One way to increase cycle efficiency of a gas turbine engine is to operate at higher turbine inlet temperature (TIT). In most engines, the turbine inlet temperatures have increased to be well above the metallurgical limit of engine components. Film cooling of gas turbine components (blades and vanes) is a widely used technique that allows higher turbine inlet temperatures by maintaining material temperatures within acceptable limits. In this cooling method, air is extracted from the compressor and forced through internal cooling passages within turbine blades and vanes before being ejected through discrete cooling holes on the surfaces of these airfoils. The air leaving these cooling holes forms a film of cool air on the component surface which protects the part from hot gas exiting the combustor. Design optimization of the airfoil film cooling system on an engine scale is a key as increasing the amount of coolant supplied yields a cooler airfoil that will last longer, but decreases engine core flow—diminishing overall cycle efficiency. Interestingly, when contemplating the physics of film cooling, optimization is also a key to developing an effective design. The film cooling process is shown to be a complex function of at least two important mechanisms: Increasing the amount of coolant injected reduces the driving temperature (adiabatic wall temperature) of convective heat transfer—reducing heat load to the airfoil, but coolant injection also disturbs boundary layer and augments convective heat transfer coefficient due to local increase in freestream turbulence. Accurate numerical modeling of airfoil film cooling performance is a challenge as it is complicated by several factors such as film cooling hole shape, coolant-to-freestream blowing ratio, coolant-to-freestream momentum ratio, surface curvature, approaching boundary layer state, Reynolds number, Mach number, combustor-generated high freestream turbulence, turbulence length scale, and secondary flows just to name a few. Until computational methods are able to accurately simulate these factors affecting film cooling performance, experimental studies are required to assist engineers in designing effective film cooling schemes. The unique contribution of this research work is to experimentally and numerically investigate the effects of coolant injection rate or blowing ratio and exit Reynolds number/Mach number on the film cooling performance of a showerhead film cooled first stage turbine vane at high freestream turbulence (Tu = 16%) and engine representative exit flow conditions. The vane was arranged in a two-dimensional, linear cascade in a heated, transonic, blow-down wind tunnel. The same facility was also used to conduct experimental and numerical study of the effects of freestream turbulence, and Reynolds number on smooth (without film cooling holes) turbine blade and vane heat transfer at engine representative exit flow conditions. The showerhead film cooled vane was instrumented with single-sided platinum thin film gauges to experimentally determine the Nusselt number and film cooling effectiveness distributions over the surface from a single transient-temperature run. Showerhead film cooling was found to augment Nusselt number and reduce adiabatic wall temperature downstream of injection. The adiabatic effectiveness trend on the suction surface was also found to be influenced by a favorable pressure gradient due to Mach number and boundary layer transition region at all blowing ratio and exit Mach number conditions. The experimental study was also complimented with a 3-D CFD effort to calculate and explain adiabatic film cooling effectiveness and Nusselt number distributions downstream of the showerhead film cooling rows of a turbine vane at high freestream turbulence (Tu = 16%) and engine design exit flow condition (Mex = 0.76). The research work presents a new three-simulations technique to calculate vane surface recovery temperature, adiabatic wall temperature, and surface Nusselt number to completely characterize film cooling performance in a high speed flow. The RANS based v2-f turbulence model was used in all numerical calculations. CFD calculations performed with experiment-matched boundary conditions showed an overall good trend agreement with experimental adiabatic film cooling effectiveness and Nusselt number distributions downstream of the showerhead film cooling rows of the vane. / Ph. D.

An experimental study was performed to measure surface Nusselt number and film cooling effectiveness on a film cooled first stage nozzle guide vane using a transient thin film gauge (TFG) technique. The information presented attempts to further characterize the performance of shaped hole film cooling by taking measurements on a row of shaped holes downstream of leading edge showerhead injection on both the pressure and suction surfaces (hereafter PS and SS) of a 1st stage NGV. Tests were performed at engine representative Mach and Reynolds numbers and high inlet turbulence intensity and large length scale at the Virginia Tech Transonic Cascade facility. Three exit Mach/Reynolds number conditions were tested: 1.0/1,400,000; 0.85/1,150,000; and 0.60/850,000 where Reynolds number is based on exit conditions and vane chord. At Mach/Reynolds numbers of 1.0/1,450,000 and 0.85/1,150,000 three blowing ratio conditions were tested: BR = 1.0, 1.5, and 2.0. At a Mach/Reynolds number of 0.60/850,000, two blowing ratio conditions were tested: BR = 1.5 and 2.0. All tests were performed at inlet turbulence intensity of 12% and length scale normalized by leading edge diameter of 0.28. Film cooling effectiveness and heat transfer results compared well with previously published data, showing a marked effectiveness improvement (up to 2.5x) over the showerhead only NGV and agreement with published showerhead-shaped hole data. NHFR was shown to increase substantially (average 2.6x increase) with the addition of shaped holes, with only a small increase (average 1.6x increase) in required coolant mass flow. Heat transfer and effectiveness augmentation with increasing blowing ratio was shown on the pressure side, however the suction side was shown to be less sensitive to changing blowing ratio. Boundary layer transition location was shown to be within a consistent region on the suction side regardless of blowing ratio and exit Mach number. / Master of Science