An Experimental Study on the Leakage Flow on the Film Cooling Effectiveness of Gas Turbine Shroud

In the vicinity of the gas turbine blades, a complex flow field is formed by various secondary flows, and results in locally non-uniform heat transfer on the surfaces. In particular, the tip leakage flow caused by the pressure difference between the blade pressure side and the suction side is accelerated through the narrow tip clearance, which promotes heat transfer not only at the tip surface but also at the shroud surface. In this study, the effects of the leakage flow through the slot between gas turbine vane and blade rows on the film cooling effectiveness of the forward region of the ring segment shroud were experimentally investigated. Leakage flow was introduced through the slot between vane and blade rows and additional coolant air was injected from the cooling holes installed at the vane outer zone. The effects of the slot geometry, hole size, and blowing ratio were experimental investigated by using the pressure sensitive paint technique. The experiment was carried out in a linear cascade with five blades and four passages. CO₂ and mixture of SF₆ and N₂ (25%+75%) were used as leakage flow in order to simulate the leakage flow to mainstream density ratios of 1.5 and 2.0, respectively. Results showed that the overall film cooling effectiveness was more affected by the slot width than the cooling hole size at the same blowing ratio, and the lower density ratio cases showed higher film cooling effectiveness than that of higher density ratio case at the same cooling configuration.