It is widely accepted that climate change is a very serious environmental concern. Levels of carbon dioxide (CO2) and other emissions in the global atmosphere have increased substantially since the industrial revolution and now increasing faster than ever before. There is a thought that this has already led to dangerous warming in the Earth’s atmosphere and relevant changes around. Emissions legislations are going to be stringent as the years will pass. Hydro carbon fuel cost is also increasing substantially; more over this is non- renewable source of energy. There is an urgent need for novel combustor technologies for reducing emission as well as exploring alternative renewable fuels without effecting combustor performance. Development of novel combustors needs comprehensive understanding of conventional combustors. The design and development of gas turbine combustors is a crucial but uncertain part of an engine development process. At present, the design process relies upon a wealth of experimental data and correlations. Some major engine manufacturers have addressed the above problem by developing computer programs based on tests and empirical data to assist combustor designers, but such programs are proprietary. There is a need of developing design methodologies for combustors which would lead to substantial contribution to knowledge in field of combustors. Developed design methodologies would be useful for researchers for preliminary design assessments of a gas turbine combustor. In this study, step by step design methodologies of dual annular radial and axial combustor, triple annular combustor and reverse flow combustor have been developed. Design methodologies developed could be used to carry out preliminary design along with performance analysis for conventional combustion chambers. In this study the author has also proposed and undertaken preliminary studies of some novel combustor concepts. A novel concept of a dilution zone less combustor has been proposed in this study. According to this concept dilution air would be introduced through nozzle guide vanes to provide an optimum temperature traverse for turbine blades. Preliminary study on novel dilution zone less combustor predicts that the length of this combustor would be shorter compared to conventional case, resulting in reduced weight, fuel burn and vibrations. Reduced fuel burn eventually leads to lower emissions. Another novel concept of combustor with hydrogen synthesis from kerosene reformation has been proposed and a preliminary studies has been undertaken in this work. Addition of hydrogen as an additive in gas turbine combustor shows large benefits to the performance of gas turbine engines in addition to reduction in NOx levels. The novel combustor would have two stages, combustion of ~5% of the hydrocarbon fuel would occur in the first stage at higher equivalence ratios in the presence of a catalyst, which would eventually lead to the formation of hydrogen rich flue gases. In the subsequent stage the hydrogen rich flue gases from the first stage would act as an additive to combustion of the hydrocarbon fuel. It has been preliminary estimated that the mixture of the hydrocarbon fuel and air could subsequently be burned at much lower equivalence ratios than conventional cases, giving better temperature profiles, flame stability limits and lower NOx emissions. The effect of different geometrical parameters on the performance of vortex controlled hybrid diffuser has also been studied. It has been predicted that vortex chamber in vortex controlled hybrid diffuser does not play any role in altering the performance of diffuser. The overall contribution to knowledge of this study is development of combustor preliminary design methodologies with different variants. The other contribution to knowledge is related to novel combustors with a capability to produce low emissions. Study on novel combustor and diffuser has yielded application of two patent applications with several other publications which has resulted in a contribution to knowledge. A list of research articles, two patents, awards and achievements are presented in Appendix C.

The simple gas turbine engine Operates on the basic Joule-Brayton cycle and it is notorious for its poor thermal efficiency. Several modifications have been made to the simple cycle in order to increase its thermal efficiency but, within the thermal and mechanical stress constrains, the efficiency still ranges between 28 and 35%. However, higher values of energy utilisation efficiency have been claimed in recent years by using low grade heat from the engine exhaust either for district heating or for raising low pressure steam for chemical processes. Both applications are not very attractive in hot countries. The concept of using the low grade thermal energy from the gas turbine exhaust to raise steam in order to drive a steam turbine and generate additional electricity, i. e. the combined power and power or CPP plant would be more attractive in hot countries than the CHP plant. It was hypothesized that the operational parameters, hence the performance of the CPP plant, would depend on the allowable gas turbine entry temperature. Hence, the exhaust gas temperature could not be decided arbitrarily. This thesis deals with the performance of the gas turbine engine operating as a part of the combined power and power plant. In a CPP plant, the gas turbine does not only produce power but also the thermal energy that is required to operate the steam turbine plant at achievable thermal efficiency. The combined gas turbine-steam turbine cycles are thermodynamically analysed. A parametric study for different configurations of the combined gas-steam cycles has been carried out to show the influence of the main parameters on the CPP cycle performance. The parametric study was carried out using realistic values in view of the known constraints and taking into account any feasible future developments. The results of the parametric study show that the maximum CPP cycle efficiency would be at a point for which the gas turbine cycle would have neither its maximum efficiency nor its maximum specific work output. It has been shown that supplementary heating or gas turbine reheating would decrease the CPP cycle efficiency; hence, it could only be justified at low gas turbine inlet temperatures. Also it has been shown that although gas turbine intercooling would enhance the performance of the gas turbine cycle, it would have only a slight effect on the CPP cycle performance. A graphical method for studying operational compatibility, i.e. matching, between gas turbine components has been developed for a steady state or equilibrium operation. The author would like to submit that the graphical method offers a novel and easy to understand approach to the complex problem of component matching. It has been shown that matching conditions between the compressor and the turbine could be satisfied by superimposing the turbine performance characteristics on the compressor performance characteristics providing the axes of both were normalised. This technique can serve as a valuable tool to determine the operating range and the engine running line. Furthermore, it would decide whether the gas turbine engine was operating in a region of adequate compressor and turbine efficiencies. A computer program capable of simulating the steady state off-design conditions of the gas turbine engine as part of the CPP plant has been developed. The program was written in Visual Basic. Also, another program was developed to simulate the steady state off-design operation of the steam turbine power plant. A combination of both programs was used to simulate the combined power plant. Finally, it could be claimed that the computer simulation of the CPP plant makes significant contribution to the design of thermal power plants as it would help in investigating the effects of the performance characteristics of the components on the performance of complete engines at the design and off-design conditions. This investigation of the CPP plant performance can be carried out at the design and engineering stages and thus help to reduce the cost of manufacturing and testing the expensive prototype engines.

