In order to meet the increasingly strict emissions targets required in modern civil aviation, lean burn combustors are being pursued as a means to reduce the environmental impact of gas-turbine engines. By adopting a lean air/fuel mixture NOx production may be reduced. The increase in proportional amount of high pressure air entering directly into the combustor reduces the amount available for cooling of the combustor liner tiles. A reduced mass of air places restrictions on the porosity of cooling arrays, requiring a departure from applications of pedestal and slotted film cooling typically used to cool double skin combustor liners. An alternative approach applied to lean burn combustors places impingement and effusion arrays on the cold and hot skins respectively for cooling of both sides of the hot liner skin. Although impingement cooling is well established as a means of promoting forced convection cooling, there are many areas on a liner tile where cooling behaviour is not well characterised. Additionally, film cooling reduces combustive efficiency and increases the production of NOx and CO, prompting interest in reducing its use in combustor cooling. The research for this thesis has focussed on investigations into current and proposed geometries to identify methods to enhance cold side cooling in lean burn applications. A fully modelled combustor liner tile has been used for investigation into the impact of structural and pressure blockages on cold side cooling performance of an impingementeffusion array using a transient liquid crystal technique to measure heat transfer performance. Research has found structural blockages can reduce heat transfer performance to ~60% of typical values, with crossflow development due to pressure blockage producing similar reductions in Nusselt values to ~70% of typical. A second investigation explored enhanced cooling geometries combining a distributed impingement feed over roughened channels of pedestals at variable height (H/D) and pitch (P/D). A newly proposed 'Shielded Impingement' concept combines full height pedestals, to protect impingement jets from developing crossflow, with quarter height pedestals for turbulence enhancement of crossflow cooling. The research has found that Shielded Impingement geometries displayed the strongest cooling performance of all tested designs due primarily to increased downstream Nusselt numbers. Pressure losses were comparable to short pedestal geometries, with little apparent effect of full height pedestals. Low pressure losses mean that application to extended channels in line with the full tile geometry is possible.

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As modern gas turbines implement more and more complex geometry to increase life and efficiency, attention to unsteady aerodynamic behavior becomes more important. Computational optimization schemes are contributing to advanced geometries in order to reduce aerodynamic losses and increase the life of components. These advanced geometries are less representative of cylinder and backward facing steps which have been used as analogous geometries for most aerodynamic unsteadiness research. One region which contains a high degree of flow unsteadiness and a direct influence on engine performance is that of the MidFrame. The MidFrame (or combustor-diffuser system) is the region encompassing the main gas path from the exit of the compressor to the inlet of the first stage turbine. This region contains myriad flow scenarios including diffusion, bluff bodies, direct impingement, high degree of streamline curvature, separated flow, and recirculation. This represents the most complex and diverse flow field in the entire engine. The role of the MidFrame is to redirect the flow from the compressor into the combustion system with minimal pressure loss while supplying high pressure air to the secondary air system. Various casing geometries, compressor exit diffuser shapes, and flow conditioning equipment have been tested to reduce pressure loss and increase uniformity entering the combustors. Much of the current research in this area focuses on aero propulsion geometries with annular combustors or scaled models of the power generation geometries. Due to the complexity and size of the domain accessibility with physical probe measurements becomes challenging. The current work uses additional measurement techniques to measure flow unsteadiness in the domain. The methodology for identifying and quantifying the sources of unsteadiness are iv developed herein. Sensitivity of MidFrame unsteadiness to compressor exit conditions is shown for three different velocity profiles. The result is an extensive database of measurements which can serve as a benchmark for radical new designs to ensure that the unsteadiness levels do not supersede previous successful levels.

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Modern Dry Low Emissions (DLE) combustors are characterized by highly swirling and expanding flows that makes the convective heat load on the combustor liner gas side difficult to predict and estimate. A coupled experimental-numerical study of swirling flow and its effects on combustor liner heat transfer inside a DLE annular combustor model is presented. A simulated scaled up annular combustor shell was designed with a generic fuel nozzle provided by Solar Turbines to create the swirl in the flow. The experiment was simulated with a cold flow and heated walls. An infrared camera was used to obtain the temperature distribution along the liner wall. Experimentally measured pressure distributions were compared with the heat transfer results. The experiment was conducted at various Reynolds Numbers to investigate the effect on the heat transfer peak locations and pressure distributions. A CFD study was performed using Fluent and turbulence models and used to corroborate and verify the experimental results. Results show that the heat transfer enhancement in the annulus has slightly different characteristics for the concave and convex walls. Results also show a much slower drop in heat transfer coefficient enhancement with increasing Reynolds number compared to can combustors from a previous study. An introductory study of the effect of a soft wall on the heat transfer on the combustor liner is also presented. / Master of Science

