Sawyer's Gas Turbine Combustor - Chapter 50001
Short Description
Sawyer's Gas Turibine on Combustor...
Description
CHAPTER
5 COMBUSTOR
DESIGN
by Jerry O. Melconian and Ashok T. Modak
3.3
CONTENTS
1.0 2.0
3.0
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5-2 5-5 5-5 5-5 5-6
.. .
.......
5-7 5-7 5-8 5-8 5-9
.............
5-9
Nomenclature Introduction .. .• .• .• .• .• .• .• .• 2.1 Background •••••••••••• Operating Conditions .• 2.2 Today's Requirements 2.3 2.4 The Combustion ChamberFirst Principles ••••••••• 2.4.1 Combustion Zone. 2.4.2 Dilution Zone 2.4.3 Wall Cooling •••• Combustor Configurations 3.1 Combustors Classified by Geometry 3.1.1 Multi-Can Combustors 3.1.2 Annular Combustors 3.1. 3 Can -Annular Combustors 3.1.4 Single Can Combustors 3.2 Combustors Classified by Air Distribution 3.2.1 Straight- Through Combustors 3.2.2 Reverse Flow Combustors 3.2.3 Regenerative Combustors 3.2.4 Single Vortex Combustors 3.2.5 Two-Stage Combustors 3.2.6 Variable Geometry Combustors ••••• 3.2.7 Fully Premixed Combustors 3.2.8 Catalytic Combustors
4.0
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5.0 5-10
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·....... 5-10 ·..... 5-10 ·..... 5-10 ....... 5-10
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6.0
5-11
5-14
5-14
5-15 5-15 5-16 5-16 5-16 5-16 5-17 5-17
.. 5-18
5-10
5-14
...............
...........
5-9
5-14
..........
........... 5-14 ... 5-15 ......... 5-15
5-9
5-10
Combustors Classified by Fuel Injection 3.3.1 Liquid Fuels Downstream Injection 3.3.2 Slinger-Type Fuel Injection 3.3.3 Two-Fluid Atomizers 3.3.4 Vaporizers •••••• 3.3.5 Gaseous Fuels ••• 3.3.6 Liquid - Gaseous System ••••••••• The Design Method 4.1 Combustor Performance •• 4.1.1 Combustion Efficiency 4.1. 2 Pressure Drop ••• 4.1. 3 Temperature Profile ••••••••• 4.1.4 Stability Limits 4.1. 5 Altitude Limits •• Design Specifications •••• 4.2 4.3 Design Sequence Selection of Combustor Type ••• Introduction ••••••••••• 5.1 5.2 Aircraft Systems 5.3 Industrial and Other Engines 5.3.1 Small Units (Includes Automotive and APU Applications) •••• 5.3.2 Medium and Large Units The Fuel Injection System ••••• 6.1 Introduction 6.2 Selection of Fuel Injection System Combustor Dimensions •••••••• 7.1 Determination of Reference Area 7.1.1 Aerodynamic Considerations •••
7.0
............. .................... .............. ................
5-18 5-18 5-19 5-19 5-19 5-20 5-20
5-20 5-20 5-22 5-22 5-22 5-23 5-23 5-23 5-1
CHAPTER 5
8.0 9.0 10.0
11.0
12.0 13.0
14.0
15.0
16.0 5-2
Chemical CCombustion) 5-23 Considerations 7.2 Determination of 5-25 Combustor Area ••••••• 7. 3 Selection of Appropriate Casing and Combustor 5-27 Areas .•.............. 7.4 Preliminary Estimate of Remaining Features of 5-28 Combustor •••••••••••• 5-29 Diffuser Design ••••••••••••• 5-33 Swirier Design •••••••••••••• Calculation of Flame 5-35 Temperature ••.••.•••••••••• 5-35 10. 1 Recirculation Zone •••••• 10. 2 Remainder of Primary Zone . 5-36 10. 3 Secondary Zone Temperature •.••••••••• 5-36 10.4 Dilution Zone Temperature ••••••••••• 5-36 10.5 The Equivalent Gas Temperature Adjacent 5-36to the Film Tg •••••••• Heat Transfer to Combustor Walls . 5-36 11. 1 Uncooled Wall Temperature ••••••••••• 5-36 11 . 2 Film Cooling •••••••••• 5-37 11 . 2. 1 Film Cooling Calculations ••••• 5-38 5-42 Design of Air Admission Holes Ignition Considerations •••••••• 5-45 13 . 1 Selection of Igniter Type •............•... 5-46 13 . 2 Igniter Location •••••••• 5- 46 Performance and Pollutant Predictions ••••••••••••••••• 5 -46 14.1 Combustion Efficiency) Stability and Ignition •••• 5 - 46 14.2 Pollutant Prediction (with no Test Data Available) ••••••••••••• 5-48 14.2.1 Carbon Monoxide· 5-48 14.2.2 Hydrocarbons •••• 5 -48 14.2.3 Carbon and Smoke Formation •••••• 5- 50 14.2.4 Oxides of Nitrogen (NOx) •• 5-51 14.2.5 Oxides of Sulfur-· 5-53 14 . 2 ..6 Ash··.········· 5- 53 A Review of Modeling Techniques •••••••••••••••••• 5 - 54 15. 1 Zero - Dimensional Models • 5 - 54 15. 2 One -Dimensional Models • 5 - 55 15. 3 Two -Dimensional Models • 5 - 56 15 .4 Three -Dimensional Models ••••••••••••••• 5- 56 References •••••••••••••••••• 5- 57
ABSTRACT
7 . 1. 2
s
-
The combustor is likely to be the life -Iirniting component of a modern, high compression ratio gas turbine engine. Durable, high performance engines require superior combustor designs. The present work surveys the state of the art of combustor design. The aim here is to provide information for producing designs which require minimum development time. As such, the emphasis is on the practical rather than theoretical aspects of combustor design. Basic design principles and performance constraints are discussed [1] and the various combustor configurations are surveyed. Criteria for selecting a suitable combustor configuration and fuel injection system are examined, followed by design calculations for the dimensions of the casing, the liner, the diffuser, and the swirlers. Calculations of gas temperatures in the various zones of the combustor and liner wall temperatures in the presence of film cooling are performed along with design calculations for the dimensions of the air admission holes. Criteria for selecting the type of igniter and its location are analyzed, followed by calculations for combustor performance and pollutants emitted from the combustor. The present work concludes with a brief review of theoretical modeling techniques. 1.0 NOMENCLATURE Symbol A
Name Area Cft2; m2) Cross-sectional area of flame tube (liner) (ft 2; m 2)
Aref
Maximum casing cross-sectional Cft2; m2)
area
AR
Area ratio
A/F
Air-to-fuel
B
Blockage factor
b
Inlet temperature factor (deg R; deg K)
C
E/R
mass ratio
(deg R;
deg K)
Heat transferred by convection to inner flame tube surf ace CBtul (s . ft 2) ; W·m-2) Heat abstracted by convection from outer flame tube surface CBtul (s . ft 2) ; W·m-2) Coefficient of discharge
COMBUSTOR m*
Ratio-moles
m
Mass flow (lbm/s ; kg-s-1)
Empirical constant
n
Number
D
Diameter,
n
Reaction order
Dref
Maximum casing diameter or width
P
Total pressure (Ibflft2 Pa)
DZ
Dilution
Cd ,s
Coefficient
of discharge of snout
C/H
Carbon-to-hydrogen
D*
mass ratio
usually casing (ft ; m)
(ft ; m)
usually flame tube (ft;
Hydraulic mean diameter, flame tube (ft; m) Hole diameter Activation J/(mol-
(ft ; m)
or psi or atm;
PZ
Primary zone
p
Static pressure (Ibflft 2; Pa)
q
Dynamic pressure (Ibflft 2; Pa)
R
Universal gas constant [Btul (lb mole - R); J - mol- lK-l)
energy [Btul (lb mol- R) ;
K)]
EI FAR
Fuel/air
f
Reaction order in fuel
K
Rate constants
K
Velocity correction
K
Hole pressure loss factor (Pan - Pi) Iqan
k
m)
usually
Emission index (g/kg fuel)
k
oxygen
Pressure ratio
zone
Diameter,
inerts/moles
Heat transferred by radiation to inner surface of flame tube wall (Btul (s - ft 2); W -m - 2)
mass ratio Heat lost through radiation from outer surface of flame tube wall (Btul (s - ft 2); W -m - 2) Gas constant for air [Btul (lbm - R) ; J - kg-1K-l]
factor =
r
Radius of soot particle
(ft ; m)
SZ
Secondary zone
Heat transferred by conduction across flame tube wall (Btul (s -ft2) ; W -m - 2)
T' g
Equivalent hot gas temperature
A constant used in swirler blade design
T.Q.
Traverse quality
Thermal conductivity W-m-1-K-l)
T
Temperature
t
Time (s ; s)
tw
Wall thickness
(ft;
U,u
Velocity
m-s-1)
u
Mean velocity over an area (ftl s ; m - s - 1)
V
Volume (ft3;
[Btul (s - ft - R) ;
(R; K)
(R; K)
.Any constant
L
Length
L
Length of combustion
Lu
Luminosity factor
m)
(ft ; m)
Radiation
(ft/s;
zone (ft ; m)
beam length (ft ; m)
m 3)
In
Logarithm to base e
V
Reference velocity
log
Logarithm to base 10
w
Slot gap width (ft ; m)
MC
Molecular
x
Film cooling correlation group (ftm-O.2)
weight of carbon kg -mol- 1)
(lbm/Tb - mole;
(ft/s;
m - s-1)
0
.2;
5-3
CHAPTER 5 x
Mass percent of sulphur in fuel
(02/Fuel)
x
Distance
Kinetic fuel loading [Ibm/' (s - ft 3atmn) ; kg-s-1m-3Pa-n]
x
Rate constant ratio
y
A parameter
a
Hole area ratio = Ah/ Aan
(ft ; m)
Angle of divergence between wall and axis (degrees; degrees (or rad) )
Swirler blade stagger angle (degrees; degrees (or rad)) Hole bleed ratio Psw
=
mh/man
Swirler - air turning degrees (or rad))
angle
(degrees;
w
Surface oxidation rate [lbm/' (ft2 . s) ; kgrn - 2s-1)
IT
Log log 1/7]
Difference
Subscripts
Pressure drop (Ibflft 2; Pa)
1
Overall Pa)
total
pressure loss
(Ibflft2 ;
Diffuser total pressure loss (Ibflft2 ; Pa)
act
Inner surface of flame tube wall
2
Compressor inlet, flame tube wall
outer
surface
3
Diffuser inlet, chamber inlet
of
Flame tube total pressure loss (Ibflft2 ; Pa)
4
Turbine inlet
a
Air, annulus
o
Momentum loss factor
an
Annu1us
€
Emissivity
c
Coolant,
€
Oxygen consumption efficiency
CO
Carbon monoxide
7]
Efficiency,
diff
Diffuser
7]c
Film cooling efficiency
e
8
Theta parameter [lbf1.75 -s/(lbm/ft°.75); kgO.75m-s-3.5]
Equivalent - usually refers to aerodynamic equivalents due to discharge coefficients
F, f
Fuel
Dynamic viscosity (lbm/' (ft - s}; Pa- s)
ft
Flame tube, combustor
Hole area ratio
g
Hot gas
p
Density
H
Flame tube-total
a
Stefan - Boltzmann constant [Btu/ (s . ft 2R 4); W· m - 2K - 4]
h
Hole
ha
Hydraulic diameter-annulus
T
Time, usually of residence (s ; s) he
Hydraulic diameter-coolant
HC
Hydrocarbons
usually of combustion
(lbm/ft 3; kg- m - 3)
Jet penetration grees (or rad))
angle
(degrees;
Equivalence ratio (02/Fuel) 5-4
cooling
de-
stoic :
Hydrogen
open hole area
film
COMBUSTOR r,i
Inner
n
Inlet
area. A number of examples of this basic conflict are as follows: •
Oxides of nitrogen Outer
(sometimes refers to origin)
Outlet
•
Primary zone Pertaining
to rich mixture •
ef
Reference, usually place of maximum cross - sectional area
tZ
Recirculation
zone
The mixing within the combustor can be increased to improve the uniformity of the exit temperature distribution at the expense of increasing either the pressure loss or the combustor length. Emissions of nitrogen oxides and smoke can be reduced by designing for a lean combustion zone. However, doing so results in decreased ignition performance, turndown ratio, and combustion efficiency. The frontal area of the combustor can be increased to improve combustion efficiency and flame stability, but this leads to a larger and heavier configuration which becomes more difficult to cool.
