INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY
(ICRAMET’ 15)
Journal of Chemical and Pharmaceutical Sciences
ISSN: 0974-2115
DESIGN AND IMPROVEMENT OF COMBUSTION CHAMBER FOR SMALL GAS
TURBINE POWER PLANT
Jai Ganesh Chetiyar R, Hemanathan, Guru prasath M, B. Selva babu*.
Aarupadai veedu institute of technology.Vinayaka Missions University.
*Corresponding author: Email: rganesh34@yahoo.in
ABSTRACT
The combustion chamber of gas turbine unit is one of the most critical components to be designed. Scanning
through literature reveals that the design methodologies for combustion chamber are available in a discrete manner and
there exist a need to compile this information and evolve a systematic design procedure for combustion chamber. The
present paper is an attempt towards presenting such a complete design methodology of combustion chamber for small
gas turbine applications. The combustion chamber for the 20 kW gas turbine engine has been designed and fabricated
as per these summarized design guidelines then checked for the axial and radial temperature profiles as well as liner
wall temperatures, experimentally. It is worth to mention that the liner wall temperatures achieved is in the vicinity of
350°C when centerline temperature is of the order of 1300°C. The pressure loss is in the range of 10% of the delivery
pressure which suggest the aerodynamic design of the combustion chamber.
Keywords: combustion chamber; gas turbine, diffuser
INTRODUCTION
During the past 40 years, gas turbine combustor technology has developed gradually and continuously, rather
than through dramatic change. The fact that many combustors now in service closely resemble their predecessors in
size, shape and general appearance should not be interpreted as evidence that little progress in combustor design has
been made. The close family resemblance among combustors stems from the fact that the basic geometry of the
combustor is dictated largely by the need for its length and frontal area to remain within the limit set by other engine
components, by the necessity for a diffuser to minimize pressure loss, and by the requirement of a liner to provide
stable operation over a wide range of air/fuel ratios. In spite of more arduous operating conditions–higher pressure,
temperature and inlet velocity–present-day combustors continue to exhibit 100 percent combustion efficiency over
their normal working range, demonstrate substantial reduction in pressure loss and pollutant emissions, and allow a
liner life that is significantly longer than those of many other engine components.
Despite these advances, the challenge to ingenuity in design is greater than ever before. New concepts and
technology are still needed to satisfy current and projected pollutant emissions regulations and to respond to the
growing emphasis on engines that can utilize a much broader range of fuels. This change of emphasis has not been
accompanied by relaxation of the more conventional requirements of durability, pattern factor and relighting capability.
In fact, except for relighting, these requirements have become more stringent as operating temperatures within the hot
section of engine have continued to rise. Relighting requirements, on the other hand, have changed very little and, until
the advent of pollution regulations, represented the most severe combustor loading condition; hence, they were
responsible for determining the size of the combustor. The desired performance requirements, will call for higher
turbine inlet temperatures and closer adherence to the design temperature profile at the turbine inlet. At the same time
the demand for greater reliability, increased durability and lower manufacturing, development and maintenance costs
seems likely to assume added importance in the future. To meet these challenges, designers have searched for concepts
that would simplify both the basic design and methods of fabrication. The paper summarizes the design concepts
available in the literature and arrives at a simplified design concept for the combustion chambers for small gas turbine
applications.
DESIGN PHILOSOPHIES
The basic objective is to summarize the design philosophies available in the literature for the design of
complete combustion chamber. The designed philosophies are presented for each component of combustion chamber
in subsequent sections.
DIFFUSER
The geometry of straight-walled diffusers may be defined in terms of their geometric parameters, as shown in
Fig.1. Area ratio AR is a major parameter, since it is directly related to the primary function of the diffuser. Nondimensional length is another parameter because as pointed out by Sovran and Klomp, in combination with the area
ratio such a length defines the overall pressure gradient, which is the principal factor in the boundary layer
development. A third parameter is the divergence angle 2 , which is not an independent variable but is related to the
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INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY
(ICRAMET’ 15)
Journal of Chemical and Pharmaceutical Sciences
ISSN: 0974-2115
other parameters by
(1)
Expression for annular diffuser generally includes the additional parameter of the annulus radius ratio at entry and an
additional wall angle. We have
(2)
where,
R1
R0
∆R1
ϕi
ϕo
inner radius at inlet
outer radius at inlet
Annulus height
inner core angle
outer core angle
(a) Twodimensional (b) Conical (c) Equiangular annular (d) Straight-core annular
Fig 1 Diffuser Geometries
AERODYNAMICS
Aerodynamic processes play a vital role in the design and performance of gas turbine combustion systems. It is
probably no great exaggeration to state that when good aerodynamic design is allied to a matching fuel-injection
system, a trouble-free combustor requiring only nominal development is virtually assured.
For straight-through combustors the optimal cross-sectional area of the casing Aref is determined from considerations of
overall pressure loss and combustion loading. However, for most industrial combustors and some aircraft combustors,
the casing area needed to meet the combustion requirements is so low as to give an unacceptably high-pressure loss.