The current demands for higher performance in gas turbine engines can be reached by raising combustion temperatures to increase thermal efficiency. Hot combustion temperatures create a harsh environment which leads to the consideration of the durability of the combustor and turbine sections. Improvements in durability can be achieved through understanding the interactions between the combustor and turbine. The flow field at a combustor exit shows non-uniformities in pressure, temperature, and velocity in the pitch and radial directions. This inlet profile to the turbine can have a considerable effect on the development of the secondary flows through the vane passage. This thesis presents a computational study of the flow field generated in a non-reacting gas turbine combustor and how that flow field convects through the downstream stator vane. Specifically, the effect that the combustor flow field had on the secondary flow pattern in the turbine was studied. Data from a modern gas turbine engine manufacturer was used to design a realistic, low speed, large scale combustor test section. This thesis presents the results of computational simulations done in parallel with experimental simulations of the combustor flow field. In comparisons of computational predictions with experimental data, reasonable agreement of the mean flow and general trends were found for the case without dilution jets. The computational predictions of the combustor flow with dilution jets indicated that the turbulence models under-predicted jet mixing. The combustor exit profiles showed non-uniformities both radially and circumferentially, which were strongly dependent on dilution and cooling slot injection. The development of the secondary flow field in the turbine was highly dependent on the incoming total pressure profile. For a case with a uniform inlet pressure in the near-wall region no leading edge vortex was formed. The endwall heat transfer was found to also depend strongly on the secondary flow field, and therefore on the incoming pressure profile from the combustor. / Master of Science

The detonation-based engine has the key advantage of increased thermodynamic efficiency over the traditional constant pressure combustor. These detonation-based engines are also known as Pressure Gain Combustion systems (PGC) and Rotating Detonation Combustor (RDC) is a form of PGC, in which the detonation wave propagates azimuthally around an annular combustor. Prior researchers have performed a high fidelity 3-D numerical simulation of a rotating detonation combustor (RDC) to understand the flow physics such as detonation wave velocity, pressure profile, wave structure; however, performing these 3-D simulations is computationally expensive. 2-D simulations are a potential alternative to reduce computational cost. In most RDCs, fuel and oxidizer are injected discretely from separate plenums, and this discrete fuel/air injection results in inhomogeneous mixing within the domain. Due to the discrete fuel injection locations, fuel/oxidizer will stratify to form localized pockets of rich and lean mixtures. The motivation of the present study is to investigate the impact of unmixedness and stratification length scales on the performance of an RDC using a 2-D numerical approach. Unmixedness, which is defined as the standard deviation of equivalence ratio normalized by the mean global equivalence ratio, is a measure of the degree of fuel-oxidizer inhomogeneity. To model the effect of unmixedness in a 2-D domain, a lognormal distribution of the fuel mass fraction is generated with a mean equivalence ratio of 1 and varying standard deviations at the inlet boundary as a numerical source term. Moreover, to model the effects of stratification length scales, fuel mass fraction at the inlet boundary cells is bundled for a given length scale, and the mass fractions for these bundles are updated based on the lognormal distribution after every three-time steps. Using this methodology, 2-D numerical analyses are carried out to investigate the performance of an RDC for an H2-air mixture with varying unmixedness and stratification length scales. Results show that mean detonation velocity decreases and wave speed variation increases with an increase in unmixedness. However, with an increase in stratification length scale mean velocity remain relatively unchanged but variation in local velocity increases. The detonation wave front corrugation also increases with an increase in mixture inhomogeneity. The mean detonation cell size increases with an increase in unmixedness. The cell shape becomes more distorted and irregular with an increase in stratification length scale and unmixedness. The combined effect of unmixedness and stratification length scale leads to a decrease in pressure gain. Overall, this concept is able to elucidate the effects of varying unmixedness and stratification length scales on the performance of an RDC. / Master of Science / Pressure Gain Combustion (PGC) system has gained significant focus in recent years due to its increased thermodynamic efficiency over a constant pressure Brayton Cycle. Rotating Detonation Combustor (RDC) is a type of PGC system, which is thermodynamically more efficient than the conventional gas turbine combustor. One of the main aspects of the detonation process is the rapid burning of the fuel-oxidizer mixture, which occurs so fast that there is not enough time for pressure to equilibrate. Therefore, the process is thermodynamically closer to a constant volume process rather than a constant pressure process. A constant volume cycle is thermodynamically more efficient than a constant pressure Brayton cycle. In an RDC, a mixture of fuel and air is injected axially, and a detonation wave propagates continuously through the circumferential section. Numerical simulation of an RDC provides additional flexibility over experiments in understanding the flow physics, detonation wave structure, and analyzing the physical and chemical processes involved in the detonation cycle. Prior researchers have utilized a full-scale 3-D numerical simulation for understanding the performance of an RDC. However, the major challenge with 3-D analyses is the computational expense. Thus, to overcome this, an inexpensive 2-D simulation is used to model the flow physics of an RDC. In most RDCs, the fuel and oxidizer are injected discretely from separate plenums. Due to the discrete fuel injection, the fuel/air mixture is never perfectly premixed and results in a stratified flow field. The objective of the current work is to develop a novel approach to independently investigate the effects of varying unmixedness and stratification length scales on RDC performance using a 2-D simulation.