Snout t
Stoichiometric
w
Swirler
;Z
Secondary zone Sulfur dioxide Wall
'I
e
Pertaining
to weak mixtures
2.0 INTRODUCTION . I Background gas turbine combustor is a device for raismg ~e temperature of the incoming air stream by e addition and combustion of fuel. In serving .is purpose, the combustor must satisfy many ifferent requirements. It must be capable of iniating ignition easily and must operate stably ver a wide range of conditions. At all operating oints, it must provide for essentially complete bmbustion of the fuel while minimizing the forlation and emission of undesirable pollutants. o avoid damaging the turbine, sufficient mixing ust be achieved in the combustor to obtain an pceptably uniform exit gas temperature distrition. The combustor must also operate with as w a pressure loss as practical to maintain high O'er all cycle efficiency. Finally, all of these nctions must be performed in a configuration hich has a minimum size, weight, and cost, id which is sufficiently durable to achieve an :ceptable operating life. In many respects, these requirements are utually incompatible. Achieving an improveent in one aspect of performance very often quires a corresponding sacrifice in some other
These are typical considerations in the design and development of a combustor for any given application. Thus, achieving a successful combustor configuration involves trade-offs among the various relevant design and performance criteria until the optimum compromise has been reached, which best satisfies all of the imposed specifications and constraints. 2.2 Operating Conditions In the open-cycle gas turbine, fuel is burned continuously under pressure to heat air to moderate temperatures. The combustor is essentially a direct-fired air heater in which fuel is burned with less than one third of the air; the combustion products are then mixed with the remaining air to arrive at a specified temperature distribution at the turbine inlet. Combustor inlet temperature depends upon engine pressure ratio and load, varying from about 394 to 789 K (250 to 960 F) in various nonregenerative engines. With regeneration, combustor-inlet temperatures may be 644 to 1033 K (700 to 1400 F). Combustor outlet temperatures range from about 922 to 1255 K (1200 to 1800 F) for heavy-duty industrial turbines, and from 1061 to 1644 K (1450 to 2500 F) for aircraft-type engines. Engines utilizing combustor outlet temperatures in the range of 1644 to 1922 K (2500 to 3000 F) are under development. Combustor pressures for full-load operation vary from about 310 kPa (45 psia) for small, simple engines, to as much as 3.1 MPa (450 psia) in complex engines. The mass flow of air through a combustor varies as a function of load, but the static pressure varies similarly so that the volumetric air flow rate is nearly constant. This leads to the important concept that the gas turbine combustor operates with nearly constant air velocities at all loads, and the use 5-5
CHAPTER 5 of velocity is a highly useful combustor design criterion. Fuel rates vary with load, and wide ranges are encountered in aircraft applications, leading to the need for fuel atomizers having a flow range as great as 100: 1. However, the range of variation of fuel- air ratio is narrow (less than 3: 1) , which simplifies combustor design. Figure 1 shows the relationship between combustor pressure, combustor fuel- air ratio, temperature rise, altitude, engine speed, and thrust for steady-state operation of a turbojet engine with a pressure ratio of 5. It is evident that the combustor pressure varies widely with both engine speed, and altitude. The fuel- air ratio varies by a factor of 2 from idling to full load. Other turbojet engines would have similar characteristics, differing in combustor pressures and turbine rotational speeds. The effect of flight speed is to raise combustor pressure with increasing Mach number, displacing the curves of Fig. 1 upward. Another effect is to raise combustor inlet temperature without permitting an increase in outlet temperature, thereby reducing maximum temperature rise and fuel-air ratio. An industrial turbine of similar pressure ratio operating at fixed altitude, would operate similarly, 80r---------------,----,r----;
70
60 .~ a.
'" ~ ::J '"'"
.,
e 2
0..
'"
::J .c
50
40
30
E 0
U
20
10
0 400
800 Combustor I
0.008
1000
Temperature
!
I
!
0.012
Approximate
Fuel/Air
1200 Rise,
I
!
0.016
1400
F I
0020
Ratio
Figure 1. Relation of Combustor Temperature Rise and Fuel- Air Ratio to Pressure, Engine Speed, Altitude, and Thrust For a Turbojet Engine [96]
5-6
following an appropriate constant-altitude curve on Fig. 1. During transient operation, fuel- air ratios may vary significantly from steady-state ratios. On lightoff , and during acceleration, higher fuel-air ratios and greater jemperature rise are required. On deceleration, conditions may be appreciably leaner. A combustor which can operate over a wide range of mixtures without blowout simplifies engine controls, as scheduling of changes in fuel rate is less critical. Operating conditions for various types of engines are discussed in more detail in [2-13]. 2.3 Today' s Requirements The many applications of gas turbines include aircraft, automo biles, ships, hovercraft, electric generators, pipeline compressors, and pumps. The combustor designer has to solve the problems of stability, ignition, durability, exit temperature distribution, and cost. In addition, high priority is now given to the problems of emissions and alternate fuel utilization. In aircraft engine combustors, the aim is towards a reduction of pollutant emissions while using fuels with relaxed specifications. The relaxations will probably be brought about by the increasing scarcity of existing aircraft fuels within the next decade , forcing the blending of existing fuels and their gradual replacement with synthetic fuels. Oxides of nitrogen (NOx) will remain the most difficult pollutant to control. With the tendency for some aircraft to operate at higher altitudes (in the interests of fuel economy and probable traffic density), demands will be made for less and less NOx emissions. Attempts to reduce the NOx emissions have met with some success, using devices such as air blast atomization' two-stage combustors, etc. But NOx emissions are still considered to be excessive, and serious consideration is being given to the use of premixed, prevaporized combustors, with or without variable geometry. For marine purposes, as with aircraft, the fuels which will be used within the next decade will probably have some change in specification. An increased aromatic content is anticipated, and this may well result in higher wall temperatures, due to increased flame radiation and the possibility of carbon deposition, unless a modified design is achieved. For ground transportation, especially for military application, the trend is to introduce multifuel combustors. The restriction in the supplies of natural gas and low-ash distillate fuels has promoted interest in other possible fuels for stationary industrial gas turbines. These fuels are heavy distillates, crudes, residuals and blends, medium and low Btu gases and, in the longer term, dual systems
COMBUSTOR involving partial coal gasification. Along with this demand for multi-fuel application, there is a requirement (dictated by the need for improved efficiency) for higher turbine inlet temperatures. Often, the various conflicting requirements placed on the combustor can only be resolved by specialized designs. Nevertheless, all combustor designs have to meet the following basic requirements: 1. Maximum flame stability at all operating conditions 2. A high combustion efficiency at all conditions 3. Minimum pollutant formation at all conditions 4. Minimum pressure loss commensurate with operation and performance 5. Satisfactory ignition and relight at altitude, as well as ground starting at low temperatures 6. A satisfactory outlet temperature distribution tailored to the demands of the turbine 7. Absence of smoke and solids from the exhaust, as well as deposits in the combustor 8. Minimum manufacturing cost, size, and weight for the particular application 9. A long operating life 10. Ease of maintenance The priority given to each of the above requirements will vary with the intended engine application. For example, minimum size and weight are more important in an aircraft engine than in an industrial engine. Similarly, long life is more import ant in an industrial engine than in an aircraft engine for military applications. 2.4 The Combustion Chamber-First Principles Although the early combustors were derived by empirical techniques, it is possible to demonstrate the fundamental reasoning behind the development. The general approaches are presented in [14]. 2 .4 . 1 Combustion Zone In a combustor the flame is stabilized by recirculation of hot combustion products. This means that the fresh combustibles are vitiated by the returned products at the instant of ignition. Thus, the rate of combustion (and stability) are influenced by a tradeoff between temperature rise and vitiation necessitating a successful compromise. If one assumes an ignition temperature around 1500 K (2240 F) , an ignition mixture of about one part combustibles to 0.2 parts burnt products (i , e., a recirculation ratio of about 0.2) is implied. The highest laminar flame speed (at 300 K) of the fuel is about 0.4 mls (1.5 ft/s). Although flame speed increases as the inlet tem-
perature increases, it never approaches the mean gas velocity within the combustor. Turbulent flame speeds are somewhat higher, but not significantly so. Hence, there is a need for a stabilizing region which provides a sheltered wake for ignition and flame propagation. The simplest way of stabilizing a flame is to utilize a simple baffle, Fig. 2B. Its chief defect is that the only method for transfer of flow into the wake is by the diffusion of turbulent eddies across across the boundary streamline. A second defect is that it creates a series of shedding vortices, but this latter problem may be offset by the introduction of a secondary baffle, Fig. 2C. As indicated in Fig. 2D, the two baffles may be replaced by a series of holes which admit air in a way that promotes recirculation. In fact, this provides stronger recirculation although there is still instability due to fluid dynamic effects. In industrial boiler furnaces, a swirler is used to stabilize flames. If such a device is incorporated into the previous system, then the resulting recirculating flow pattern is strengthened and stabilized. The resulting configuration, Fig. 2E, closely resembles that of a conventional primary zone, Fig. 2F. The circular symmetry developed with the flame stabilizer lends itself to the use of a conical fuel spray. The cone angle is capable of being varied ovet a wide range, but experience has shown that, for most purposes, an angle between 90 and 120 is best suited to the chamber configuration. For the early engines a simple atomizer proved satisfactory. As operating conditions were widened, it became necessary 0
A
OitfuSMJn alone
C
Added Secondary Baffle
E
Holes & SWider
0
B
OiftUStOli
& Baftle
F Represent.lllve
PrlOl.lry
Zcoe
Figure 2. Evolution of Primary Zone
5-7
CHAPTER 5 to replace this with a dual injection system having a pilot flow for weak mixture, low fuel-flow operation, and a main high-flow injector for fullload operation. It is only recently that other designs of injectors have been shown to be beneficial to combustor operation. One of the major problems was to decide on the amount of air which should be admitted to the primary zone (PZ) . Initially, the air was added in stoichiometric amounts, but experience showed that better combustion occurred when the mixture was slightly lean (A/F = 18/1). However, this often resulted in flameout and ignition problems at high altitudes. To overcome the ignition problems the secondary zone (SZ) was created. This enabled the primary zone to operate fuelrich with good stability and light-up, while the secondary zone functioned as a region where the primary zone gases were diluted to a lean mixture for completing combustion prior to the dilution zone (DZ). The primary and secondary zones together comprised the combustion zone. In summary, the primary zone serves to: • •
•
Evaporate the fuel and diffuse it with the air to form a combustible mixture Provide a sheltered recirculation zone where the combustible mixture is ignited and stabilized by means of recirculation of some of the hot products Burn enough mixture to ensure nearly complete burnout in the secondary zone (the efficiency of the primary zone varies from about 40%at idle to about 80%at full load)
The function of the secondary zone is to take the hot gases leaving the primary zone and, by gradual admission of air, accomplish maximum combustion within a minimum distance. 2.4.2 Dilution Zone The function of the dilution zone is to reduce the bulk temperature of the gas to a level acceptable to the nozzle guide vanes and the turbines. This mixing of the hot and cold gases has to be accomplished in a minimum distance and with the lowest possible pressure loss. Despite much research, the understanding of dilution mixing is still incomplete. Most combustors still have their dilution regions designed on an empirical basis rather than on a fundamental design philosophy. The mean turbine inlet temperature is determined by the required life of the nozzle guide vanes and the turbine blades. Significant departures from this requirement can result in premature failure. Generally, the designer of the dilution zone is required to ensure a satisfactory radial and circumferential profile. The traverse quality (T. Q .) is often defined 5-8
as:
TO=~~~~_~~~E~r~~~re=Me~~_~~~perature Mean Temperature R1se x 100 For aircraft combustors values of T. Q. = 25% are generally acceptable; for industrial combustors T. Q. = 10%is a requirement. The peak temperature generally affects the life of the nozzle guide vanes. The turbine blades have their life more closely defined by the radial temperature distribution, since the rapid rotation of the turbine ensures that the blades "see" only a radial profile. As a rule, this profile must be tailored such that the turbine blades are cooler at the roots than the tips, thereby minimizing thermal stress. 2. 4 . 3 Wall Cooling The early gas turbine engines operated with low compressor exit temperature and low turbine inlet temperature. These conditions, together with a low combustor loading, meant that the combustor was not the life-limiting component of the engine. As the compression ratio increased and higher turbine inlet temperatures became possible, the combustor wall temperatures correspondingly increased and began to limit combustor life. To reduce combustor wall temperatures, film cooling was introduced. The earliest form guided cold air through a number of small drilled holes at a local hot spot. More general cooling requirements were met by introducing splash cooling devices, Fig. 3A, and later wigglestrips, Fig. 3B. The latter showed superior cooling performance and material strength [15] . In recent years, other devices, e .g. machined rings, Fig. 3C , and forced convective cooling have been introduced for longer combustor life.