The goal of the present work was to identify and implement modifications to a density-based unstructured RANS CFD algorithm, as typically used in turbomachinery flows (represented here via the RoIIs-Royce 'Hydra' code), for application to Iow Mach number gas-turbine combustor flows. The basic algorithm was modified to make it suitable for combustor relevant problems. Fixed velocity and centreline boundary conditions were added using a characteristic based method. Conserved scalar mean and variance transport equations were introduced to predict scalar mixing in reacting flows. Finally, a flarnelet thermochemistry model for turbulent non-premixed combustion with an assumed shape pdf for turbulence-chemistry interaction was incorporated. A method was identified whereby the temperature/ density provided by the combustion model was coupled directly back into the momentum equations rather than from the energy equation. Three different test cases were used to validate the numerical capabilities of the modified code, for isothermal and reacting flows on different grid types. The first case was the jet in confined cross flow associated with combustor liner-dilution jetcore flow interaction. The second was the swirling flow through a multi-stream swirler. These cases represent the main aerodynamic features of combustor primary zones. The third case was a methane-fueled coaxial jet combustor to assess the combustion model implementation. This study revealed that, via appropriate modifications, an unstructured density-based approach can be utilised to simulate combustor flows. It also demonstrated that unstructured meshes employing nonhexahedral elements were inefficient at accurate capture of flow processes in regions combining rapid mixing and strong convection at angles to cell edges. The final version of the algorithm demonstrated that low Mach RANS reacting flow simulations, commonly performed using a pressure-based approach, can successfully be reproduced using a density-based approach.

Gas turbine engines for aircraft applications are required to meet multiple performance and sizing requirements, subject to constraints established by the best available technology level. The performance requirements and limiting values of constraints that are considered by the cycle analyst conducting an engine cycle design occur at multiple operating conditions. The traditional approach to cycle analysis chooses a single design point with which to perform the on-design analysis. Additional requirements and constraints not transpiring at the design point must be evaluated in off-design analysis and therefore do not influence the cycle design. Such an approach makes it difficult to design the cycle to meet more than a few requirements and limits the number of different aerothermodynamic cycle designs that can reasonably be evaluated. Engine manufacturers have developed computational methods to create aerothermodynamic cycles that meet multiple requirements, but such methods are closely held secrets of their design process. This thesis presents a transparent and publicly available on-design cycle analysis method for gas turbine engines which generates aerothermodynamic cycles that simultaneously meet performance requirements and constraints at numerous design points. Such a method provides the cycle analyst the means to control all aspects of the aerothermodynamic cycle and provides the ability to parametrically create candidate engine cycles in greater numbers to comprehensively populate the cycle design space from which a "best" engine can be selected. This thesis develops the multi-design point on-design cycle analysis method labeled simultaneous MDP. The method is divided into three different phases resulting in an 11 step process to generate a cycle design space for a particular application. Through implementation of simultaneous MDP, a comprehensive cycle design space can be created quickly for the most complex of cycle design problems. Furthermore, the process documents the creation of each candidate engine providing transparency as to how each engine cycle was designed to meet all of the requirements. The simultaneous MDP method is demonstrated in this thesis on a high bypass ratio, separate flow turbofan with up to 25 requirements and constraints and 9 design points derived from a notional 300 passenger aircraft.

