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.

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

Heat management is an essential part of combustor design, as operating temperatures within the combustor generally exceed safe working temperatures of the materials employed in its construction. Two principal methods used to manage this heat are impingement and film cooling. Impingement heat transfer refers to jets of impinging fluid delivered by orifices integrated into internal structures in order to remove undesired heat. This mode of heat transfer has a relatively high effectiveness, making it an attractive method of heat management. As such, a considerable number of studies have been done on the subject providing a substantial body of useful knowledge. However, there are innovative cooling configurations being used in gas turbines which generate compressibility and temperature ratio effects on heat transfer which are currently unexplored. Presented here are data showing that these effects have a significant impact on heat transfer and new correlations are presented to account for temperature ratio and Mach number effects for a range of conditions. These findings are significant and can be applied to impinging flows in other areas of a gas turbine engine such as turbine blades and vanes. Film cooling refers to the injection of coolant onto a surface through an array of sharply angled holes. This is done in a manner that allows the coolant to remain close to the surface where it provides an insulating layer between the hot gas freestream and the cooler surface. In order to improve turbine efficiency, research efforts in film cooling are directed at reducing film cooling flow without decreasing turbine inlet temperatures. Both impingement cooling and film cooling are heavily utilized in combustor liners. Frequently, cooling air first impinges against the back side of the liner, then the spent impingement fluid passes through film cooling holes. This arrangement combines the convective heat transfer of the impinging jets convection as the coolant passes through the film cooling holes and the benefits that come from having a thin film of cool air between the combustor wall and the combustion products. In order to improve the understanding of internal cooling in gas turbine engines, the influence of previously unexplored physical parameters such as compressible flow effects and temperature ratio in impingement flows and variable blowing ratio in a film cooling array must be examined. Prior to this work, there existed in the available literature only an extremely limited exploration of compressibility effects in impingement heat transfer and the results of separately examining the effects of Mach number and Reynolds number. The film cooling literature provides no information for a full array of film cooling holes along a contraction at high blowing ratios. Exploring these effects and conditions adds to the body of available data and allows the validation of numerical predictions.

Combustion designers have considered back-side impingement cooling as the solution for modern DLE combustors. The idea is to provide more cooling to the deserved local hot spots and reserve unnecessary coolant air from local cold spots. Therefore, if accurate heat load distribution on the liners can be obtained, then an intelligent cooling system can be designed to focus more on the localized hot spots. The goal of this study is to determine the heat transfer and pressure distribution inside a typical can-annular gas turbine combustor. This is one of the first efforts in the public domain to investigate the convective heat load to combustor liner due to swirling flow generated by swirler nozzles. An experimental combustor test model was designed and fitted with a swirler nozzle provided by Solar Turbines Inc. Heat transfer and pressure distribution measurements were carried out along the combustor wall to determine the thermo-fluid dynamic effects inside a combustor. The temperature and heat transfer profile along the length of the combustor liner were determined and a heat transfer peak region was established. Constant-heat-flux boundary condition was established using two identical surface heaters, and the Infrared Thermal Imaging system was used to capture the real-time steady-state temperature distribution at the combustor liner wall. Analysis on the flow characteristics was also performed to compare the pressure distributions with the heat transfer results. The experiment was conducted at two different Reynolds numbers (Re 50,000 and Re 80,000), to investigate the effect of Reynolds Number on the heat transfer peak locations and pressure distributions. The results reveal that the heat transfer peak regions at both the Reynolds numbers occur at approximately the same location. The results from this study on a broader scale will help in understanding and predicting swirling flow effects on the local convective heat load to the combustor liner, thereby enabling the combustion engineer to design more effective cooling systems to improve combustor durability and performance. / Master of Science

A comparison study between axial and radial swirler performance in a gas turbine can combustor was conducted by investigating the correlation between combustor flow field geometry and convective heat transfer at cold flow conditions for Reynolds numbers of 50,000 and 80,000. Flow velocities were measured using Particle Image Velocimetry (PIV) along the center axial plane and radial cross sections of the flow. It was observed that both swirlers produced a strong rotating flow with a reverse flow core. The axial swirler induced larger recirculation zones at both the backside wall and the central area as the flow exits the swirler, and created a much more uniform rotational velocity distribution. The radial swirler however, produced greater rotational velocity as well as a thicker and higher velocity reverse flow core. Wall heat transfer and temperature measurements were also taken. Peak heat transfer regions directly correspond to the location of the flow as it exits each swirler and impinges on the combustor liner wall. Convective heat transfer was also measured along the liner wall of a gas turbine annular combustor fitted with radial swirlers for Reynolds numbers 210000, 420000, and 840000. The impingement location of the flow exiting from the radial swirler resulted in peak heat transfer regions along the concave wall of the annular combustor. The convex side showed peak heat transfer regions above and below the impingement area. This behavior is due to the recirculation zones caused by the interaction between the swirlers inside the annulus. / Master of Science

In modern, low emission, gas turbine combustion systems the amount of air available for cooling of the flame tube liner is limited. This has led to the development of more complex cooling systems such as cooling tiles i.e. a double skin system, as opposed to the use of more conventional cooling slots i.e. a single skin system. An isothennal experimental facility has been constructed which can incorporate 10 times full size single and double skin (cooling tile) test specimens. The specimens can be tested with or without effusion cooling and measurements have been made to characterise the flow through each cooling system along with the velocity field and cooling effectiveness distributions that subsequently develop along the length of each test section. The velocity field of the coolant film has been defined using pneumatic probes, hot-wire anemometry and PIV instrumentation, whilst gas tracing technique is used to indicate (i) the adiabatic film cooling effectiveness and (ii) mixing of the coolant film with the mainstream flow. Tests have been undertaken both with a datum low turbulence mainstream flow passing over the test section, along with various configurations in which large magnitudes and scales of turbulence were present in the mainstream flow. These high turbulence test cases simulate some of the flow conditions found within a gas turbine combustor. Results are presented relating to a variety of operating conditions for both types of cooling system. The nominal operating condition for the double skin system was at a coolant to mainstream blowing ratio of approximately 1.0. At this condition, mixing of the mainstream and coolant film was relatively small with low mainstream turbulence. However, at high mainstream turbulence levels there was rapid penetration of the mainstream flow into the coolant film. This break up of the coolant film leads to a significant reduction in the cooling effectiveness. In addition to the time-averaged characteristics, the time dependent behaviour of the .:coolantfilm was. also investigated. In particular, unsteadiness associated with large scale structures in the mainstream flow was observed within the coolant film and adjacent to the tile surface. Relative to a double skin system the single skin geometry requires a higher coolant flow rate that, along with other geometrical changes, results in typically higher coolant to mainstream velocity ratios. At low mainstream turbulence levels this difference in velocity between the coolant and mainstream promotes the generation of turbulence and mixing between the streams so leading to some reduction in cooling effectiveness. However, this higher momentum coolant fluid is more resistant to high mainstream turbulence levels and scales so that the coolant film break up is not as significant under these conditions as that observed for the double skin system. For all the configurations tested the use of effusion cooling helped restore the coolant film along the rear of the test section. For the same total coolant flow, the minimum value of cooling effectiveness observed along the test section was increased relative to the no effusion case. In addition the effectiveness of the effusion patch depends on the amount of coolant injected and the axial location of the patch. The overall experimental data suggested the importance of the initial cooling film conditions together with better understanding of the possible mechanisms that results in the rapid cooling film break-up, such as high turbulence mainstream flow and scales, and this will lead to a more effective cooling system design. This experimental data is also thought to be ideal for the validation of numerical predictions.

No description available.