The highly competitive gas turbine industry has been motivated by consumer demands for higher power-to-weight ratios, increased thermal efficiencies, and reliability while maintaining affordability. In its continual quest, the industry must continually try to raise the turbine inlet temperature, which according to the well-known Brayton cycle is key to higher engine efficiencies. The desire for increased turbine inlet temperatures creates an extremely harsh environment for the combustor liner in addition to the components downstream of the combustor. Shear layers between the dilution jets and the mainstream, as well as combustor liner film-cooling interactions create a complex mean flow field within the combustor, which is not easy to model. A completely uniform temperature and velocity profile at the combustor exit is desirable from the standpoint of reducing the secondary flows in the turbine. However, this seldom occurs due to a lack of thorough mixing within the combustor. Poor mixing results in non-uniformities, such as hot streaks, and allow non-combusted fuel to exit the combustor. This investigation developed a database documenting the thermal and flow characteristics within a combustor simulator representative of the flowfield within a gas turbine aero-engine. Three- and two-component laser Doppler velocimeter measurements were completed to quantify the flow and turbulence fields, while a thermocouple rake was used to quantify the thermal fields. The measured results show very high turbulence levels due to the dilution flow injection. Directly downstream of the dilution jets, an increased thickness in the film-cooling was noted with a fairly non-homogeneous temperature field across the combustor width. A highly turbulent shear layer was found at the leading edge of the dilution jets. Measurements also showed that a relatively extensive recirculation region existed downstream of the dilution jets. Despite the lack of film-cooling injection at the trailing edge of the dilution hole, there existed coolant flow indicative of a horse-shoe vortex wrapping around the jet. As a result of the dilution jet interaction with the mainstream flow, kidney-shaped thermal fields and counter-rotating vortices developed. These vortices serve to enhance combustor mixing. / Master of Science

Acoustic Transfer Functions Derived from Finite Element Modeling for Thermoacoustic Stability Predictions of Gas Turbine Engines Design and prediction of thermoacoustic instabilities is a major challenge in aerospace propulsion and the operation of power generating gas turbine engines. This is a complex problem in which multiple physical systems couple together. Traditionally, thermoacoustic models can be reduced to dominant physics which depend only on flame dynamics and acoustics. This is the general approach adopted in this research. The primary objective of this thesis is to describe how to obtain acoustic transfer functions using finite element modeling. These acoustic transfer functions can be coupled with flame transfer functions and other dynamics to predict the thermoacoustic stability of gas turbine engines. Results of this research effort can go beyond the prediction of instability and potentially can be used as a tool in the design stage. Consequently, through the use of these modeling tools, better gas turbine engine designs can be developed, enabling expanded operating conditions and efficiencies. This thesis presents the finite element (FE) methodology used to develop the acoustic transfer functions of the Combustion System Dynamics Laboratory (CSDL) gaseous combustor to support modeling and prediction of thermoacoustic instabilities. In this research, several different areas of the acoustic modeling were addressed to develop a representative acoustics model of the hot CSDL gaseous combustor. The first area was the development and validation of the cold acoustic finite element model. A large part of this development entailed finding simple but accurate means for representing complex geometries and boundary conditions. The cold-acoustic model of the laboratory combustor was refined and validated with the experimental data taken on the combustion rig. The second stage of the research involved incorporating the flame into the FE model and has been referred to in this thesis as hot-acoustic modeling. The hot-acoustic model also required the investigation and characterization of the flame as an acoustic source. The detailed mathematical development for the full reacting acoustic wave equation was investigated and simplified sufficiently to identify the appropriate source term for the flame. It was determined that the flame could be represented in the finite element formulation as a volumetric acceleration, provided that the flame region is small compared to acoustic wavelengths. For premixed gas turbine combustor flames, this approximation of a small flame region is generally a reasonable assumption. Both the high temperature effects and the flame as an acoustic source were implemented to obtain a final hot-acoustic FE model. This model was compared to experimental data where the heat release of the flame was measured along with the acoustic quantities of pressure and velocity. Using these measurements, the hot-acoustic FE model was validated and found to correlate with the experimental data very well. The thesis concludes with a discussion of how these techniques can be utilized in large industrial-size combustors. Insights into stability are also discussed. A conclusion is then presented with the key results from this research and some suggestions for future work. / Master of Science