Figure 3. Film Cooling Devices
COMBUSTOR The continual increase in combustor gas temperatures has demanded that more and more of the total combustor air be utilized for film cooling. This can result in some reduction in overall combustion performance due to:
• • •
•
An increased possibility of flame quenching near the combustor walls (resulting in increased pollutants) Less air available for effective dilution (resulting in high temperature distribution at the outlet) A reduction in primary zone performance
• •
Figure 4 identifies the various zones and subzones within the combustor, together with typical· flow patterns and air/fuel distributions. For any given combustor, both the sizes of the zones and the air/fuel distributions will depend on the function of the combustor and the preferences of the designer. One result of this is the proliferation of combustor types, which are only superficially dissimilar. 3.0 COMBUSTOR CONFIGURATIONS To define a combustor, it is necessary to specify three principal design features:
The inefficient use of the frontal area available for combustion (which could result in a longer combustor) The requirement for interconnectors for ignition purposes The likelihood of requiring a higher pressure loss
3 . 1.2 Annular Combustors Annular combustors appear to have all the advantages. First, they make use of the entire available combustion space (after allowance for the shaft). Second, with the use of axial discharge cornpressors , the discharge is already annular in form. Third, there is no need for transition prior to delivery at the turbine. One might also expect some savings in pressure loss, but this is generally negated due to the mismatch of the annular geometry with the fuel injectors. Thus, if a good exhaust traverse quality is demanded, then additional pressure loss is required to overcome the uneven pattern of gases leaving the primary zone. Sometimes attempts are made to overcome this mismatching by the use of fan sprays or similar devices. If the engine is large, and if it has a high compression ratio, the material thicknesses needed to overcome the strength problems may result in a system heavier than a corresponding tubo-system.
1. Geometric classification 2. Classification by air distribution 3. Classification by type of fuel injection
~ I :
However, even this triple identification is sometimes insufficient for a combustor for special applications.
COMBUSTION-t-
I:
I
*
**
I
0.3
**
*
= 0.35) to be burned effectively. Because of this, the combustor temperatures are kept at such a low level that the formation of NOx is negligible. 3 . 3 Combustors Classified by Fuel Injection 3.3.1 Liquid Fuels-Downstream Injection In the simplest form of atomizer, the fuel is pumped tangentially from the outside to the inside of the swirl chamber where it is deposited as a thin sheet on the walls. It then travels down a conical section to the swirler orifice where it is discharged at a very high velocity. At this point, the fuel has both axial and tangential velocities, and is ejected as a hollow conical sheet. The conical sheet rapidly thins and is disrupted into small filaments which then form droplets. The size of the droplets depends upon the velocity of injection (i . e., pressure loss through the atomizer) , the viscous shearing forces, and the surface tension of the fuel. This type of atomizer conforms (approximately) to the orifice law, and the flow rate is thus proportional to the square root of the atomizing pressure across the nozzle. The use of a simple atomizer demands a pump capable of efficient operation over a wide range of pressure and flow to cover the operating conditions of the gas turbine engine.
COMBUSTOR For early engines, the simple atomizer was sufficient, but as the range of operating conditions increased, a more sophisticated injection system was required. The first (and still widely used) technique was to construct two concentric nozzles. The inner one was small, and covered the low flows; the outer one was larger for higher flows. A difficulty of this type of atomizer is that somewhere along the operating line of the engine, there is a fuel flow which demands that the main nozzle operate with relatively low through flow (and hence very large droplets). Under these conditions there may be a considerable drop in combustion efficiency. In designing the combustor, care should be taken to ensure that such conditions are only transient. An alternative to this concentric nozzle type is a spill atomizer. Using this type, a larger quantity of fuel is discharged at a reasonably high pressure through tangential slots, so that at all conditions a high velocity exists. The bulk of this fuel is then removed through a hole in the rear of the swirl chamber and returned to the low pressure side of the fuel system. The small amount of fuel which is not returned is ejected with a high swirl velocity through the nozzle and into the combustor. A disadvantage of this system is that the fuel pump has to have large capacity at all operating conditions. Additionally, the atomizer characteristics seem to be less well defined than those of the previous system. 3.3.2 Slinger-Type Fuel Injection A sketch of this system developed by Turbomeca is given in Fig. 15. It necessitates the use of an annular combustor. The fuel is injected radially outwards from holes in the hollow shaft. The centrifugal action generates very high fuel pressures, even at low speed engine conditions such as idling, thereby ensuring good atomization over the entire operating range. However, since the strength of the hollow shaft is weakened by the fuel injection holes, it must be kept as short as possible. The chamber has its primary zone normal to the shaft, and the gases are directed through a right angle to the dilution zone. 3.3.3 Two-Fluid Atomizers It is possible to use a stream of high velocity air to assist in the breakup of the liquid sheet. Two general forms are recognized: air- assist and air blast. The former system uses a relatively high air pressure drop, while the latter uses the pressure drop available in the engine, usually across the combustor liner. When air-assist is used for only a portion of the operating range, say low power, it is usually called an air-boost system. The main function of the air atomizer is to improve the quality of atomization at low fuel flows. To achieve this, a small amount of air is ar-
ranged so that it blows across the fuel nozzle at about the same angle as the fuel spray cone. The quantity of air is about half that of the mass of the fuel at the low flow conditions. By contrast, the use of an airblast atomizer implies that the air is largely responsible for the atomization at all operating conditions. The amount of air to fuel ratio is usually between 4/1 to 5/1. Below the lower figure, atomization begins to decline; above the higher figure, there is generally little to be gained. Both types extend the flow range and may be used to atomize viscous fuels. Both also reduce the tendency toward amoke formation and the formation of oxides of nitrogen. They can generally reduce the latter by about a factor of two. 3 . 3.4 Vaporizers The term vaporizer is really a misnomer since at nearly all conditions there is insufficient heat transfer surface to heat and fully vaporize the fuel, even with the assistance of the inlet air which is admitted within the vaporizer tubes. Hence the device as a whole functions as a cross between a carburetor, an airblast atomizer, and a true vaporizer. A typical arrangement is illustrated in Fig. 16 [19]. The relative merits of spray system versus vaporizer have been argued for years. From a combustion viewpoint, there is probably little to choose. The main objection to vaporizers lies in the complexity of the system and the slow response to changes in operating conditions. They are also less amenable to the use of alternative fuels. 3 . 3 . 5 Gaseous Fuels Gaseous fuels are divided into three types: high, medium, and low calorific value. For the first type, the fuel injector is usually simple, and of a
__ A B C
o E f G
Cold Ai,
~Hot
Gas
Air enters from Compressor Out ••. casino HoUow nonle guidevanes (cold air mside) Primary lone Dilutionzone Fuel fed through holes in hollow shaft Hot gAs exit via turbine
Figure 15. Slinger Combustor Schematic
5-15
CHAPTER 5
COOLING
AI"
AI'-
GILLS
MIXING
BAFFLES
/
rut.L
~~;; ,. 'fE R~
SECO",,[,)AR1
AIR NOZZLE!>
""
INHRCO
•..•">IECT()"
Figure 16. Early Vaporizer Combustion Chamber (Courtesy
of Rolls-Royce
"pepper pot" type construction. If diffusion is a problem, as for example when hydrogen is the fuel (it burns so rapidly that the hot products remain "undiffused"), then it may be necessary to replace the pepper pot with a "spider" or multi -ring type of injector. Very wide burning limits sometimes require twin injectors. These are, similar to the dual orifice liquid fuel injectors. One of the major difficulties in designing gas nozzles is ensuring that, on shutdown, the residual gas cannot become mixed with air inside the injector to form an explosive mixture. The low calorific gases present a special problem in that their volume flow (to ensure good combustion) is much higher than high calorific gases, and is substantial in comparison with the air flow requirements. Consequently, the whole front end of the combustor has to be redesigned to ensure good distribution of fuel and air. The low calorific value gases usually have the advantage that flame temperatures are low and hence the oxides of nitrogen present no problem (unless there are nitrogen compounds in the gas) . 3 . 3 .6 Liquid - Gaseous System An industrial combustor is often required to operate with either liquid or gaseous fuels. This may generally be accomplished by using concentric injectors. Usually, the liquid injector is in the center, the gaseous fuel injector having the form of an annular pepper pot. If methane (natural gas) is the gaseous fuel, then there are substantial differences in the combustion properties of the two fuels, and careful attention must be paid to matching the flame tube to the gaseous fuel rather than the liquid one. Past experience has shown that a flame tube developed pri5-16
Limited)
[19]
marily for the liquid fuel may have only a comparatively short life when operated on the gaseous fuel. Although combustion appears to be satisfactory, it has been claimed that the difference in the frequency of the flame oscillation has brought about failure of the combustor walls. 4.0 THE DESIGN METHOD 4.1 Combustor Performance Performance of the gas turbine combustor is defined here to include all those characteristics which affect engine performance: combustion efficiency, pressure loss, outlet temperature profile, and limits of stable operation. Combustion efficiency is important because it influences specific fuel consumption; pressure loss affects both specific fuel consumption and power output and, thus, engine size and weight. Outlet temperature profile affects mean turbine inlet temperature and, thus, limits power output and efficiency. Stability limits and altitude limits define the operating limits of the engine which may limit its application. 4. 1. 1 Combustion Efficiency Combustion efficiency should always be close to 100%if fuel and air are well mixed in proper proportions, ignited, and given time to burn. In the usual industrial apparatus these conditions are easily met. In the gas turbine, and especially in turbojet engines, however, combustor size is critical, and it has proven advantageous to design for operation near the limits of combustion intensity. Furthermore, turbojet combustors operate over a wide range of altitudes, inlet temperatures, and fuel- air ratios, with the result
COMBUSTOR that combustion may deteriorate at high altitude and under lean conditions. A principal objective of combustor design and development has been to assure satisfactory performance up to the specified engine altitude limit, with minimum combustor volume and pressure drop. , A great deal of systematic testing of turbojet combustors has been carried out to correlate combustion efficiency with design variables and operating conditions. Much of the significant results are reported in [2], [20], [21] and [22]. Combustion efficiency in typical combustors has been found to be adversely affected by high reference velocity (small combustion volume), low inlet air temperature, and low pressure. In an annular combustor for a turbojet of early design, combustion efficiency fell rapidly at reference .velocities above 28 mls (92 fps) , inlet temperature below 290 K (65 F), and pressure below 82.7 kPa (12 psia). One way to correlate combustion efficiency over a range of operating conditions is to plot it against PT IV where P is absolute pressure, T is combustion inlet temperature, and V is reference velocity. Combustion efficiency has also been found to vary considerably with fuel type [23] . However, much of this variation can be traced to differences in atomization resulting from differences in viscosity, and to differences in evaporation rate resulting from differences in volatility. It appears possible to obtain good efficiency with .any usual fuel by close attention to fuel atomization and dispersion. 4 . 1. 2 Pressure Drop Pressure drop in combustion systems is usually defined as the difference in total pressure between the compressor outlet and the turbine inlet. It consists of three components: the diffusion loss associated with slowing the high velocity air from the compressor outlet, the friction loss taken as pressure drop through the combustion chamber if unheated, and the momentum loss associated with accelerating the inlet, low temperature air to a higher exit velocity. Representative values of these losses are 40% for diffusion loss, 40% for friction loss, and 20% for momentum loss. Total pressure loss is usually in the range of 2 to 8% of static pressure. This loss has the same effect as a decrease in compressor efficiency; it results in both lower power output and higher specific fuel consumption. Pressure loss of 1% results in a loss of about 1%in output power and a 1%increase in specific fuel consumption, depending upon the engine cycle. Although there are several ways of expressing and correlating combustor pressure drop, the only expression which has significance relative to engine performance is percentage of static .pressure. Ef-
