To prevent melting of turbine blades, numerous cooling schemes have been developed to cool the blades using cooler air from the compressor. Unfortunately, the clearance gap between adjacent hub sections allows coolant to leak into the hub region. Coolant flow also leaks into the hub region through gaps between individual stages. The results of a combined experimental and computational study of cooling along the hub of a first stage turbine blade caused by leakage flows are discussed in detail. Additionally, this study examines a novel cooling feature, called a microcircuit, which combines internal convective cooling with external film cooling. For the experimental investigation, scaled up blades were tested in a low speed wind tunnel. Adiabatic effectiveness measurements were made with infrared thermography of the entire hub region for a range of leakage flow conditions. For the computations, a commercially available computational fluid dynamics (CFD) code, FLUENT 6.0, was used to simulate the various flows. Results show that featherseal leakage flows provide small cooling benefits to the hub. Increases in featherseal flow provide no additional cooling to the hub region. Unlike the featherseal, leakage flows from the front rim provide ample cooling to the hub region, especially the leading edge of the blade passage. None of the leakage flows provide significant cooling to the pressure side region of the hub or trailing edge suction side. With the addition of the hub microcircuits, there is improved hub cooling of the suction side of the blades. Though the coolant exit uniformity was low and affected by the featherseal flow, the microcircuits were shown to provide more cooling along the hub region. Good agreements were observed between the computational and experimental results, though computations over-predicted front rim cooling and microcircuit uniformity. / 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 possibility of misalignments at interfaces would be increased due to individual parts' assembly and external factors during its operation. In actual engine representative conditions, the upstream misalignments have effects on turbines performance through the nozzle guide vane passages. The current experimental aerodynamic investigation over the nozzle guide vane passage was concentrated on the backward-facing step of upstream misalignments. The tests were performed using two types of vane endwall platforms in a 2D linear cascade: flat endwall and axisymmetric converging endwall. The test conditions were a Mach number of 0.85, Re_ex 1.5*10^6 based on exit condition and axial chord, and a high freestream turbulence intensity (16%), at the Virginia tech transonic cascade wind tunnel. The experimental results from the surface flow visualization and the five-hole probe measurements at the vane-passage exit were compared with the two cases with and without the backward-facing step for both types of endwall platforms. As a main source of secondary flow, a horseshoe vortex at stagnation region of the leading edge of the vane directly influences other secondary flows. The intensity of the vortex is associated with boundary layer thickness of inlet flow. In this regard, the upstream backward-facing step as a misalignment induces the separation and attachment of the inlet flow sequentially, and these cause the boundary layer of the inlet flow to reform and become thinner locally. The upstream-step positively affects loss reduction in aerodynamics due to the thinner inlet boundary layer, which attenuates a horseshoe vortex ahead of the vane cascade despite the development of the additional vortices. And converging endwall results in an increase of the effect of the upstream misalignment in aerodynamics, since the inlet boundary layer becomes thinner near the vane's leading edge due to local flow acceleration caused by steep contraction of the converging endwall. These results show good correlation with many previous studies presented herein. / Master of Science / In response to climate change and limited resources, fossil fuel prices are expected to rise and energy policies are expected to change. Under these circumstances, there is a growing demand in the industry to provide an affordable option for improving the efficiency of technology. Energy efficiency is one of most cost effective ways to improve the competitiveness of all businesses and reduce energy costs for consumers. Regarding the current study topic in particular, the gas turbine is an internal combustion engine that extracts energy, which is resultant from the liquid fuel flow, and is then converted into mechanical energy to drive a compressor or other devices. Gas turbines are used in many applications such as, to power aircraft, electrical generators, pumps, and gas compressors in industrial fields. Because the gas turbine has a probability of unaligned connections of components due to assembly characteristics of its huge size, performance is affected. To consider issue, an experimental study was conducted related to the energy efficiency for an actual engine’s representative conditions; the current study focuses on the upstream backward facing step of the unaligned connections and highlights the practical effects of the unaligned connection and converging geometry. These backward facing unaligned connections are shown to have positive effects for reducing aerodynamic losses by weakening a main source of the loss, even despite the development of the additional losses. And, the application of converging geometry to the gas turbine also results in loss reduction due to local flow acceleration. These results show good correlation with the many previous studies presented herein.

