INTERNATIONAL CONFERENCE ON RECENT ADVANCEMENT IN MECHANICAL ENGINEERING &TECHNOLOGY (ICRAMET 15)

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1 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: 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 JCHPS Special Issue 9: April Page 287

2 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, R 1 R 0 R 1 ϕ i ϕ o 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 A ref 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 A ref 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 JCHPS Special Issue 9: April Page 288

3 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 q pz. 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 p L/q pz. 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 C D 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. JCHPS Special Issue 9: April Page 289

4 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, T K Mass Flow Rate of Fuel, m f 6.33 x 10-3 kg/s Inlet Pressure, P bar Mass Flow Rate of Air, m a 0.74 kg/s Fuel/Air Ratio 8.45 x 10-3 The fuel chosen for the engine was kerosene with chemical formula C 12H 24. 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 to fuel lean ratio of Figures 8 and 9 show the photographic view of JCHPS Special Issue 9: April Page 290

5 the setup. Figures show the temperature distribution along the length of combustion chamber at air/fuel ratios of , 26.72, 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 JCHPS Special Issue 9: April Page 291

6 Fig.10 Temperature Distribution along Length of Fig.11 Temperature Distribution along Length of Fig.12 Temperature Distribution along Length of Fig.13 Temperature Distribution along Length of Fig.14 Pressure Distribution along Length of Fig.15 Pressure Distribution along Length of JCHPS Special Issue 9: April Page 292

7 Fig.16 Pressure Distribution along Length of Fig.17 Pressure Distribution along Length of Pressure drop along the length of the combustion chamber is illustrated in Figs at different air/fuel ratios of , 26.72, 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, A. H. Lefebvre and E. R. Norster, A Design of Tubular s for optimum Mixing Performance, Technical Advances in Gas Turbine Design, pt. 3N, Proc. Inst. Mech. Eng., A. H. Lefebvre, A. R. A. F. Ibrahim, and N. C. Benson, Arthur W. Lefebvre, Gas Turbine Combustion, McGraw Hill, Factors Affecting Fresh Mixture Entrainment in Bluff-BODY Stabilized Flames, Combust. Flame, volume number 10, pp , 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 , Elsevier, New York, 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, 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 (A ), Symposium (International) on Combustion, , New York, V. W. Greenhough and A. H. Lefebvre, Some Applications of Combustion Theory of Gas Turbine Development, 16 th. W. Deacon, A Survey of the Current State of Art in Gas Turbine Design, Proc. Inst. Mech. Eng. London Part 3N, volume number 183, 1-8, JCHPS Special Issue 9: April Page 293