fects of pressure drop on engine performance are discussed in [2], [3] and [4]. 4 . I . 3 Temperature Profile The average gas temperature level for satisfactory turbine life is limited by the peak gas temperature, so that a large temperature gradient reduces the average gas temperature and, thus, limits turbine output and efficiency. For purposes of turbine design, it is necessary to provide for a reasonable variation of gas temperature at the turbine nozzles in both radial and circumferential directions. Turbine nozzles are directly exposed t.o local gas temperatures, and the highest local gas temperature can affect the life of the nozzle on which it impinges. Thus, the highest local gas temperature limits. nozzle design. Turbine buckets, however, rotate past the entire nozzle ring, and effectively average out circumferential temperature gradients. For bucket design, the average radial temperature profile, as compared with a stress-limited theoretical radial profile, is of interest. However, large circumferential temperature differences must be avoided to minimize excitation of bucket vibration by differences in gas velocity. Temperature profiles are frequently described, for turbine design purposes, in terms of a "Temperature Factor" or a "Traverse number." For nozzle design, the Traverse number would be defined as peak gas temperature minus mean gas temperature divided by mean temperature rise. For bucket design, the Traverse number would be the difference between the highest average radial temperature and the mean. radial temperature, based on either a flat radial profile or a tilted profile with the higher temperature at the bucket tip. The usual range of Traverse number is between 0.05 and O. 30, and is the result of optimizing all cycle variables. For example, it is possible to improve turbine output and efficiency by improving Traverse number, but this improvement requires greater combustor pressure drop and length, so that a limiting point is soon reached where loss overbalances gain. The Traverse number may grow larger with turbine age if the combustor structure is such that areas of various film -cooling slots, joints, or other gaps can change as the result of warping or other deterioration. Thus, the effect of age on the combustor is an important aspect of selecting a realistic temperature factor for the life of the gas turbine. Good mixing, as indicated by a good Traverse number, is most important for operation at full load with maximum turbine inlet temperature, since turbine life can vary significantly with small differences in local metal temperatures. Larger variations can be tolerated at lower levels of turbine inlet temperature. 5-17
CHAPTER 5 4.1.4 Stability Limits Combustor stability limits are the limits of fuelair ratio, or temperature rise, within which a combustor can operate. If air flow through a combustor is constant and fuel flow is varied, rich blowout will occur at some high fuel rate and lean blowout will occur at some low fuel rate. The fuel-air ratios at which blowout occurs must be relatively far from steady-state operating conditions to permit engine acceleration and deceleration. For example, a step change from full load to idling will reduce fuel flow by about 70%, but the corresponding change in air flow will take place gradually as turbine speed changes. The instantaneous combustor temperature rise will also decrease 70%, then gradually reach equilibrium as air flow drops, Likewise, acceleration from idling to full power requires momentary operation with a temperature rise well above that for steady-state operation, Fortunately, most combustors have an extremely wide operating range and are hard to blowout by normal changes in fuel flow, However, at high altitudes the operating range of turbojet combustors narrows, so that the useful power range of the engine is narrowed, Combustor blowout can also occur as the result of compressor surge, During acceleration, a large increase in fuel rate without a corresponding increase in air flow will reduce flow through the compressor. If the compressor goes' into surge, the air flow will be reduced still more which may cause rich blowout of the combustor or excessive turbine inlet temperatures, Many turbojet engine control systems schedule the rate of change of fuel flow with altitude to avoid compressor surge or blowout during accelerations and decelerations. 4 .1. 5 Altitude Limits The altitude limit of a turbojet engine is determined primarily by combustor limitations, Although Reynolds number effects result in some loss in compressor and turbine efficiency, this is minor compared to the large loss in combustor efficiency which takes place as the altitude limit is approached, With increasing altitude, the rich and lean blowout limits converge gradually, narrowing the usable thrust range of the engine, Combustion efficiency also declines gradually, and. at the altitude limit, becomes so low that the turbine inlet gas temperature needed for continued operation cannot be generated. The loss of efficiency at high altitude occurs because the volume required for combustion becomes larger at extremely low pressures, Consequently, combustion reactions are only partially completed by the time the burning mixture is quenched in the dilution zone [2], [20-22], Thus, the altitude specification for the engine 5-18
determines, to a large extent, the reference v locity or combustion volume required in U combustor. Since altitude limit is primarily lower limit of combustor pressure, engines ha ing very high pressure ratios also have superi altitude performance, Poor fuel atomization may also contribute poor combustion at high altitude, Fuel rates which are proportional to ambient pressure, ai only 7 to 10% of those at sea level, and fu pressure may be too low for satisfactory atomiz tion. It is especially important to size the primal and secondary orifices of a dual- orifice nozz so that the flow-divider opening pressure do not fall within the high altitude operating regim
4,2 Design Specifications Hitherto, it was often acceptable to select a si gle operating condition as the design point for t combustor and to accept the performance oli tained at the other conditions, Since, generall the combustion efficiency was high at all co ditions other than idling, this did not significant detract from the overall engine performanc With the added requirement of minimum poll tion, this technique is no longer acceptable, a many conditions must be examined in order t ensure satisfactory performance from all view: points, Currently, the status of combustor desi does not impose significant performance re straints upon the engine. However, dependi upon the application, there are still some per formance requirements which demand greate attention than others. Attention to these detail can assist in defining the type of combustor an fuel system required, Next it is necessary to re fine the information required to specify the type size, and design of the combustor, Because 0 the emphasis on controlling air pollution, th variables within the combustor have to be can sidered for as many operating conditions as pas, sible . The items to be specified at each condi tion are: • • • • • • • •
Air mass flows Fuel mass flows Combustor inlet conditions of temperature, pressure and velocity distribution Turbine inlet temperature and traverse quality Pressure loss limitations Combustion efficiency limitations Pollutant types and permitted levels Allowable wall temperature
The design engineer has to decide how manj of the operating conditions are relevant, For example, in considering a possible aircraft application, the following represents the minimum
COMBUSTOR conditions which require investigation: • • • • • • •
Maximum thrust Normal cruise Maximum altitude Windmilling conditions at maximum altitude for relight Ground idling Approach cruise
In addition to specifying the above performance variables, it is also necessary, at this stage, to specify the physical limitations within which the combustor designer will have to operate. These are: • • • • •
Space limitations Weight limitations Total life requirements and life between overhauls Fuel types Any other special restrictions or requirements, for example, compressor speed limitations at lightup , any thrust augmentation, or special acceleration requirements
4.3 Design Sequence With the information outlined above, the designer is now in a position to decide on the type, size, and design features of the combustor needed to meet the specifications. Before attempting a detailed design, there exists simple empirical formulae which will enable the designer to arrive at the overall and zone sizes, the air/fuel distribution, and the film cooling requirements of the combustor. A preliminary approach is as follows: •
Select the combustor type. This will be defined by the engine application and space limitations. • Specify the casing and combustor cross- sections. These may be limited by the available engine space. If, so the engine designer should be warned of any possible performance losses. • Specify the fuel system. Again, this is largely defined by the engine application (and, possibly, by pollutant restrictions) . • Specify the primary zone dimensions and operating air/fuel ratios. • Specify primary zone film cooling air requirements. Sometimes it is better to define all the film cooling requirements after establishing the combustor zone requirements. • Specify secondary zone size and air requirements '. • Specify secondary zone film cooling requirements, • Specify dilution zone size and air require-
• •
ments. Specify dilution zone film cooling requirements. Specify ignition requirements and type of system. Specify interconnectors, etc.
The complete design of a combustor involves the following steps: • • • • • •
Preliminary design and sizmg Evaluation of preliminary design at all operating conditions Modifications to preliminary design to meet all operating conditions Detailed design Evaluation of detailed design at all operating conditions Modifications to detailed design to meet all operating conditions
On completing the above steps, the design is ready for hardware procurement and the start of the development cycle. To illustrate the method, a hypothetical example will be given. This design represents neither a currently used combustor nor a future design. The selected conditions are not applicable to any real engine; they are used here for illustration only. It will be seen that, even with the simple design formulae proposed here, an iterative procedure is required in order to achieve a combustor which represents the best compromise betweenaerodynamic and combustion requirements. Basically, the method for both aircraft and industrial combustors is to utilize a series of simple correlation formulae to achieve a "first-design." It is then supposed that a limited amount of test rig development will be carried out. At this point, the original correlation formulae may be modified to enable firmer predictions of performance to be made at operating conditions other than those tested. If required, experimental information may be used to obtain more accurate predictions. 5.0 SELECTION OF COMBUSTOR TYPE 5 . 1 Introduction Currently, there is no known method for designing a combustor that will guarantee that no development will be required. The aim of the present technique is simply to minimize the development time. The logical arrangement of the major steps is indicated in Fig. 17. The selection of combustor type is usually predetermined by the overall engine specific a tions but, in cases of doubt, it may be necessary to conduct a design study for two or more types of combustors. The essential points of the major 5-19
CHAPTER 5 types of designs are summarized here.
hp or 3.4xl06
5 .2 Aircraft Systems The combustor space of an aircraft engine is the annular volume between the compressor and the turbine, with the inner dimensions limited by the shaft and bearings and the outer dimensions by the diameters of the turbine and/or the compressor casing. Although there are no specific restrictions on the choice of combustor type, over the years certain trends have evolved. In the older engines, multi-can combustors were generally employed, especially for those engines having compression ratios up to about 15/1. While they do not fully use the maximum available combustion area, can-annular combustors offer some of the advantages of both annular and multi-can units. They are currently used in some medium sized aircraft engines. The current trend is toward annular combustors. Generally, straight-through combustors are used for the larger engines; for helicopter engines, because of length limitations, reverseflow combustors are favored. For air pollution purposes, EPA classifies gas turbine aircraft engines as follows (Federal Register, Vol. 43, No. 58, March 24, 1978):
Large units, over 15 Mw (over 20,000 hp 5lxl06 Btu/h)
"Class P2" means all aircraft turboprop gines.
en-
"Class Tl" means all aircraft turbofan or turbojet engines except engines of Class T5 of rated power less than 35,600 Newtons thrust. "Class T2" means all turbofan or turbojet aircraft engines except engines of Class T3, T4 and T5 of rated power of 35,600 Newtons thrust or greater. "Class T3" means all aircraft gas turbine engines of the JT3D model family. "Class T4" means all aircraft gas turbine engines of the JT8D model family. "Class T5" means all gas turbine engines employed for propulsion of aircraft designed to operate at supersonic flight speeds. 5 . 3 Industrial and Other Engines It is convenient to classify non-aviation gas turbines into three categories by power output. For the purposes of this report, the division is arbitrarily defined as follows: Small units, less than 1 Mw (appro x . 1340 hp or 3. 4xl06 Btu/h) Medium units, 1 Mw to 15 Mw (1340 to 20,000
to 5lxl06
Btu/h) 0
5 . 3 . 1 Small Units (Includes Automotive and APU Applications) Since the frontal area of the combustor for sma units is not as important as in aircraft practice the size of the combustor is often determined b the limitations of the fuel system. In order t avoid difficulties in manufacturing very small at omizers (which may block in use), the size 0 the combustor is often set by prior choice of th smallest convenient and most economical atom izer. Hence, single flame tubes or reverse -flo systems are quite normal especially for engine of about 0.2 Mw or less. Annular combustors also find application for small engines, but fO very narrow annuli some loss in performance may occur, and there may be difficulties with filn cooling. If pollution legislation is applied to smal engines, it may become necessary to use the single combustor concept due to the quenchinj action of the walls of a narrow annular combus· tor. Currently, it is not known at what level war quenching becomes significant, but evidence il accumulating that it is of the order of 25 to 4( mm gap width (about 1.0 to 1.5 in).
5.3.2 Medium and Large Units In this category there is considerable choice, When designing a combustor specifically for industrial application, the usual choice is betweer multi -can and single can systems. However, annular combustors, originally designed for aircraft use, have been found to be satisfactory fo industrial application after suitable modification. Primarily, the selection between multi-can 0: single can combustors depends on the econom· ics of manufacturing and development. Where space is limited, the multi -can systerr is usually employed, often with reverse flow ir order to minimize the length of the engine shaft. Development costs are also reduced, since ( single combustor may be tested using only ( fraction of the main air flow. However, manufacturing costs are higher, accessibility for repair: and maintenance is reduced, and a more com· plex fuel system is required. Of all the combustors, the single can is probably cheapest to manufacture. The use of a single large atomizer is particularly advantageou: with heavy high -viscosity fuels, which require preheating in the lines. Development of single cans is often accom· plished using scaled models, but particular can is required in scaling heat transfer results, anc final development is required on the full scak combustor. The single can combustor has founc
COMBUSTOR
Design specifications (All operating conditions).
Select combustor type. Provisionally select fuel injector. Determine reference area and other reference values. Determine combustor area.
Determine Primary Zone air requirements and size.
Select ignition system.
Determine Secondary Zone air requirements and size.
Estimate total film cooling air.
Estimate Dilution Zone air and size.
Design combustor holes.
Est. pressure loss due to combustion.
