The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

In today's modern gas turbine engines, the region between the rotor and the stationary shroud has the most extreme fluid-thermal conditions in the entire turbine and is characterized by a periodically unsteady threedimensional flowfield. The purpose of the present work is to conduct an unsteady study of the tip leakage flow adjacent to the shroud in real gas turbine engines using an in-house industrial computational fluid dynamics code. Both time-averaged and time-dependent data for the velocity, temperature, and mass flow rate in the tip clearance region are presented in parts 1 and 2, respectively. In part 1, it was found that near the pressure side of the tip clearance region and near the blade tip on the suction side, the leakage flow is dominant, whereas opposing flows entering through the suction side dominate near the shroud and at the suction side. This opposing flow is the combined effect of the shroud relative motion and the crossflow originating from the adjacent blade passage on the suction side. A small recirculation region was observed above the rotor passage and was attributed to the bladepassage crossflow interacting with the high-pressure region found at the suction side of the blade. This high-pressure region is caused by the combined effect of the crossflow with the shroud boundary-layer flow interacting with the tip leakage flow inside the tip clearance region. Nomenclature b = blade span C x = blade axial chord h = tip clearance height M = Mach number P = pressure r = radial coordinate T = temperature v = velocity vector W = mass flow rate x = axial coordinate y = Cartesian coordinate oriented along the circumferential direction z = Cartesian coordinate oriented along the radial direction Subscripts aw = adiabatic wall c = core flow o = stagnation properties rel = relative s = isentropic

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The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on thermal film cooling characteristics, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, the exit Mach number is 0.35, and the tests are conducted using the first row of holes only, second row of holes only, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.73 to 1.92 similar to values present in operating gas turbine engines. A mesh grid is used to give a magnitude of longitudinal turbulence intensity of 5.7% at the inlet of the test section. Results show that the best overall protection over the widest range of blowing ratios and streamwise locations is provided by either the RC holes or the RR holes. This result is particularly significant because the RR hole arrangement, which has lower manufacturing costs compared with the RC and SA arrangements, produces better or equivalent levels of performance in terms of magnitudes of adiabatic film cooling effectiveness and heat transfer coefficient. Such improved performance (relative to RA and SA holes) is most likely a result of compound angles, which increases lateral spreading. As such, the present results indicate that compound angles appear to be more effective than hole shaping in improving thermal protection relative to that given by RA holes.

The experimental investigation of the film cooling performance of louver schemes using Thermochromic Liquid Crystal technique is presented in this paper. The louver scheme allows the cooling flow to pass through a bend and impinges with the blade material, which then exits to the outer surface of the aerofoil through the film cooling hole. The cooling performance for the louver scheme was analyzed across blowing ratios of 0.5 to 1.5 at a density ratio of 0.94. The results showed that the louver scheme enhances the local and the average film cooling performances in terms of film cooling effectiveness, and net heat flux reduction better than other published film hole configurations. As well, it provides a widely spread of the secondary flow extensively over the downstream surface, thus, it enhances the lateral film cooling performance. Moreover, the louver scheme produces a lower heat transfer coefficient ratio than other film hole geometries at low and high blowing ratios. As a result, the louver scheme is expected to reduce the gas turbine airfoil’s outer surface temperature and provides superior cooling performance which increases airfoil life time.