Gas turbines are widely used for power generation, producing megawatts of usable energy, but consume fossil fuels in order to do so. With gas prices on the rise, all eyes have turned to operating cost and fuel efficiency. To increase efficiency, manufactures raise the temperature of the gas that is combusted. This temperature is high above the melting point of the turbine components. In order for the gas turbine to work under these conditions, its parts must be protected. This study focuses on two aspects of cooling for turbine components. Over the last decades, researchers have investigated many aspects of film cooling, The present study investigates the impact of the stagnation region created by a downstream airfoil on endwall film cooling effectiveness with and without the presence of wake. Experimental measurements are presented for a single row of cylindrical holes inclined at 35° with hole length to diameter ratio, LID= 7.5, pitch to diameter ratio, Pl/D = 3 with a constant density ratio of 1.26, and with nitrogen as the coolant. Twelve different configurations were studied. The airfoil was positioned at X/D equal to 6.35, 12.7, and 25.4. A wake plate was added upstream of the film holes at -12.7 and -50.8 X/D. The effect of stagnation and wake was combined by placing both the airfoil and the wake plate in the test section, combining all positions of each. Baseline cases for the cooling holes alone, and the cooling holes with the airfoil and wake individually were compared to the combined effects. The experimental data shows that as the airfoil stagnation region inhibits film cooling close to the airfoil, and strong wake decreases film effectiveness. With both stagnation region and wake combined, an overall decrease in film cooling performance is observed. Higher blowing ratio increase lateral spreading of the jet promoting jet to jet interaction and mainstream interaction enhancing mixing. The presence of wake promotes jet mixing with the mainstream resulting in lower film cooling effectiveness. High performance turbine airfoils are typically cooled with a combination of internal cooling channels and impingement/film cooling. In such applications, the jets impinge against a target surface, and then exit along the channel formed by the jet plate, target plate, and side walls. Local convection coefficients are the result of both the jet impact, as well as the channel flow produced from the exiting jets. Numerous studies have explored the effects of jet array and channel configurations on both target and jet plate heat transfer coefficients. However, little work has been done in examining effects of height variation and heating on all channel walls, in which both target wall and side wall data is taken, as was neglected by previous literature. This study examines the local and averaged effects of channel height on heat transfer coefficients for target and side walls. High resolution local heat transfer coefficient distributions were measured using temperature sensitive paint and recorded via a scientific grade CCD camera. Streamwise pressure distributions for both the target and side walls was recorded and used to explain heat transfer trends. Results are presented for average jet based Reynolds numbers 17K to 45K. All experiments were carried out on a large scale single row, 15 hole impingement channel, with X/D of 5, YID of 4, and Z/D of 1, 3 and 5. Providing high quality results will aid in the validation of predictive tools and development of physics-based models.

The study of flow fields and heat transfer characteristics inside a gas turbine combustor provides one of the most serious challenges for gas turbine researchers because of the harsh environment at high temperatures. Design improvements of gas turbine combustors for higher efficiency, reduced pollutant emissions, safety and durability require better understanding of combustion in swirl flows and thermal energy transfer from the turbulent reacting flows to solid surfaces. Therefore, accurate measurement and prediction of the flows and heat loads are indispensable. This dissertation presents flow details and wall heat flux measurements for reacting flow conditions in a model gas turbine combustor. The objective is to experimentally investigate the effects of combustor operating conditions on the reacting swirl flows and heat transfer on the liner wall. The results shows the behavior of swirling flows inside a combustor generated by an industrial lean pre-mixed, axial swirl fuel nozzle and associated heat loads. Planar particle image velocimetry (PIV) data were analyzed to understand the characteristics of the flow field. Experiments were conducted with various air flow rates, equivalence ratios, pilot fuel split ratios, and inlet air temperatures. Methane and propane were used as fuel. Characterizing the impingement location on the liner, and the turbulent kinetic energy (TKE) distribution were a main part of the investigation. Proper orthogonal decomposition (POD) further analyzed the data to compare coherent structures in the reacting and non-reacting flows. Comparison between reacting and non-reacting flows yielded very striking differences. Self-similarity of the flow were observed at different operating conditions. Flow temperature measurements with a thermocouple scanning probe setup revealed the temperature distribution and flow structure. Features of premixed swirl flame were observed in the measurement. Non-uniformity of flow temperature near liner wall was observed ranging from 1000 K to 1400 K. The results provide insights on the driving mechanism of convection heat transfer. As a novel non-intrusive measurement technique for reacting flows, flame infrared radiation was measured with a thermographic camera. Features of the flame and swirl flow were observed from reconstructed map of measured IR radiation projection using Abel transformation. Flow structures in the infrared measurement agreed with observations of flame luminosity images and the temperature map. The effect of equivalence ratio on the IR radiation was observed. Liner wall temperature and heat transfer were measured with infrared thermographic camera. The combustor was operated under reacting condition to test realistic heat load inside the industrial combustors. Using quartz glass liner and KG2 filter glass, the IR camera could measure inner wall surface temperature through the glass at high temperature. Time resolved axial distributions of inner/outer wall temperature were obtained, and hot side heat flux distribution was also calculated from time accurate solution of finite difference method. The information about flows and wall heat transfer found in this work are beneficial for numerical simulations for optimized combustor cooling design. Measurement data of flow temperature, velocity field, infrared radiation, and heat transfer can be used as validation purpose or for direct inputs as boundary conditions. Time-independent location of peak location of liner wall temperature was found from time resolved wall temperature measurements and PIV flow measurements. This indicates the location where the cooling design should be able to compensate for the temperature increase in lean premixed swirl combustors. The characteristics on the swirl flows found in this study points out that the reacting changes the flow structure significantly, while the operating conditions has minor effect on the structure. The limitation of non-reacting testing must be well considered for experimental combustor studies. However, reacting testing can be performed cost-effectively for reduced number of conditions, utilizing self-similar characteristics of the flows found in this study. / Ph. D. / The study of flow fields and heat transfer characteristics inside a gas turbine combustor provides one of the most serious challenges for gas turbine researchers because of the harsh environment at high temperatures. Design improvements of gas turbine combustors for higher efficiency, reduced pollutant emissions, safety and durability require better understanding of combustion in swirl flows and thermal energy transfer from the turbulent reacting flows to solid surfaces. Therefore, accurate measurement and prediction of the flows and heat loads are indispensable. This dissertation presents flow details and wall heat flux measurements for reacting flow conditions in a model gas turbine combustor. The information about flows and wall heat transfer found in this work are beneficial for numerical simulations for optimized combustor cooling design. Measurement data of flow temperature, velocity field, infrared radiation, and heat transfer can be used as validation purpose or for direct inputs as boundary conditions. Time-independent location of peak location of liner wall temperature was found from time resolved wall temperature measurements and PIV flow measurements. This indicates the location where the cooling design should be able to compensate for the temperature increase in lean premixed swirl combustors. The characteristics on the swirl flows found in this study points out that the reacting changes the flow structure significantly, while the operating conditions has minor effect on the structure. The limitation of non-reacting testing must be well considered for experimental combustor studies. However, reacting testing can be performed cost-effectively for reduced number of conditions, utilizing self-similar characteristics of the flows found in this study.