Figure 17. Preliminary Design Procedure
5-21
CHAPTER 5 application in all the ranges of industrial engines and, because of its favorable cost and accessibility, it is particularly favored for very large units, for units which have to be adapted for multi-fuel application, or for dual-fuel use (e. g . liquid/gas). For the latter two applications, the ease of interchanging fuel injection systems is particularly advantageous. 6.0 THE FUEL INJECTION
SYSTEM
6 . I Introduction To achieve rapid burning of a liquid fuel, it must be thoroughly mixed with the air in roughly stoichiometric proportions. However, before mixing and combustion can proceed, the fuel must be vaporized With the vaporizing system type of fuel injector, the fuel is mixed with some air and vaporized before admission to the combustion zone. In other types, the fuel is atomized into a large number of small drops, thereby producing a liquid spray of high surface-to-volume ratio and hence a high rate of fuel evaporation. Both methods of fuel injection have found widespread application in aircraft engines, but the current trend is towards a greater use of atomizers in one form or another. Considerable experience has now been gained on fuel atomizers in both aircraft andIndustrial engines. For conventional combustors, the fuel spray should contain a reasonable proportion of drops in the size range below 50 micrometers. These small drops of high specific surface area readily evaporate and burn to provide a source of high temperature products that initiate and sustain combustion of the spray. In general, fuel/air mists composed of drops less than 20 micrometers in diameter behave as perfectly uniform mixtures. To realize the high rates of heat release associated with uniform mixtures, it might be argued that the spray should be composed entirely of finely atomized drops. Unfortunately, uniform mixtures are also characterized by a very narrow burning range. In practice, therefore, some imperfections in fuel/air mixing may be desirable in order to attain stable burning over a wide range of mixture strengths. The atomizers used in gas turbines are often of the pressure swirl type, in which fuel is Iorced through tangential ports into a swirl chamber from which it is discharged in the form of a conical sheet. With the early "simplex" types of swirl atomizer used on aircraft gas turbines, stable burning was possible only over a fairly narrow range fuel flow. Various limitations on performance were experienced, notably in terms of low combustion efficiency and inadequate relighting capability at extreme altitudes. To eliminate the shortcomings of the "simplex" atomizer, various "wide -range" atomizers were developed,
0'
5-22
outstanding examples of which were probably the so-called "duple" or "dual-orifice," and "duplex" spray atomizers. These developments, along with parallel advances in the performance of vaporizing systems, provided a period of many years in which fuel injection was one of the most satisfactory and trouble-free aspects of combustor operation. However, with the continuing trend towards engines of higher compression ratio, combustors using pressure atomizers increasingly suffered from excessive exhaust smoke and poor pattern factor. Vaporizing systems, on the other hand, began to pose serious mechanical problems due to overheating and coking of the vaporizing tubes. These problems led to a revival of interest in the airblast atomizer, in which atomization is achieved by injecting the fuel into a high velocity airstream. This type of fuel injector is ideally suited for gas turbine applications, since high velocity air is always available due to the pressure drop across the liner. An important feature of the airblast atomizer is that it provides an element of fuel/air mixing prior to combustion, which leads to worthwhile reductions in exhaust smoke and in the emissions of oxides of nitrogen. INJECTOR
REQUIREMENTS
Ideally, a fuel injector should possess all of the following characteristics: •
Good atomization over the entire range of fuel flow • Rapid response to changes in thrust or load • Low cost, light weight, ease of manufacture and ease of removal for servicing • Freedom from flow instabilities • Low susceptibility to blockage by contaminants and low carbon buildup on tip • Low susceptibility to gum formation by heat soakage • Be capable of scaling to provide design flexibility The injector should also provide: • • • •
An easily ignitable mixture A ratio of maximum to minimum fuel flow that exceeds the ratio of maximum to minimum engine air flow Rapid evaporation and dispersion of the fuel throughout the primary combustion zone An exhaust gas temperature distribution that is insensitive to the amount of fuel supplied
6 . 2 Selection of Fuel Injection System At this stage it is useful to consider the likely fuel injection system for the combustor. For liquid
COMBUSTOR fuels the choice is large, ranging from a simple pressure atomizer, to the complicated system required for a premixed, prevaporized system. For most modern combustors it is necessary to utilize some form of airblast or air-assist atomizers in order to comply with environmental restrictions on HC and CO emissions. Although the choice may be made at this stage, the initial calculations of combustor size will ignore the effects of the fuel injector. Little difficulty is anticipated when high Btu gaseous fuels are used, but for low Btu gases, the fuel injector must be accepted as an integral part of the combustor, since it significantly affects the design of the head of the combustor. 7.0 COMBUSTOR
DIMENSIONS
7.1 Determination of Reference Area The reference area, Aref, is selected by considering the possibility of either chemical or pressure loss limitations. Each operating condition is considered and, from the tabulated answers, the optimum value is assessed. The technique will be illustrated with a simple example. It applies equally to aircraft and industrial combustors. The calculations are made even for combustors where there is no outer casing (e. g. an industrial combustor with regenerator) to estimate the combustor liner dimensions. (This is based upon the assumption that Aft = 0.7 Aref) . The data for the example are given in Table 1. They have been chosen for the purely hypothetical case of a multi-can combustor. The data are for six single cans. The engine bears no relation to any known combustor and is not necessarily typical. This particular engine also contains fewer operating conditions than a real case. The aircraft operates at Mach 1.4 condition, and an exit velocity of 150 ms-1 has been assumed for the compressor, which also has an exit cross section area of O. 096 m 2. The fuel is assumed to be a typical aviation kerosene. The permissible pressure loss for this combustor is higher than usual. Table 2 lists representative values of pressure loss terms for aircraft combustors. 7 . 1. 1 Aerodynamic Considerations The dimensions of a combustor might be determined either by aerodynamic or by reaction rate control. Generally, if the combustor is sized to meet a specific pressure loss, it will be large enough to accommodate the chemical reactions. However, it is necessary to consider all the possible factors before making a final choice of size. The combustor and casing diameters are estimated using the equations which follow. The values obtained, see Tables 3 and 4, are then assessed against the value derived from Eq , (1) below. This latter has been fixed by the
specification given in Table 1 and is repeated here only for completeness. Equation (1) is used to calculate the value of (m -If 3) / Aref p 3 in Table 2.
A
ref ( 1)
The constant, k , has the value of 143.5 in SI units and 0.83 in lb ft R units. Based upon the data of Table 1, the corresponding diameters are given in Table 3. Sometimes, at one or more conditions, the diameter corresponding to one or another of the reaction rate equations exceeds that determined by aerodynamic considerations. When this occurs, it calls for a judgment by the combustion engineer as to which is more correct. 7.1.2
Chemical (Combustion) Considerations For any given fuel/air ratio, the combustion efficiency, 1}, is given as a function of the correlating parameter, (), [24] where:
o
P 1.75 A nO.75 exp(T /b) 3 ref 3 -------------------------m3
(2)
All combustors have combustion efficiencies close to 100%at a value of () = 73 X 106 (SI units), Fig. 18. Values of the temperature correction factor, b , for a constant overall air/fuel ratio have been defined by Eq , (3), [25]. b = 245 (1. 39 + .tn pZ) for 0.6 b
<
PZ< 1.0
(Kelvin)
170 (2.00 - .tn pZ) for 1. 0
<
PZ< 1. 4
(Kelvin) (3)
To obtain the appropriate value of PZ(the primary zone equivalence ratio), the following simple analysis is performed. For most hydrocarbon gases burning ~in air, examination of the determined flame limit temperatures for weak mixtures (assuming no quenching at the wall) shows that they lie around 1600 K. Pressure is assumed to have no effect upon these limits. Thus, the limiting equivalence 5-23
CHAPTER 5
Table 1. Hypothetical
Operating Conditions for a Fictitious
Combustor
The engine is assumed to be a multi-can configuration having six combustors. cruise
Mach
number
is assumed
as 1.4;
the
compressor
exit
has
art
area
The aircraft's m2,
of 0.096
and
the
exit velocity is 150 ms+ I at the normal cruise .condition. Pressure loss characteristics the design point are assumed to be:
llP _ 3 4
llP _ 3 4
7%
~ Condition
Condo
p)
p)
MPa
MPa
Design Point Thrust
5
Ground Idle
llP _ 3 4
Turbine Inlet Data
~
T3
m3
T4
K
-1 kg. s
K
~
Q % Pattern
overall
Fac tor
Comb. Eft. % min
m
f
kg.s
-1
1.93
814
18.1
1600
0.347
20
99.7
0.427
0.070
0.7
0.68
707
6.8
1387
0.286
20
99.5
0.132
0.070
308
0.22
20
99.5
0.140
0.070
99.0
0.0091
0.017
Relight Alt. 0.030 0.0298 Windmi 11ing Normal
-3
2.0
Max.
Note:
P ref 3
SLS
Max. Altitude
4
).0 x 10
A
Compressor Outlet Data for a single combustor
Number
Max.
mvT3
53
qref
at
1.8
1.77
1060
14.2
1393
0.145
0.15
0.148
343
1.05
703
0.128
Cruise
One condition is selected as preemptive and referred to as the 'Design Point'. this is Max. Thrust SLS (as chosen he~e).
Usually
",
Table 3. Outer Casing Airflow Reference Values
Table 2. Representative Values of Pressure- Loss Terms for Aircraft Engine Combustors
OUTER
Chamber Type
llP _ 3 4 qref
llPJ_4 PJ
%
5.3
Multi-can
Annular
40
6.0
Can annular
llP _ J 4 P3
5.4
143.5
ref PJ
30
3.5 x 10-J
(~y
AIRFLOW
REFERENCE
VALUES
0 ---1&
A
~ Eqn 1 m
2
Eqn 2 m
Eqn
2
1
Eqn 2
m
m
0.335
0.076
1.
2 8.81 x 10-
4.54
2.
-2 8.76 x 10
-2 1.47 x 10
0.334
0.137
3.
-2 4.26 x 10
-1 1.69 x 10
0.233
0.464
4.
-2 8.605 x 10
9.079 x 10-4
0.331
0.034
5.
-2 4.26 x 10
0.234
0.201
x 10-3
-2
3.173 x 10
ref PJ
ratio (weak mixture) corresponding to any condition will be given by a mixture having a temperature rise of (1600 - T 3)' Let 1600 - T 3 = t.T. Using Fig. 19, the value of t.T is read against the appropriate curve corresponding to the inlet temperature (T 3)' The corresponding value of 4> yields the equivalence ratio for the weak extinction .' The corresponding rich extinction value is obtained in a similar way using Fig. 20. This particular figure has been chosen to minimize the possible effects of chemical dissociation. 5-24
Condo No.
J.O x 10-J 4.5 x 10-3
x
qref
A
20
llP3_4
CASING
m/f3_
Table 5 is derived in this way. From Table 5, if the primary zone is designed to use between 14 and 24%of the total air, then the combustor should remain lit at the four major operating conditions. To design for minimum smoke, CO, and HC, it is desirable that the primary zone equivalence ratio never be richer than 4> = I. 5 at any single operating condition. ¢ = I.5 corresponds to 23% of the air entering the primary zone at Condition 1. Since this is the richest condition (i . e., probability of maximum
COMBUSTOR
Table 4. Combustor Liner Airflow Reference Values 2
d
Aft m
m
ft
Condo No. __ Eqn 1 & 4 1
6.17xlO
2
6.13xlO
3
2.98xlO
4
6.02xlO
5
2.98xlO
Eqn 2 -2
-3
3.18xlO
-2
-1
1.18xlO
-2
6.36xlO
-2
3.73xlO
-2
1.03xlO
-2
Egn 5
4
&
-4
2.221xlO
-2
Eqn 7 -2
6.88xlO
NA
4.38xlO
NA
2.35xlO
NA
2.41xlO
NA
1.81xlO
Note - Selected value of d
ft
1. 0
Eqn
Eqn 1 & 4 -2 -2 -1 -2 -1
2.80xlO 2.79xlO 1.95xlO 2.77xlO 1.95xlO
2.80 x 10
-1
2
&
-,
-1
6.36xlO -
-1
1.15xlO
-1
3.88xlO
-1
2.85xlO
-1
1.68xlO
Eqn 7
Eqn 5
4
2.18xlO
-1 -1 -1 -1
-1-
2.96xlO
NA
2.50xlO
NA
5.47xlO
NA
1. 75xlO
NA
4;80xlO
-1 -1 -1 -1 -1
m.
obtain Aref for each operating condition. For comparison with the earlier aerodynamic values, the results are indicated in Table 3. It will be noted that there are considerable differences in the estimated values for each condition.
,....----.---r--,-----r---,
7.2 Determination of Combustor Area There are several methods which yield the combustor area, Aft. A simple relationship is: Aftl Aref
2U
iu
30
Theta l bm , l bm ,
-**
Sl
lbf, l bf ,
40 t
ft, in,
R units K «n t c s
s, s ,
so
50
Pu rame
e r- -
70
HOxIOb
•• _
IJ
units
Figure 18. Theta Parameter Correlation smoke formation), it is recommended that this value (23% of the total air) be the minimum used. Since some form of fuel spray injection 'is envisaged, the real limits of combustion will be considerably wider than those determined here, and hence it will probably be safe to use an even weaker mixture at take-off. For initial design purposes, a split of 25%will be assumed. Table 6 indicates the predictions based upon this split. Using the value of b determined in this manner, and making use of the fact that Ar~f = (17D2)/4, it is now possible to solve Eq , (2) to
=
(4)
O. 7
This seems to be quite satisfactory for single can, multi-can, and annular combustors, but for can-annular combustors, a value of about 0.65-0.67 is more appropriate, since it becomes necessary to "draw-in" the combustors. Using Eq , (4), the data of Table 3 have been used to derive the equivalent combustor data of Table 4. It is recommended that the following techniques also be used to estimate the combust or dimensions. The value finally selected will then be made on the basis of engineering judgment. An alternative method of predicting combustor size was determined by Bragg [26]:
-2
1.621 x 10
m f
ff
P
0.5
3 3 -----(--) P 6P
r
(5)
The combustor diameters calculated using the above formulae are listed in Table 4. Yet another possibility is to utilize the OdgersCarrier equation [27]. Substituting a value of 5-25
CHAPTER 5 Table 5. Equivalence Ratio Limits for Sample Combustor Condo No.
T3 deg
6T = 1600-T deg K 3
K
Theoret ieal limits (¢)
Operating
Ratio
(¢)
¢Overall/ ¢Limit
Ovrnll wea.k
rich
weak
rich
l.
814
786
0.345
2.52
o , 347
1.
2.
707
891
0.391
2.38
0.286
0.7J
0.12
3.
