Transfer 357

5. Fishenden, M., and Saunders, O. A., An Introduction to Heat Transfer, Oxford

University Press, New York, 1950.

6. Kretshmer, D., and Odgers, J., “A Simple Method for the Prediction of Wall

Temperatures in a Gas Turbine Combustor,” ASME Paper 78-GT-90, 1978.

7. Lefebvre, A. H., “Inuence of Fuel Properties on Gas Turbine Combustion

Performance,” AFWAL-TR-84-2104, 1985.

8. Hottel, H. C., “Radiant Heat Transmission,” in W. H. McAdams, ed., Heat

Transmission, 3rd ed., Chap. 4, McGraw-Hill, New York, 1954.

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Region,” Journal of Heat Transfer, Vol. 95, pp. 265–6, 1973.

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Cooled Wall Temperatures in Gas Turbine Combustion Chambers,” ASME

Paper 72-WA/HT-24, 1972.

12. Kacker, S. C., and Whitelaw, J. Η., “An Experimental Investigation of the

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Oxford, 1969.

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a Plate Inserted in the Exhaust Stream of a Gas Turbine Combustor,” NASA TN

D-7156, 1973.

15. Mularz, E. J., and Schultz, D. F., “Measurements of Liner Cooling Effectiveness

within a Full-Scale Double-Annular Ram-Induction Combustor,” NASA TN

D-7689, 1974.

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Washington, DC, 1980.

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“Energy Efcient Engine Combustor Test Hardware—Detailed Design Report,”

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Improved Combustor Wall Cooling Techniques,” ASME Paper 80-GT-66, 1980.

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pp. 268–76, 1980.

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Technology Program,” Phase II, Final Report, 1989.

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

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359

Emissions

9.1 Introduction

Pollutant emissions from combustion processes have become of great pub-

lic concern due to their impact on health and the environment. The past

decade has witnessed rapid changes both in the regulations for controlling

gas turbine emissions and in the technologies used to meet these regula-

tions. During this period, the consumption of fuel by civil aviation has

increased to the extent that air transport is now perceived as one of the

world’s fastest growing energy-use sectors. At the same time, stationary

gas turbines have become rmly established as prime movers in the gas

and oil industry, and have acquired new ranges of application in combined

cycle plants and in many areas of utility power generation. All these devel-

opments have been accompanied by continuous and increasing pressure

on the combustion engineer to reduce pollutant emissions from all types

of gas turbines.

The material presented in this chapter is divided into six main sections:

A general overview of emissions concerns and the regulations that •

have been introduced to address these concerns

The mechanisms of pollutant formation and the methods employed •

in alleviating pollutant emissions from conventional gas turbine

combustors

The use of variable geometry and staged combustion for emissions •

reduction by control of ame temperature

Basic approaches to the design of “dry” low oxides of nitrogen (NO•

)

and “ultralow” NO

combustors

Alternative methods for achieving ultralow NO•

emissions, includ-

ing rich-burn, quick-quench, lean-burn, and catalytic combustors

Correlations for NO•

and carbon monoxide (CO) emissions.

Emissions from new alternative and bio-fuels used in gas turbines are pre-

sented and discussed in Chapter 10.