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The growing need for increased performance from gas turbines has fueled the drive to raise turbine inlet temperatures. This results in high thermal stresses especially along the first stage nozzle guide vane cascade as the hot combustion products exiting modern day gas turbine combustors generally reach temperatures that could endanger the structural stability of these vanes and greatly reduce the vane life. The highest heat transfer coefficients in the vane passage occurs near the endwall, particularly in the leading edge-endwall junction where vortical flows cause the flow of hotter fluid in the mainstream to mix with relatively lower temperature boundary layer fluid. This work documents the computational investigation of air injection at the end wall through a cylindrical hole placed upstream of the nozzle guide vane leading edge-end wall junction. The effect of the secondary jet on the formation of the leading edge horseshoe vortex and the consequent formation of the passage vortex has been studied. For the computations, the Reynolds averaged Navier–Stokes (RANS) equations were solved with the commercial software ANSYS Fluent using the SST k-ω model. Total pressure loss coefficient and kinetic energy loss Coefficient contour plots at the exit of the cascade to estimate the effect of the endwall fluid injection on loss profiles at the vane cascade exit. Swirling strength contours were plotted at several axial chord locations in order to track the path of the passage vortex in and downstream of the vane cascade. Two different hole-positions (located at 1 hole diameter and 2 hole diameters from the leading edge) along a plane parallel to the incident flow were considered in order to study the effect of the hole position with respect to the vane leading edge-endwall junction. Three different streamwise hole inclination angles with respect to the mainstream flow direction were studied to identify the best angle for the injection of fluid through the endwall. This angle was combined with five different compound angles (0°, 30°, 45°, 60° and 90°) in order to study the effect of varying the compound angle on the leading edge vortex and the passage vortex. Each of the above studies were conducted at two different injected fluid-to-mainstream mass flow ratios (0.5% and 1%) in order to study the effect of varying injected flow rate on the formation of the leading edge vortex and the vane passage vortex. From the results it was observed that suitable selection of the secondary injection mass flow rate, injection angle and hole-position caused an absence of the leading edge horseshoe vortex and delayed migration of the passage vortex across the guide vane passage. Heat Transfer studies were also conducted to observe the absence/weakening of the leading edge vortex and the delayed pitch-wise movement of the passage vortex. / Master of Science / Gas turbines are a kind of Internal Combustion engine that convert chemical energy to mechanical energy by way of burning an air-fuel mixture to cause turbine blades to spin and produce power. A typical gas turbine consists of a compressor which compresses the air intake into the combustion chamber, the combustion chamber in which energy is released from fuel by the combustion of the air-fuel mixture, and a turbine coupled to the compressor that is made to spin by the high pressure high temperature exhaust from the combustor. In order to increase the amount of power produced per unit (by weight or volume) of fuel consumed and increase the performance of the engine, the turbine inlet temperature i.e. the temperature of the hot gas products leaving the gas turbine combustor is increased by changing the fuel flow rate into the combustors and the amount of compression of the air entering the combustor. Consequently, the first component of the turbine, the nozzle guide vane faces high thermal loading which could structurally endanger vane life. The existence of complex secondary flows (leading edge vortex, passage vortex, corner vortices) near the junction of vane’s leading edge and the turbine endwall to which the vane is connected to causes increased heat transfer at this point as opposed to other points on the vane surface. The aim of this work is to study through computational simulations how injecting high momentum fluid (air) near the leading edge junction to observe any changes to the secondary flow near the endwall. The angle at which this fluid is injected and the rate of injection of this fluid are, among others, the parameters varied in this study. The flow near the leading edge and through the vane passage is visualized and the pressures at the inlet and outlet of the test domain measured at each step to compute parameters which decide how further studies are designed. The ultimate aim of this project is to identify if injecting fluid through the endwall would prove useful in reducing the vortical flows near the endwall (thereby reducing the thermal load on the endwall).

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The turbine blade platform can be protected from hot mainstream gases by injecting cooler air through the gap between stator and rotor. The effectiveness of this film cooling method depends on the geometry of the slot, the quantity of injected air, and the secondary flows near the platform. The purpose of this study was to measure the effect of the upstream vane or stator on this type of platform cooling, as well as the effect of changes in the width of the gap. Film cooling effectiveness distributions were obtained on a turbine blade platform within a linear cascade with upstream slot injection. The width of the slot was varied as well as the mass flow rate of the injected coolant. Obstacles were placed upstream to model the effect of the upstream vane. The coolant was injected through an advanced labyrinth seal to simulate purge flow through a stator-rotor seal. The width of the opening of this seal was varied to simulate the effect of misalignment. Stationary rods were placed upstream of the cascade in four phase locations to model the unsteady wake formed at the trailing edge of the upstream vane. Delta wings were also placed in four positions to create a vortex similar to the passage vortex at the exit of the vane. The film cooling effectiveness distributions were measured using pressure-sensitive paint (PSP). Reducing the width of the slot was found to decrease the area of coolant coverage, although the film cooling effectiveness close to the slot was slightly increased. The unsteady wake was found to have a trivial effect on platform cooling, while the passage vortex from the upstream vane may significantly reduce the film cooling effectiveness.

This research work investigates the thermal performance of a film-cooled gas turbine endwall under two different mainstream flow conditions. In the first part of the research investigation, the effect of unsteady passing wakes on a film-cooled pitchwise-curved surface (representing an endwall without airfoils) was experimentally studied for heat transfer characteristics on a time-averaged basis. The temperature sensitive paint technique was used to obtain the local temperatures on the test surface. The required heat flux input was provided using foil heaters. Discrete film injection was implemented on the test surface using cylindrical holes with a streamwise inclination angle of 35? and no compound angle relative to the mean approach velocity vector. The passing wakes increased the heat transfer coefficients at both the wake passing frequencies that were experimented. Due to the increasing film cooling jet turbulence and strong jet-mainstream interaction at higher blowing ratios, the heat transfer coefficients were amplified. A combination of film injection and unsteady passing wakes resulted in a maximum pitch-averaged and centerline heat transfer augmentation of ? 28% and 31.7% relative to the no wake and no film injection case. The second part of the research study involves an experimental and numerical analysis of secondary flow and coolant film interaction in a high subsonic annular cascade with a maximum isentropic throat Mach number of ? 0.68. Endwall (platform) thermal protection is provided using discrete cylindrical holes with a streamwise inclination angle of 30? and no compound angle relative to the mean approach velocity vector. The surface flow visualization on the inner endwall provided the location of the saddle point and the three-dimensional separation lines. Computational predictions showed that the leading-edge horseshoe vortex was confined to approximately 1.5% of the airfoil span for the no film injection case and intensified with low momentum film injection. At the highest blowing ratio, the film cooling jet weakened the horseshoe vortex at the leading-edge plane. The passage vortex was intensified with coolant injection at all blowing ratios. It was seen that increasing average blowing ratio improved the film effectiveness on the endwall. The discharge coefficients calculated for each film cooling hole indicated significant non-uniformity in the coolant discharge at lower blowing ratios and the strong dependence of discharge coefficients on the mainstream static pressure and the location of three-dimensional separation lines. Near the airfoil suction side, a region of coalesced film cooling jets providing close to uniform film coverage was observed, indicative of the mainstream acceleration and the influence of three-dimensional separation lines.

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No description available.