308
1292
0.550
2.01
4.
1060
540
0.240
2.68
0.145
0.60
0.05
5.
343
1257
0.525
7.17
0.128
0.24 **
0.06
zone will operating
en su r«
* **
14 per cent of the air in the primary combustor will remain lit at all rich
oi
0.14*
the cnodittoo.
24 per cent of the air in the primary zone will ensure the combustor will remain lit at all weak operating condition.
T3 -
300
550
200 300 400 500 600 700 800 900 1000
AT K
500 T3 100 4 00
fJ
200
/,
AT K
T3
200 300 400 500 600 700
1000
K
6 00 8 00 1000 K
1100
900
800 900 1000
400
K
A 'I
~Z 1'/ V 100
700
i
3UO
500.~t=
500 200
.,HH-+-+-+-+++-+-H
o
400
0.02
0.04
0.06
~
0.08
0.08
0.1
0.2
0.15
0.3
0.2
0.4
Equivalence UNDISSOClATED (all
TEMPERATURE
RISE
pressures)
Figure 19. Undissociated Temperature 99% for the combustion efficiency, and making the assumption that the primary zone volume is given by:
v then, 5-26
=
(1714)
d;t
with 11 = 0.99,
Ratio
Rise
duces (in SI units) to: log "300 = -1. 39 - 4.40n
(6)
the original equation re-
-1. 10 D* (7)
where n (reaction order)
2 ¢J (weak) or 2/¢J
COMBUSTOR
2~00
2000
1800
1400
.... :::1::' ........... 10')0
,'".
'''1: ' ::+.
"::'::~ ,:"i:: ';:L:
iliSJ ,.•..
..... '•• :F
"""".r
,;
~'I.~ "': ::' ''I
.•,:1:,:::::; :'~f....,,' ::.::·:::Y
.. .... :".·'1
.••'::V'"
::.::: :~ ;~-;i-~~:; :::::::: .... :: ::::j:::: :~~:g:'::I;::
0.4
0.5
0.6
I::
§33:::;"~:
=
3 MPa
- ='~·:~I:::
(29.6 atm)
E5=t~:
. . . .. . . ..
=:. "" ~,~""
,: .::j' :: :':r::: ::: ::=:1 --: ':::t:::-!
E "'" :c:F: .::=:-:
.......• :.:1":1: I.c.LjeJ J':J::vrLLiJ::I:,
0.7
0.8
0.9
Equivalence
Ratio
j~ :-:~~~._
EEF.,-s'~
....;ii:I::~l'1.:l ••• :Ft~l"¥'FLTTl'.TTcFoTc.r
.................J'
J :::;:i:J.
0.3
"'11
RISE
(Chemical Equilibrium)
P3
::::i
600 I
.
TEMPERATURE
..........• :::1'''::':L:
>i:,:: 'I::', ::~:::,: :·T:' :::'i:: :::J :: >L
I ::':, ::::'1':1
.
- ADIABATIC
::'::',,' "'":,:,::~,::, ,
.
:J :;;':.1 ,.::'
81)0
...........
.
1.0
1.1
1.2
-!:: ::.+::-
JCJ~~:::;,::~~~
14, ..:J:.:
I
1.3
1.4
1.5
Figure 20. Adiabatic Temperature Rise Table 6. Initial Combustor Air Distribution Overall
Condo
No.
Primary Zone
b (Eqn
3)
Predicted Premixed gas Extinction
Weak
o A/F
o A/F
log 7/1'300 is corrected to 7/l'T using the following equation, [28]. 3
Rich
o A/F
(T 3
0.347 42.36
1.39 10.59
272
170
23.3
where:
0.286 51.40
1.14 12.85
318
206
24.7
y
0.145 101.38
0.58 25.35
207
245
22.0
0.128 114.84
0.51 28.71
176
112
28.7
o A/F
17/1' T3
.
=
205
(1O-3.054y-1.
)
300 1.2327y-1.205
= 4>
Finally,
A/F
o
~
(9)
)
for
4>
,,1.0;
y
=
1.0 for
4>
> 1.0.
the value of dft is obtained from: 4m
1/3
f
(-------_.- ) 11 'l/l'T P3
n
(10)
3
(rich). The value of the empirical constant D* may be obtained from [28]. D*
=
0.736 - 0.0173
(p3/~P)
(8)
D* refers only to the design point values, (i . e. , Condition 1, Table 1).
Values calculated in this way are also listed in Table 4. 7.3 Selection of Appropriate Casing and Combustor Areas Tables 3 and 4 indicate that there are considerable discrepancies between the four methods, 5-27
CHAPTER 5 suggesting that the methods available for prediction are not of high accuracy. Furthermore, some of the curves (e. g. e curve) are asymptotic as the combustion efficiency approaches 100%. Hence, the engineer must use some discretion in interpreting the results. Consider first the range of combustor diameters. For Condition 1, Table 1, pressure loss considerations are met by a combustor having a diameter of 0.28 m; it is, however, desirable to justify this value against others calculated from the alternative equations. First, the above diameter suffices for most of the reaction -based estimates. Second, it is not possible to specify the relight condition since it is a transient condition with a varying fuel/air ratio. It is possible that, if the igniter is adequate for inducing initiation, subsequent acceleration would move the combustion into a stable condition. In the extreme, a starter assist could be considered. Third, Eq . (7) suggests that a diameter of o . 48 m is required to meet the specified idling combustor efficiency (Condition 5). Acceptance of this latter value would demand a total revision of design and probably require higher pressure loss commitments. This value should, therefore, be noted as an area of probable difficulty which would require development and possible modification of the fuel injection systems. Th:is latter approach would be the lesser of two evils. Fourth, the other diameters, O. 28 5 for Condition 4, and 0.296 for Condition 1, may be regarded as being within prediction error and in satisfactory agreement with the "aerodynamic" value of 0.28 m first considered. Based upon the above evidence, the following dimensions are chosen for the combustor: Aft
6.17
x 10- 2 m2
Dft
2.80
x 10-1 m
Aref
8.81
x 10- 2 m2
Dref
3.35 x 10-1 m
where T 3 is the design point condition. This is generally taken to be the maximum thrust S. L. S. condition (Condition I, Table 1); substitution yields a value of 51.4% of the total air required for film cooling. This is a considerable quantity of the available air, but it must be remembered that some fraction of it will be used in the various zones of the combustor; it is not all parasitic. The primary zone air flow has already been estimated at 25%; the length of this zone may be assessed initially as being equal to 2/3 to 3/4 of the combustor diameter. The latter value has been selected here since it is commensurate with a high combustion efficiency. The secondary zone air is determined by considering the richest operating condition. The gases leaving the secondary zone at this condition should not be richer than 4> = O. 8. The richest overall operating condition is Condition 1, Table 1. At· Condition 1, 4>overall = 0.347. Hence, the percent air going into primary and secondary zones is given by: 0.347 x 100
----O~8---
43.38%
The length of the secondary zone is taken as half the combustor diameter. With respect to the dilution zone, if all the film cooling air were parasitic, the amount left would be 5.2%. This clearly demonstrates the designer's problem--to cool the walls adequately but, at the same time, to use a major portion of the film cooling air for combustion or dilution purposes. At this stage, the dilution air will be specified. However, the design requirement (Table 1) calls for a traverse quality of 20%, and knowledge of this enables the length of the dilution zone to be estimated. From Table 1, the value of:
Inserting this value into Fig. 21 yields LDZ/Dft 1. 3.
Primary zone air = 25% of total air flow. 7 .4 Preliminary Estimate of Remaining' Features of Combustor At this stage it is useful to estimate the size of the combustor zones and the remaining air distribution. The amount of film cooling air may be estimated from [28]. Percent film cooling air = 0.1 T 3
-
30 (11)
The role that the transition piece (from combustor to nozzle guide vanes) plays in the definition of traverse quality is controversial. Obviously, it has some effect, even though the gas is accelerating. It is suggested that , for can -type combustors, if the transition is short in comparison with the length of the dilution section, it may be neglected. Otherwise, it is suggested that half of the length of the transition be included in the length of the dilution zone. For the present example, the transition piece has been assumed to be short.
COMBUSTOR they cannot be changed, then the combustor size must be adjusted to obtain the maximum use of space, although the LID of the outer casing should be maintained within the range of 1.5 to 2.2 so as to enable the air flow to be controlled adequately. Similarly Aftl Aref should be kept within the range of 0.6 to O.72 .
Thus, by using simple empirical relationships, the primary elements of the combustor have emerged, Fig. 22, together with some zone requirements for the air distribution. To complete the design, the first step is to design the diffuser so that it has a minimum pressure loss, thereby giving maximum design possibilities for the air distribution and usage within the combustor. It will be noted that the above dimensions take no account of the engine space limitations. If the space limitations enable a longer combustor to be used, this is advantageous. Similarly some increase in the outer casing diameter might be useful. Finally, if the space limitations do not allow a combustor of the estimated size, and if
30
8.0 DIFFUSER DESIGN Analytical techniques for the design of diffusers which can adequately describe the non-uniform turbulent flow issuing from the compressor are still not available. Hence, an empirical approach based upon experimental data is necessary. Available theory, together with adequate references to the sources of diffuser maps, etc., are in [1]. The design of the diffuser is usually dictated by the space restrictions of the engine. Therefore, the objective of the designer is to generate the most efficient diffuser within that space, consistent with a minimum pressure loss. Hence, the final design will represent a compromise among the restrictions of available space, pressure loss, and uniformity of the exit flow. The following method represents a very simple approach but one that is adequate to take the combustor design to the development stage" For the present example, it will be assumed that space limitations are not restrictive and that a pressure loss of 1%is acceptable for the diffuser. The two main generic types of diffusers are illustrated in Fig. 23. For most practical applications the shortage of space will preclude the use of, a plain diffuser', and either a snout must
1111I PUSSUR! LOSS FACTOR
"p)./qnt
15
".'+~'
0.25
20
4
;::.
_
'~ 0.20
7~
4
__
4
••
~.
.::::c: .;..;...:. ,~::-; "-.:.j:.
~:'"::..:....'"
0.
,~.
_
__:~
.~"..
._
,.-_:~
)0
.
~
-
:-.;::::=:
:,._.
•• ~~~.~-";;;:j-:± i D.ISt::1±~~"~~=~c_~':"F.,'SS:";2T~·!::.::;t·I==::-:~:~~:2, 1
COMBUSTOR
Table 7. Combustor Temperature Distri bution Condo No. 1
Recirculation
Tin T T
out mean
Tin T T
3
out mean
2226 deg K
Tin
2384 deg K
T
23S4 deg K
T
1639 deg K
T. 1n
2142 deg K
T
2142 deg K
T
2252 deg K
T
T3
814 deg K
2303 deg K
T
2226 deg K
out
out
1807 deg K
707 deg K
T3
2038 deg K
T
707 deg K
out
out
2252 deg K
in
1422 deg K
out
1594 deg K
Tin
1060 deg K
T3
1060 deg K
Tin
2150 deg K
T. 1n
1890 deg K
T
2572 deg K
T
2150 deg K
T
1890 deg K
T
1420 deg K
343 deg K
T
1086 deg K
T. 1n
1086 deg K
T
1191 deg K
T
T
4
Tin
814 deg K
out
Dilution Zone
Secondary Zone
Primary Zone
Zone
out mean
T. 1n T
out
T mean
out
out
out
2068 deg K
343 deg K 1137 deg K
T3 T
out
in out
1191 deg K 726 deg K
out
872 de g K
incoming air contained within the casing and by radiation from the combustor wall to the casing wall. Under these circumstances (and in the absence of longitudinal conduction), the calculation of the wall temperature involves the following heat balance:
.
(mgl Ag) 0 • 8 dha - 0
. 2
(40)
(Tg-Tw1)
(41)
(38) The situation to be considered is a combustor burning a typical commercial liquid hydrocarbon. The combustor can be annular, canannular, or multi-can, but proven results have been obtained only for combustors with a pressure injection atomizer or with a simple diffusion gas injection system. No published literature exists for predicting wall temperatures for systems which are uncooled and use air-blast or airassist atomizers. Conduction will be limited to that across the wall. Under these circumstances, Eq , (38) may be solved, using the equations proposed by Lefebvre and Herbert [33], plus the standard conduction equation: Rl
(112) a(1 .
+ €w) €g Tgl.
