Abstract
Secondary flows can significantly alter the surface flow field in a blade or vane passage of a turbomachine. The relationship between the three-dimensional flow field and the surface flow is essential to understand for performance improvement in gas turbines, where a secondary air flow (coolant) is injected through discrete holes provided on the endwall for cooling. The goal of the present investigation is to quantify the impact of secondary flows on film cooling and vice versa using surface flow visualization, film cooling effectiveness, and film cooling hole discharge coefficients. A high subsonic cascade with an isentropic throat Mach number of 0.68 was used for the experiments. Discrete film injection was provided on the inner endwall using a single row of cylindrical film cooling holes. It was observed that the saddle point shifted towards the airfoil leading-edge with an increase in film cooling blowing ratio. Streamlines showed that the leading-edge horseshoe vortex was confined to approximately 2% of the airfoil span for the no blowing case and intensified with film injection. Increasing the blowing ratio improved the film cooling effectiveness on the endwall. The film cooling hole discharge coefficients showed a maximum variation at the lowest blowing ratio and exhibited strong dependence on the local mainstream static pressure and the location of separation lines.
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