Abstract

In the vicinity of gas turbine blades, a complex flow field is formed due to the flow separation, reattachment, and secondary flows, and results in locally non-uniform and high heat transfer on the surfaces. In this study, the effects of leakage flow through the slot between gas turbine vane and blade rows on the film cooling effectiveness of the forward region of the shroud ring segment were experimentally investigated. The experiment was carried out in a linear cascade with five blades. Instead of vane, a row of rods at the location of the vane trailing edge was installed to consider the wake effect. The 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 on the film cooling effectiveness were experimentally investigated by using a pressure sensitive paint technique. CO2 gas and the mixture of SF6 and N2 (25%+75%) were used as leakage flow in order to simulate leakage flow to mainstream density ratios of 1.5 and 2.0, respectively. Results showed that the area averaged 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.

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