(Tg2·5-Tw12.S)
S
(39)
C2
0.02
(ka/~aO.8)T 3
(mal Aa)
0 .8
(TW2
Ta)
-
dha - 0
.2
(42) (43)
If the wall is thin, then Tw 1 approximates Tw 2 . Equation (43) is then eliminated, and the only unknowns are R 1, R 2' C 1, C 2 and Tw. Typical values of predicted and measured wall temperatures for an uncooled combustor are given in Figs. 29, 30 and 31, [33]. The technique seems satisfactory for uncooled combustors, but requires modification for combustors having some form of air cooling at the walls . II . 2 Film Cooling The three common ways of cooling walls are indicated schematically in Fig. 32. Convection 5-37
CHAPTER 5
10
V
x
•o
0_
9
o
18
V
000
0-0-
/'"
0
_o~
16
14 a:
•
o/V
o o
~u
/0
12 ellperlmental
0
--
-
calculated
I
5
60
100 PRESSURE p.
10
I
140 k Po
180
cooling has already been discussed, and cooling by transpiration will be discussed later. This section will be limited to the theory and practice of film cooling . Figure 32b indicates the film cooling process. An air stream is blown through a slot in a direction tangential to the surface and parallel to the hot gas. The cool air thus forms a protective barrier to separate the wall from the hot gases. The cool film gradually mixes with the hot gas and, in the process its effectiveness as a coolant depreciates. Therefore. when the wall reaches its maximum safe temperature, the film will have to be renewed. In order to derive suitable analytical expressions describing heat transfer by film cooling, certain simplifying assumptions are needed. These are: •
Flame tube
diameter
Reference
flow rate
mass
Primary
air mass flow rate
Total
air mass flow rate
Inlet
temperature air-fuel
rl)
0 r-ef
= l72 mm (0.89
rt)
man/ma
= 0.9
Drl
Diameter
Annulus
Overall
= 22!.J mm (0.75
m m
P
ratio
1m
a
a
=
0.16
=
I
• (2.2
kgls
Ibis)
T3
= 498 deg K (897 deg R)
Z
=
•
n
•
Figure 29. Comparison of Calculated and Experimental Wall Temperatures Effect of Air Pressure [33]
The hot gas will be assumed to be of uniform temperature and velocity, and it will be assumed to be transparent (i . e., no heat transfer by radiation}. The coolant will be assumed to completely occupy the film cooling slot, to be transparent, and to have uniform temperature and velocity as it leaves the slot. The wall temperature will be assumed to be equal to that of the film adjacent to it and to be adiabatic. An adiabatic film effectiveness (film cooling efficiency) will be defined as follows: T
_9.
"'c
100 OR
x• o o
5
7
I
I
9 I
o~
0
7
- 13
/' """,,-
1
0
,/1
a:
•
-110
o
c
/
-
"I~
0
eo
-- ,
/
3 4 INLET TEMPERATURE
Inlet Other
pressure conditions
.pe-rlmeotal
- -9
calculated
I 5
T3 (l00
P3 = 84.4 kPa (12.2 as shown in Figure
OK)
psia) 29
Figure 30. Comparison of Calculated and Experimental Wall Temperatures Effect of Inlet Temperature [33]
5-38
,
- T~
( 44)
T
- T 9 c
11 .2. 1 Film Cooling Calculations It would appear logical to begin the calculations by estimating the characteristics of the flare cooling section. However, at the flare, the values of Aa and ma are defined by the geometry and flow characteristics of the first wall cooling device. Hence, it is better to choose and design the first wall cooling device before going to the flare cooling. For the example here, a wigglestrip design has been selected and, to simplify production, it will be assumed that all the film cooling devices will have the same form and gap width. The slot height, w, will be set at 2. 5 mrn , the material will be Nimonic 75, the material emissivity will be 0.8, and the maximum permissible operating temperature for the wall will be 1300 K. The permissible temperature of 1300 K may seem high, but the aircraft is designed for supersonic cruise and is probably for military application. Hence, a long -Iife engine is not a primary consideration. The normal blockage for wigglestrips ranges from about 30 to about 50%; here a value
COMBUSTOR of 50%will be assumed. There is little published information on the coefficients of discharge for film cooling devices other than that of Venneman [32, 34]. However, reasonable agreement of calculated with measured wall temperatures suggests the following values are appropriate for through-flow combustors with reasonable air distribution:
Ib Is 1
~
0.85
Splash device
0.65
"
11
0
Q fo-
Total head device
,
12'
•
•
....
a::
10
a:: .... Q.
-.
..J ..J
e
- 20
-18
0
0-
-14
1
conditions
toP
80
237 kPa (34.4
psia)
= 475 deg K (855 deg R)
as shown on Figure
29
Figure 31. Comparison of Calculated Experimental Wall Temperatures Effect of Air Mass Flow Rate [33]
Gas
~
0W4W4~4Co",bustor
sw
Coolant
= 43.2,
2.55 x 103 Pa
Therefore:
0.057
a
Cooling
b
Film Cooling
-
(46)
All values are now known for insertion in Eq . (45) to give mc i : The data may then be used to calculate mc , 1 at the other operating conditions. 11. 2. 1. 2 Calculation of Hot Mass Flow, mg
Convection
Casrnq
-
Ae is computed as being equal to the area of the film cooling gap less the blockage factor, B, and multiplied by Cd, i. e. , Ac (1 - B) Cd
--..
®0M:;S:~~~
qref
=
= T3
=
kg/s
(SI Units)
qref
Ae
P3
ma
RATE
ratio Z
air-fuel
Inlet temperature Other
3
2 MASS FLOW
Inlet pressure
(45)
143.5
-16
r---t-2-
8
Hot
0.07,
Q
~
AIR
Because this first device takes its air from the snout, the pressure losses will depend upon the latter and are calculated with respect to the design condition (Condition 1).
toP P3
o
~~
7
)2
k
. a::
9
Overall
p2(A
calculated
;
2
k m T -------
f?llpenmentill
0
~ .... fo-
Used in conjunction with film' cooling parameters, the subscripts 1, 2, 3, etc. , refer to the number of the film cooling device, counting from the head of the combustor. The pressu.re loss characteristics of the device may be expressed as: toP
5
__
::;) f0ot
11. 2. 1. 1 Calculation of Cooling Air Mass Flow, mc 1
P
3
c
Transpiration
Cooling
Gas
The hot gas mass flow is defined as that flow within the recirculation region. Although it is not
Figure 32. Types of Wall Cooling
5-39
CHAPTER 5 • (mcl Ac)
precisely known at this stage, it is possible to estimate it with sufficient accuracy. The distribution assumed is:
0 .8
dhc - 0
.2
rr, - TW1)
(52)
(53)
(1/2) nH.PZ ~'PZ
(47)
o . 02
(kal ILa 0
. 8)
where; msw is known mc. f is typically 0.03 m 3 .
(mal Aa)
0 .8
The air coming through the primary zone holes is given by:
(TW2
Ta)
-
T
a dha - 0
.2
(54)
(55)
The method of solving Eqs , (51) to (55) is:
Equations (47) and (48) assume a single row of primary zone holes. If two rows are used, then the final terms of Eq , (47) will be defined as the mass flow obtained by assuming that two-thirds of the mass flow entered the recirculation zone from the upstream row, and one-third of the mass flow, from the downstream row of holes. 11. 2. 1. 3 Calculation Air
Mass
of Annulus Flow. =«
Here ma is the air flow entering the annulus and is equal to the total air flow less that entering the snout. 11. 2. 1. 4 Calculation Ac.
Ag•
of the Areas. Aa
AC=l7·dc·w
(49)
where dc is the arithmetic mean diameter of the gap. Ag is the flame tube cross-section inside the cooling film [1.25 17 (dft - 2w) 2]; Aa is the cross-section of the gap between flame tube and casing (dft increases with the addition of film cooling devices). 11. 2. 1. 5 Calculation Temperature.
of
1. Specify the combustor Tinlet = Ta, in K.
inlet
condition
Pinlet = Pa = Pg = Pc (pressure losses are neglected) in Pa. 2 . Air Partitioning: ma, mc and m are derived from the estimates discusse'a above,· in kg/s. 3. Specify fuel type: a knowledge of the elH ratio is required for flame radiation calculation. For this example a kerosene is assumed, see Table 8. 4. Estimate the hot gas temperature, T' g' as given previously.
Tg and
5. Estimate the coolant gas entry temperatures, Ta, as given previously. 11. 2. 1. 6 Estimate
of Coolant
Temperature
See [35]. o•8 X
x
o• W
Wall Tw
8
K
( 56)
The heat transfer equations to be solved are:
where:
(50) K
u 1 for -«1 ~ 0.8 u
( 57)
c
(51)
K
( 58)
5-40
COMBUSTOR
Table 8. Fuel Properties Assumed for Worked Example Aniline
157.6
Point
Aniline/Gravity
7092
Constant
45.9
API Gravity IBP 10% 50% 90% FBP percent recovered percent residue
ASTM Distillation
Average
B.P.
Calorific
Volumetric Bolar
Value
390 380
472°1< 466°K 43.26 f1,T/kq 6.14
Temperature Pressure
Flash Point
423°K 437°K 468°y S14°K 533°K
Ratio
Cloud Point Critical
302 327 383 466 500 98.5 1.0
18600 Btu/lbm
Carbon/Hydroqen Critical
of
(Closed)
-58
OF
223°K
661
of
623°K
72.1 x 103 lbf/ft2 111) OF
3.45 ~~.Pa 316°K 3.64
Luminosity Holecular
160
Weight
~roma~ic
Content
Percent
Specific
Gravity
(60/60 OF)
Stoichiometric
18 0.800 14.72
Air/Fuel
Density
by mass
Surface Tension
at at at at
3
32°P/273°K 77°F/298°K 100°F/311°K 122°F/323°K
49.91 Ibm/ft3 799.4 kg/m -3 30 x 10 _~/m 26.3 x 10 N/m 23.3 x 10-3 N/m 24.2 x 10-3 N/m
Ultimate AnalvsisCarbon-percent Hydrogen-percent Viscosity
11 . 2. 1. 7 Estimate
€g
86.0 14.0 1.27 x 10-6 m2/s 1.80 x 10-6 m2/s 6.70 x 10-6 m2/s
(liquid) at 122°F/323°K at 77°P/298°K at -22°F/243°K
of Flame
Emissivity.
€g
= 1 - exp [-0.286 Lu P
11.2.
1.8
Estimate ka/J1.aO•8
of Conductivity
74.811 + l.674 (59)
with Lu
0.0081599 T1. 10-5
0.0691 (C/H ~ 1.82)2.71
(60)
T2.25
10- 8 T3
-
Term.
TO.75 5
-
+ 2.2539 x
2.5287 x (61) 5-41
CHAPTER 5
Table 9. Air Distribution and Geometry of Film Cooling
Air
Film Cooling Device m
g
m
a
m
A
c
2
per cent
per cent
m
14.0
7.0
2.92
5.94xlO
1st Wall 14.0
83.0
7.01
5.94xl0
2nd Wall 21.0 25.0
75.6 67.7
7.38
6.25xl0
32.4 3rd Wall 43.4
60.1 49.1
7.55
6.56xl0
4th Wall 50.9 76.3
41. 3 16.0
7.73
6.88xl0
5th Wall 84.0
8.10
7.92
7.22xlO
Flare
A
g
cent
per
Hydraulic Mean Diam.
Areas
m
h
2
m
-2 Varies -3 o2.5xlO
-2
-2
2.57xl0
2.25xl0
-2
1.93xl0
-2
1.60xl0
-2
4
Note - d
A
a
2.45xl0
x
-2
2.23xl0
2.29xl0
2.34xl0
2.40xl0
cross-sectional
'wetted'
wall
d
ha
m
-3
-3
-3
-3
-3
d
t
\
w
m
m
m
0.001
0.165
0.0020
hc
m
.
0.275
Varies 0.01 0.004 to 0.005
0.275
0.053
0.005
0.001
0.165
0.0025
0.282
0.046
0.005
0.001
0.169
0.0025
0.289
0.039
0.005
0.001
0.173
0.0025
0.296
0.032
0.005
0.001
0.178
0.0025
0.303
0.005
0.005
0.001
0.184
0.0025
area
perimeter
The data given in Tables 1 and 8 enable all of the above equations to be solved using an iteration technique. The results of the gas temperature calculations are also given in Table 7. Using Eqs . (45) to (61) it is then possible to calculate the wall temperatures for the conditions listed in Table I. Similarly, it is possible to predict the wall temperatures for the remaining wall cooling devices along the combustor. (For the remaining devices the pressure loss available will be that across the combustor.) Some caution must be observed in the use of temperatures for the cooling strip (or strips) that span two regions, (e. g. cooling ring no. 3 cools the latter part of the primary zone and the early part of the seeondary zone). In such cases, the relevant gas temperatures must be used for each zone. The results of the film cooling calculation are summarized in Tables 9 and 10. Condition No. 3 has been neglected, because of its transient nature, and so has Condition No. 5 (idling) because of the very low temperatures involved. 5-42
hg
m
2.18xlO
-2
-2
2
Varies -3 °1.5xl0
-2
-2
d
c
12.0 DESIGN OF AIR ADMISSION HOLES Figure 33 illustrates the geometric arrangement and definitions for the holes. Although the annulus air flow is generally parallel to the plane of the holes, considerable deflection of the streamlines occurs in their immediate vicinity. The amount of disturbance depends upon: • • • •
Hole Liner Liner Hole
geometry pressure loss geometry bleed ratio
If x is the hole area ratio, /3 is the hole bleed ratio, K is the hole pressure loss factor, and J1. /31 a, then [36], e,
K=
1 +
02
(t./.2/02)
{2t./.2
+
[4t./.4
(4/3_/32))
+
0 . 5}
(62)
COMBUSTOR
Table 10. Film Cooling Performance Data Device
Distance from Injector Face
Condo No.
nun
n.a.