Gas turbine vanes find themselves in very hostile environments – extremely high temperature combustion gases, much exceeding material melting temperatures, flowing over them at enormous pressures. It is necessitated due to the increased efficiency and power output at these conditions. However, this also means that, in spite of the technological advancements made, these parts need frequent repairing compared to parts placed in milder environments. Primarily due to economic reasons, gas turbine parts are repaired by companies other than the original equipment manufacturer (OEM). While multitude of condition monitoring techniques have been developed and are used in the industry for regular maintenance checks, there is no easy way to characterize the impact on thermal performance of the repairing processes involved. This thesis reports the development of a technique to address this issue. It also chronicles the test rig design, experiments conducted, development and significance of the thermal performance metric. Heated air (250 ̊C – 300 ̊C) is flown through the internal cooling passages of 8 samples each of OEM and repaired parts at two different pressure ratios (vane inlet over ambient pressure), 1.1 and 1.3. First, steady state mass flow rates through each airfoil (one part is a cluster of 4 airfoils) is experimentally determined and compared among the OEM and repaired sample sets. Second, a transient experiment is run and the surface temperatures of the airfoils are measured using multiple infrared cameras viewing both the pressure and suction side of the airfoils. A parameter involving localized vane surface temperature, airfoil inlet temperature and ambient temperature is formulated to characterize the vane thermal performance. Using statistical analysis, it is found that there is no significant difference between the OEM and repaired samples tested. The development of the discussed technique, it is expected, will help companies in the gas turbine vane repairing business to qualify their parts in a robust and efficient manner without the need to invest a lot of money in buying precision equipment, or, control chambers. Finally, a couple of further studies are recommended to further improve the qualifying procedure and thereby increase the efficiency of the technique. / Master of Science / Most manufactured parts, during its lifetime, go through wear and tear of some form. Some much more than others – a gas turbine vane is one example, owing to the hostile environments it finds itself in. While repairing turbine vanes make economic sense instead of replacing the worn-out vanes with new ones, due care must be taken to ensure that the repairs pass high quality standards of the original manufactured parts. Most, if not all, companies in the turbine repairing business rely on room-temperature air-flow testing through the internal passages of these vanes to qualify their repaired parts. This is done partly due to the complexity in replicating engine-like conditions in a test environment in addition to being very time-intensive. While room-temperature air flow comparison between repaired and original parts is a necessary test, it does not paint the whole picture. Thermal performance, or, how the vane exchanges heat with the surrounding media, is the other part which completes the puzzle. A plurality of techniques has been developed to ascertain the thermal performance of gas turbine vanes, however, these are limited in the scope of their applicability – the reason why industry is still mostly relying on airflow measurements for their part qualification. In this study, a new technique has been proposed which is agnostic of the unavoidable variations in operating conditions and easy to apply while still upholding high quality standards. This translates to huge savings to organizations which are in the business of repairing original parts, not necessarily restricted to gas turbine industry.