Under these conditions the overall pressure loss dictates the casing size and Aref is obtained as
(3)
At first sight it might appear advantageous to make the liner cross-sectional area as large as possible, since
these results in lower velocities and longer residence times within the liner, both of which are highly beneficial to
ignition, stability, and combustion efficiency. Unfortunately, for any given casing area, an increase in liner diameter
can be obtained only at the expense of a reduction in annulus area. This raises the annulus velocity and lowers the
annulus static pressure, thereby reducing the static pressure drop across the liner holes. This is undesirable, since a high
static pressure drop is needed to ensure that the air jets entering the liner have adequate penetration and sufficient
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INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY
(ICRAMET’ 15)
Journal of Chemical and Pharmaceutical Sciences
ISSN: 0974-2115
turbulence intensity to promote rapid mixing with the combustion products. These considerations suggest that a
satisfactory criterion for mixing performance would be the ratio of the static pressure drop across the liner ∆p L to the
dynamic pressure of the flow in the combustion zone qpZ . If the ratio of liner cross-sectional area to casing
crosssectional area is denoted by k , then the optimal
value of the highest value of k is that which gives ∆pL/qpZ .
It can be shown that another important parameter is the diameter of the liner holes. The need of the liner holes is to
provide enough air in the primary zone for complete combustion, to provide enough air to the intermediate zone for
cooling the products of combustion and to provide a uniform temperature profile at the exit in the dilution zone.
The value of CD can be found from Fig. 2 for different hole shapes. The length of the liner can be found from the
correlation of data of the pattern factor and the liner diameter.
(8)
Here A = 0.07 and 0.05 for tubular and annular liners, respectively.
Fig 2 Influence of Hole Shape on Discharge Coefficient Fig 3 Design Curves for Conventional Combustor
COMBUSTION EFFICIENCY
Combustion inefficiency represents a waste of fuel, which is clearly unacceptable in view of the world’s dwindling
oil supply and escalation of fuel costs. Another important consideration is that combustion inefficiency is manifested in
the form of undesirable or harmful pollutant emissions, notably unburned hydrocarbons and carbon monoxide. These
implications relate the combustion efficiency with the casing area of the combustor which in turn relates it with the
liner diameter and the performance of the combustion chamber. There are three types of controls as far as combustion
is concerned.
(a) Reaction Rate Controlled System
(b) Mixing Rate Controlled System
(c) Evaporation Rate Controlled System
Of the above-mentioned controlled system, the reaction rate controlled system is considered. The burning velocity
model described by Greenhough and Lefebvre is used to determine the casing area. The casing area is the function of
the combustion loading parameter θ, given by the equation,
The value of θ, can be found from the Fig.3. The value of Temperature Dependence parameter b can be found from
Fig.4.
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INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY
(ICRAMET’ 15)
Journal of Chemical and Pharmaceutical Sciences
ISSN: 0974-2115
Fig 4 Variation of Temperature Dependence with
Equivalence Ratio
FLAME STABILIZATION
The designer has very little control over the amount of fresh mixture that is entrained into the recirculation
zone of a bluff-body flameholder. Usually this amount represents only a very small fraction of the mainstream flow, a
fraction that varies markedly with changes in air velocity and temperature [9]. With main combustors, however, air
enters the recirculation zone through various apertures in the liner wall, and the designer can control the amount of air
participating in primary combustion to within fairly close limits by proper selection of the number, size, and type of
aperture.
As a general rule, maximum stability is achieved by injecting the primary air through a small number of large
holes. This is because large holes produce large jets and large-scale flow re-circulations that provide ample time for
combustion. However, for a given air mass flow rate, an increase in hole size can be obtained only at the expense of a
reduction in the number of holes. Although no firm guidelines have been laid down for the optimal number of liner
holes for annular combustors, one opposing pair of holes per fuel injector should be regarded as the absolute minimum;
twice that number would be preferable.
The other parameters like Ignition and Fuel Injection are selected according to the requirement and after the
geometric design of the combustion chamber for gas turbine applications.
DESIGN CALCULATIONS OF ACTUAL COMBUSTION CHAMBER FOR 20 KW GAS TURBINE ENGINE
The design of the combustion chamber was undertaken using the initial conditions as obtained from the test
conducted on compressor. The parameters are listed in Table 1.
Table 1 Combustor Design Parameters
Inlet Temperature, T03
438.83 K
Mass Flow Rate of Fuel, mf 6.33 x 10-3 kg/s
Inlet Pressure, P03
2.95 bar
Mass Flow Rate of Air, ma 0.74 kg/s
Fuel/Air Ratio
8.45 x 10-3
The fuel chosen for the engine was kerosene with chemical formula C12H24. Other assumed properties of
kerosene are: Density and Lower Calorific Value. The design of combustion chamber using the above mentioned
criteria was carried out and the designed chamber is shown in Fig.5. The casing diameter is 110 mm. Figure 6
elaborates the 3 – D view of the combustion chamber.