1
n.a.
2
n.a.
4
68 to 143
1 2 4
143 to 168 168 to 189
1
4th
T
T
lll2 1229 1345 1410
829 861 909 943
967 988 1001 ll58
973 1077 ll81 1239
720 749 792 821
872 891 901 901
1371 1480 1598 1664
1078 1106 ll56 ll91
1279 1302 1313 1314
T
'
814
0.139
1594
707
0.0704
2068
1060
75 75 75
0.168 0.139 0.0704
1659 1462 1918
814 707 1060
1438 1264 1693
885 770 ll31
955 856 1256
4 1 2 4
25 25 25 46 46 46
0.168 0.139 0.0704 0.0943 0.0776 0.0394
1807 1594 2068 2017 1868 2109
814 707 1060 814 707 1060
1733 1528 1993 1815 1621 2037
853 742 llOO 893 778 ll38
978 882 1293 1070 966 1299
189 to 210 210 to 280
1 2 4 1 2 4
21 21 21 91 91 91
0.0943 0.0776 o .039!f 0.0543 0.0447 0.0227
2226 2142 2150 2305 2196 2133
814 707 1060 814 707 1060
2122 2005 2130 2233 2131 2095
861 753 llOl 1021 906 1230
ll40 1062 1293 ll96 1048 1275
280 to 350 350
1 2 4 1 2 4 1 2 4
70 70 70 120 120 120 170 170 170
0.0543 0.0447 0.0227 0.0309 0.0255 0.0129 0.0309 0.0255 0.0129
2384 2252 1890 2282 2138 1825 2179 2024 1761
814 707 1060 814 707 1060 814 707 1060
2345 2224 1955 2340 2212 1915 2308 2173 1879
978 863 ll74 ll44 1054 1242 1271 ll84 1312
1200 1039 1218 ll87 1044 1227 1229 1094 1255
1 2 4
264 264 264
0.0281 0.0231 0.01l7
1639 1422 1420
814 707 1060
1909 1723 1250
1287 ll58 ll62
1077 942 H18
to 450 5th
T
g a g c w dt.:g K deg K deg K deg K deg K 1807
Cooling
3rd
25 50 75 89
T
0.168
Flare
2nd
Local f/a
rnrn
1st
~
450 to 714
Z
25 50 75 89 25 50 75 89
5-43
CHAPTER 5
The annulus loss is generally neglected l should it be required for cylindrical combust. the following equation is suggested [38]:
where: () =
momentum loss factor.
This) in turn) leads to: +
p2
c
d
= ------~-=-!~------2 () [4K
The jet angle) sin
[1
is given by:
PZ(weak mixtures)
n
2/Ct>PZ
y
ct>when ct>~ 1
y
1 when ct>> 1
mf
fuel flow in kg/s
(7T14)
rJ
X
(84)
(rich mixtures)
The primary zone volume
0.2802
(82a)
where: EI
(82b) 1S
given as:
x 0.152 x 0.2802 x 0.058 m3
VPZ •.....0.106 m3
+ (7T/12) (83a)
=
Emission Index, gCO/kg fuel
For the combustor used in our example, Eq. (84) indicates that the CO will range from about EI = 100 g/kg at idling to about EI = 2 g/kg at full load. Hence, some improvement is required. 14.2.2 Hydrocarbons Almost the same comments apply to hydrocarbons as to carbon monoxide. For hydrocarbons, the corresponding equation is: 10glO [EI]HC
= 19.730-7.1915
10glO T3
(83b)
The predicted values of combustion efficiency relevant to the operating conditions and based on Eqs . (74) to (83) are given in Table 11. The agreement between the e predictions and those of Odgers- Carrier is considered adequate. As a result of the above analysis, it would seem that the efficiency is likely to be satisfactory at other than the ground idling condition. The general approach to stability is to utilize a e curve for a comparable combustor (if one exists) , and this has been done here, Fig. 36. 5-48
Pollutant Prediction (With No Test Data Available) This section will be limited to simple semiempirical empirical techniques. Discussion of alternative and more elegant models will be deferred to Section 15.0, A Review of Modeling Techniques. 14.2
10glO [EIJCO
where: n
VPZ
-1 J
The values of Ct>PZand e from Table 11 fall well within the stability limits of the diagram. If the stability loop must be predicted, a fair approximation to the extinction limits may be obtained by extrapolating the e values to 7J = 0 or by using the Odgers - Carrier equation, setting the combustion efficiency at 50% and solving for 4>. It now remains to establish the feasibility of ignition at the relight condition. The limited amount of ignition data available is shown in Fig. 36 and, at the light-up condition, this suggests that ignition is possible using an ignition system similar to that of Fig. 36.
(85) For the sample case, the amounts of hydrocarbon range from about EI = 31 g/kg fuel at idling to EI = O. 1 g/kg fuel at full load . The corresponding combustion efficiencies may be calculated (assuming a kerosene type fuel) from: 7J
IV
1 - 10-3 (0.24 EICO + EIHC)
which yields: 7J =
94.5% at idling
(86)
md TJ .,
99.9% at full load
Equation (86) implies that the hydrocarbon is .he parent fuel, and that the combustion ineff'i.iency is caused solely by CO and HC. Actually, iurther losses may occur due to carbon (smoke, ioot) , hydrogen and to partial oxidation prodicts such as aldehydes, ketones, and acids. Of :hese, only hydrogen has been found to be sigiificant and its presence modifies Eq , (86) to: TJ =
1 - 10-3 (0.24 EICO + EIHC + 2.82 EIH2)
(87a)
The order of magnitude of EIH EIH2
••••
(0.005 to 0.05)
2
Stability
(87b)
TJ)
7.20 x 109
= ---------T 4.5270 3
m/ s 3.5
A
~ 7 . 7 kPa
)(.1 7 • J k.Pa
9
41.J 14
16.7
Ign t t i on
Loop
•
3
Equations (84) to (87) are based upon the effects of compression ratio, assuming an ambient temperature of 300 K.
THETA PARAMETER IJ kg 0.75
68 kPa ,H. I kP"
5.37 X 1016 + ----------T 7.1915 (88)
LOOp
0102 kPa
+ o
(1 -
Equation (88) does not attempt to take into account variations of fuel/air ratios or atomizing efficiency. At idling conditions, both the operating fuel/air ratio and the degree of atomization (evaporation) differ considerably from engine to
IS:
EICO
The calculations (for details see Table 12) indicate the need to improve the combustion efficiency at the idling condition, which substantiates the results of combustion efficiency calculations. From Eqs . (84) to (87),
kPe kPo
kPa kP.1
102 kPa
Figure 36. Ignition and Extinction Characteristics
CHAPTER 5
Table 12. Pollutant Predictions for Design Conditions (a)
Condo NQ.""
Condition
Design Point Max. Thrust
CO
T
MPA
deg K
DlS
814
1.93
2594
16.0
99.95
2.0
0.06
94.4
707
0.68
2507
16.0
99.89
3.8
0.17
35.5
1060
1.77
2710
18.6
99.98
0.6
0.01
128.7
343
0.148
2293
44.7
94.46
99.9
31.4
5.9
deg K 1
Pollutants E.!. NOx HC
(b)
T max
P3
T3
n% (Eqn
88)
(Eqn
84) (Eqn 85) (Eqn 91)
SLS
Hax. Alt. Cruise 4
Normal Cruise
5
Ground Idle
(a)
T based upon maximum dissociated temperature estimated at inlet conditions. max
(b)
T;
pZ + TSZ
T
and where
and
engine. In particular, older engines can be very inefficient at idling conditions, and considerable scatter exists when the engine data are plotted. The amounts of pollutants predicted (plus predicted combustion efficiencies from the e and Odgers+Carrier correlation) indicate that, during development, modification will be required to improve the combustion efficiency at idling. 14.2.3 Carbon and Smoke. Formation The formation of smoke and carbon is sensitive to the equivalence ratio in the primary zone, the pressure, the degree of premix, and the type of fuel used. Provided that the primary zone operates at lean conditions, smoke should not be a problem. The effects of fuel type upon smoke are demonstrated in Fig. 37. Because of these effects' the equivalence ratio will have to be even weaker than stoichiometric to ensure no smoke in a multi-fuel combustor. A review of the factors controlling smoke in a gas turbine is given by Toone [40]. Appleton [41] has considered the rates of soot oxidation within temperature regimes akin to those of combustors. In particular, he compared shock-tube investigations of carbon pyrolytic graphite. It was concluded that the surface oxidation rates were similar, and that the Nagle and Strickland - Constable rate equation could be used to define the rates. This equation is: 5-50
~Po
(1
-
x)
(89)
2
where: w
surface oxidation rate (kg m 2 s-1)
MC
molecular weight of carbon (12 kg/kg mole)
x
[1 + kTI (PO (dimensionless'
Po
2
kB) l " 1
oxygen partial pressure
(Pa)
kA
1.97 x 10-3 exp [-1. 26 ( - 151OO/T) ] (s m " ")
kB
4.40 x 10-7 exp [-6.36 (-7640/T) ] (s m -1)
kT
1 . 51 x 106 exp [-4.06 (-48800/T) ] (kg S-1 m '
kZ
2.10 X 10-4 exp [1.72 (2060/T) ] (Pa-1)
2)
COMBUSTOR
•
10
!~
0
i
~
0
II
•I
z
~ :s
=t
_6
,~~---
5
~ 5~
.~
0"""-
D
"...4
/'
i
"
0
COMBUSTOR
fULL
.•
•
•
8 C
0
D
o. 1
-I CAR8OH-
I H'tOROGlN
IDliNG
ro••0
--
MASS
I
0
0
"
0
I
RATIO
Figure 37. Effect of Carbon-Hydrogen Ratio on Carbon Formation in Gas Turbine Combustors This expression provides a method of estimating soot oxidation rates that is suitable for use in the performance studies of most practical combustion systems such as gas turbines. The recession rate of a spherical soot particle may be written as: dr/dt = -w/p
(90)
where p is the density of soot (p varies according to formation; a value of 1800 kg m - J is reasonable) . The work suggests that, within typical gas turbine combustors, only soot particles having a diameter of less than O. 4 micrometers can be expected to be consumed within the combustor. This size is about the same as the initial size at soot formation and suggests that the success of a "clean-combustor" may be marginal. It is interesting to note that the maximum rate of soot removal occurs at ¢,..., 0.75, which would suggest that as much length as possible should be given to that part of the combustor having ¢ = 0.75. Usually this will be the secondary zone. Because of this sensitivity to equivalence ratio. improvements of the mixing process in the secondary zone will also help reduce the exhaust soot concentration. While the burn-out rate of carbon is known (even if somewhat inaccurately), the rate of formation is completely unknown. Thus, any knowledge of carbon or smoke formation is qualitative. Consequently, the prevention of carbon or smoke must also be of a qualitative nature. Apart from the design of a combustor, the tendency to form carbon is affected by the operating conditions and the fuel composition. Macfarlane et al. [42] found that. for many hydrocarbons (premixed flames), the threshold of
carbon formation appears to be attained at an equivalence ratio of about 1.5. This threshold seems to be relatively independent of temperature or pressure but. beyond the threshold, increases in pressure greatly accelerate the rate of carbon formation. Similarly the rate at which carbon is produced is a strong function of the hydrogen content of the fuel: the more hydrogen, the less the tendency to form carbon [43]. Because of the spectrum of air/fuel ratio, a diffusion flame produces carbon much more easily than a premixed flame. This is demonstrated for kerosene flames in Fig. 38 [44]. The difference in trend for the spray system between the 2. 1 and I. 1 and 1.6 MPa lines may not be significant and may represent the difficulty of measuring carbon in the burned gas. The effect of hydrogen content is shown in Fig. 39 [44]. The results are for premixed flames, but similar behavior has been noted for diffusion flames - - although ,in the latter case, the differences may be magnified. From all the foregoing, it is obvious that there are no correlations of soot or carbon or smoke with any of the operating parameters. The difficulty of measurement, the unknown mechanisms of formation, plus the variety of variables which are known to affect carbon formation (e.g. ¢, P J. T J. mixing, turbulence, fuel type) currently preclude all but the most qualitative predictions. The basic rule for preventing carbon formation is to design the combustor to have a well mixed primary zone operating under as lean conditions as possible and (with liquid fuels) never richer than ¢=1.5. 14.2.4 Oxides of Nitrogen (NOx) Three types of NOx have been identified: 1. Thermal NOx• which originates from the high -temperature reaction between oxygen and nitrogen
0.8 Pre •• ura -Spray ---
Injecz:1on
1.1
I're.lxed
MPa
1.6
0.6
0.2
0.6 0.8
1.0
1.2 Equlvahnc
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