Applications of natural gases that contain high levels of hydrogen have become a primary interest in the gas turbine market. For reheat gas turbines, understanding of the ignition delay times of high-hydrogen natural gases is important for two reasons. First, if the ignition delay time is too short, autoignition can occur in the mixer before the primary combustor. Second, the flame in the secondary burner is stabilized by the ignition delay time of the fuel. While the ignition delay times of hydrogen and of the individual hydrocarbons in natural gases can be considered well known, there have been few previous experimental studies into the effects of different levels of hydrogen on the ignition delay times of natural gases at gas turbine conditions. In order to examine the effects of hydrogen content at gas turbine conditions, shock-tube experiments were performed on nine combinations of an L9 matrix. The L9 matrix was developed by varying four factors: natural gas higher-order hydrocarbon content of 0, 18.75, or 37.5%; hydrogen content of the total fuel mixture of 30, 60, or 80%; equivalence ratios of 0.3, 0.5, or 1; and pressures of 1, 10, or 30 atm. Temperatures ranged from 1092 K to 1722 K, and all mixtures were diluted in 90% Ar. Correlations for each combination were developed from the ignition delay times and, using these correlations, a factor sensitivity analysis was performed. It was found that hydrogen played the most significant role in ignition delay time. Pressure was almost as important as hydrogen content, especially as temperature increased. Equivalence ratio was slightly more important than hydrocarbon content of the natural gas, but both were less important than pressure or hydrogen content. Further analysis was performed using ignition delay time calculations for the full matrix of combinations (27 combinations for each natural gas) using a detailed chemical kinetics mechanism. Using these calculations, separate L9 matrices were developed for each natural gas. Correlations from the full matrix and the L9 matrix for each natural gas were found to be almost identical in each case, verifying that a thoughtfully prepared L9 matrix can indeed capture the major effects of an extended matrix.

No description available.

Trapped Vortex Combustor (TVC) is a relatively new concept for potential use in gas turbine engines addressing ever increasing demands of high efficiency, low emissions, low pressure drop, and improved pattern factor. This concept holds promise for future because of its inherent advantages over conventional swirl-stabilized combustors. The main difference between TVC and a conventional gas turbine combustor is in the way combustion is stabilized. In conventional combustors, flame is stabilized because of formation of toroidal flow pattern in the primary zone due to interaction between incoming swirling air and fuel flow. On the other hand, in TVC, there is a physical cavity in the wall of combustor with continuous injection of air and fuel leading to stable and sustained combustion. Past work related to TVC has focussed on use of two cavities in the combustor liner. In the present study, a single cavity combustor concept is evaluated through simulation and experiments for applications requiring compact combustors such as Unmanned Aerial Vehicles (UAVs) and cruise missiles. In the present work, numerical simulations were initially performed on a planar, rectangular single-cavity geometry to assess sensitivity of various parameters and to design a single-cavity TVC test rig. A water-cooled, modular, atmospheric pressure TVC test rig is designed and fabricated for reacting and non-reacting flow experiments. The unique features of this rig consist of a continuously variable length-to-depth ratio (L/D) of the cavity and optical access through quartz plates provided on three sides for visualization. Flame stabilization in the single cavity TVC was successfully achieved with methane as fuel, and the range of flow conditions for stable operation were identified. From these, a few cases were selected for detailed experimentation. Reacting flow experiments for the selected cases indicated that reducing L/D ratio and increasing cavity-air velocity favour stable combustion. The pressure drop across the single-cavity TVC is observed to be lower as compared to conventional combustors. Temperatures are measured at the exit using thermocouples and corrected for radiative losses. Species concentrations are measured at the exit using an exhaust gas analyzer. The combustion efficiency is observed to be around 98-99% and the pattern factor is observed to be in the range of 0.08 to 0.13. High-speed imaging made possible by the optical access indicates that the overall combustion is fairly steady, and there is no major vortex shedding downstream. This enabled steady-state simulations to be performed for the selected cases. Insight from simulations has highlighted the importance of air and fuel injection strategies in the cavity. From a mixing and combustion efficiency standpoint, it is desirable to have a cavity vortex that is anti-clockwise. However, the natural tendency for flow over a cavity is to form a vortex that is clockwise. The tendency to blow-out at higher inlet flow velocities is thought to be because of these two opposing effects. This interaction helps improve mixing, however leads to poor flame stability unless cavity-air velocity is strong enough to support a strong anti-clockwise vortex in the cavity. This basic understating of cavity flow dynamics can be used for further design improvements in future to improve flame stability at higher inlet flow velocities and eventually lead to the development of a practical combustor.

No description available.