In this study, the film-cooling effectiveness in different regions of gas turbine blades was investigated with various film hole/slot configurations and mainstream flow conditions. The study consisted of three parts: 1) turbine blade span film-cooling, 2) turbine platform film-cooling, and 3) blade tip film-cooling. Pressure sensitive paint (PSP) technique was used to get the conduction-free film-cooling effectiveness distribution. Film-cooling effectiveness is assessed in terms of cooling hole geometry, blowing ratio, freestream turbulence, and coolant-to-mainstream density ratio. Blade span film-cooling test shows that the compound angle shaped holes offer better film effectiveness than the axial shaped holes. Greater coolant-to-mainstream density ratio prevents coolant to lift-off. Higher freestream turbulence causes effectiveness to drop everywhere except in the region downstream of suction side. Results are also correlated with momentum flux, compound shaped hole has the greatest optimum momentum flux ratio, and then followed by axial shaped hole, compound cylindrical hole, and axial cylindrical hole. For platform purge flow cooling, the stator-rotor gap was simulated by a typical labyrinth-like seal. Two different film-cooling hole geometries, three blowing ratios and density ratios, and two freestream turbulence are examined. Results showed that the shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. Moreover, the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity. The blade tip study was performed in a blow-down flow loop. Results show that a blowing ratio of 2.0 is found to give best results on the tip floor. Lift-off of the coolant jet can be observed for the holes closer to the leading edge as blowing ratio increases from 1.5 to 2.0. A stator vane suction side heat transfer study was conducted in a partial annular cascade. The heat transfer coefficients were measured by using the transient liquid crystal technique. At X/L=0.15, a low heat transfer region where transition occurs. The heat transfer coefficients increase toward the trailing edge as flow accelerates; a spanwise variation can be found at neat tip and hub portions due to passage and horseshoe vortices.

Thesis -- University of Adelaide, 1966. / [Typescript].

Thesis (M. App. Sc.)--Carleton University, 2005. / Includes bibliographical references. Also available in electronic format on the Internet.

Thesis (M. Sc.)--University of Hong Kong, 1997. / Includes bibliographical references (leaf 79-83).

The Large Eddy Simulation approach is employed to predict the flow physics and heat transfer characteristics of a film-cooling problem that is formed from the interaction of a coolant jet with a hot mainstream flow. The film-cooling technique is used to protect turbine blades from thermal failure, allowing the gas inlet temperature to be increased beyond the failure temperature of the turbine blade material in order to enhance the efficiency of gas turbine engines. A coolant fluid is injected into the hot mainstream through several rows of injection holes placed on the surface of a gas turbine blade in order to form a protective coolant film layer on the blade surface. However, due to the complex, unsteady and three-dimensional interactions between the coolant and the hot gases, it is difficult to achieve the desired cooling performance. Understanding of this complex flow and heat transfer process will be helpful in designing more efficiently cooled rotor blades. A comprehensive numerical investigation of a rotating film-cooling performance under different conditions is conducted in this thesis, including film-cooling on a flat surface and film-cooling on a rotating gas turbine blade. The flow-governing equations are discretised based on the finite-volumes method and then solved iteratively using the well-known SIMPLE and PISO algorithms. An in-house FORTRAN code has been developed to investigate the flat plate film-cooling configuration, while the gas turbine blade geometry has been simulated using the STAR-CCM+ CFD commercial code. The first goal of the present thesis is to investigate the physics of the flow and heat transfer, which occurs during film-cooling from a standard film hole configuration. Film-cooling performance is analysed by looking at the distribution of flow and thermal fields downstream of the film holes. The predicted mean velocity profiles and spanwise-averaged film-cooling effectiveness are compared with experimental data in order to validate the reliability of the LES technique. Comparison of adiabatic film-cooling effectiveness with experiments shows excellent agreement for the local and spanwise-averaged film-cooling effectiveness, confirming the correct prediction of the film-cooling behaviour. The film coverage and film-cooling effectiveness distributions are presented along with discussions of the influence of blowing ratio and rotation number. Overall, it was found that both rotation number and blowing ratio play significant roles in determining the film-cooling effectiveness distributions. The second goal is to investigate the impact of innovative anti-vortex holes on the film-cooling performance. The anti-vortex hole design counteracts the detrimental kidney vorticity associated with the main hole, allowing coolant to remain attached to the blade surface. Thus, the new design significantly improves the film-cooling performance compared to the standard hole arrangement, particularly at high blowing ratios. The anti-vortex hole technique is unique in that it requires only readily machinable round holes, unlike shaped film-cooling holes and other advanced concepts. The effects of blowing ratio and the positions of the anti-vortex side holes on the physics of the hot mainstream-coolant interaction in a film-cooled turbine blade are also investigated. The results also indicate that the side holes of the anti-vortex design promote the interaction between the vortical structures; therefore, the film coverage contours reveal an improvement in the lateral spreading of the coolant jet.

New technologies are constantly developing towards the goal of increasing the performance of gas turbine engines while reducing pollutant emissions. The design of the combustion system is vital in the drive to reduce pollutants in order to meet legislative targets. As part of this, the fuel injector is crucial in preparing the fuel for combustion through atomization and correct mixing with the air flow. Thus, it is desirable to develop techniques to allow the analysis of performance in these key criteria and improve the understanding of both fuel injector aerodynamics and fuel atomisation. Particle Image Velocimetry (PIV) allows for spatially resolved velocity data of flow fields to be recorded and therefore enables the inspection of flow behaviour.