EXPERIMENTAL VALIDATION OF DESIGN
A complete combustion chamber for a 20 kW Gas Turbine Unit has been designed as per the design
methodologies summarized herein. The photographic view of liner with different zones and wall cooling holes is
shown in Fig.7. An experimental setup was specifically developed to measure the centerline temperature, liner wall
temperature and annulus wall temperature using in total 24 numbers of K-Type thermocouples at different air/fuel
ratios ranging from fuel rich ratio of 22.7396 to fuel lean ratio of 152.4. Figures 8 and 9 show the photographic view of
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INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY
(ICRAMET’ 15)
Journal of Chemical and Pharmaceutical Sciences
ISSN: 0974-2115
the setup. Figures 10 – 13 show the temperature distribution along the length of combustion chamber at air/fuel ratios
of 22.7396, 26.72, 122.106 and 152.4, respectively. It is observed that the centerline temperature is in the vicinity of
1200°C to 1300°C and liner wall temperature is in the vicinity of 250°C to 350°C. This result clearly proves the
adequacy of the design methodologies proposed in this work.
Fig.5 Dimensional Drawing of the Combustion
Chamber Liner for 20 kW Gas Turbine Engine
Fig.6 Three Dimensional View of Combustion
Chamber
Fig 7 Liner Photographic View
Fig.8 Complete Experimental Setup
Fig.9 Setup
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INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY
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Journal of Chemical and Pharmaceutical Sciences
ISSN: 0974-2115
Fig.10 Temperature Distribution along Length of
Combustion Chamber
Fig.11 Temperature Distribution along Length of
Combustion Chamber
Fig.12 Temperature Distribution along Length of
Combustion Chamber
Fig.13 Temperature Distribution along Length of
Combustion Chamber
Fig.14 Pressure Distribution along Length of
Combustion Chamber
Fig.15 Pressure Distribution along Length of
Combustion Chamber
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INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY
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Journal of Chemical and Pharmaceutical Sciences
ISSN: 0974-2115
Fig.16 Pressure Distribution along Length of
Fig.17 Pressure Distribution along Length of
Combustion Chamber
Combustion Chamber
Pressure drop along the length of the combustion chamber is illustrated in Figs.14 – 17 at different air/fuel
ratios of 22.7396, 26.72, 122.106 and 152.4, respectively. It is worth to observe that at lower air/fuel ratios, pressure
levels on initial locations are higher than that for higher air/fuel ratios at the same locations. This is obviously due to
higher temperature levels at lower air/fuel ratios which offer more expansion of gases in a smaller volume of
combustion chamber resulting in higher pressure levels.
CONCLUSIONS
The achievement of near adiabatic flame temperature in the centerline of combustion chamber and liner wall
temperature in vicinity of 300°C may be treated as the major outcome of this work which clearly substantiates the
worthiness of the design guidelines proposed in the present work. The pressure drop along the length of the combustion
chamber being not more than 10% of the delivery pressure clearly advocates the aerodynamic design superiority.
Tolerable pressure levels achieved in primary zone also advocates for the design adequacies for such chamber as
presented in this paper.
REFERENCES
A. B. Wassell, The Design and Development of High Performance Combustors, Lecture Series 93, Von Karman
Institute of Fluid Dynamics, 1977.
A. H. Lefebvre and E. R. Norster, A Design of Tubular Combustion Chambers for optimum Mixing Performance,
Technical Advances in Gas Turbine Design, pt. 3N, Proc. Inst. Mech. Eng., 1969.
A. H. Lefebvre, A. R. A. F. Ibrahim, and N. C. Benson,
Arthur W. Lefebvre, Gas Turbine Combustion, McGraw Hill, 1984.
Factors Affecting Fresh Mixture Entrainment in Bluff-BODY Stabilized Flames, Combust. Flame, volume number 10,
pp. 231-239, 1966.
G. Sovran and E. D. Klomp, Experimentally Determined Optimum Geometries for Rectilinear Diffuser with
Rectangular, Conical or Annular Crosssection, in Fluid Mechanics of Internal Flow, pp. 270-319, Elsevier, New York,
1967.
J. H. G Howard, A. B. Thorton-trump, and H. J. Henseler,
Jack Mattingly, William Heiser and David Pratt, Aircraft Engine Design 2nd ed., AIAA Education Series, 2002.
K. S. Kaddah, Discharge Coefficients and Jet Discharge Angles for Combustor Liner Air Entry Holes, Website of
Cranfield, England.
Performance and Flow Regimes for Annular Diffusers, ASME Paper Number 67-WA/FE-21 (A68-11861), 1967.
Symposium (International) on Combustion, 858-869, New York, 1957.
V. W. Greenhough and A. H. Lefebvre, Some Applications of Combustion Theory of Gas Turbine Development, 16th.
W. Deacon, A Survey of the Current State of Art in Gas Turbine Combustion Chamber Design, Proc. Inst. Mech. Eng.
London Part 3N, volume number 183, 1-8, 1969.
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