Endwall Film Cooling Effects on Secondary Flows in a Contoured Endwall Nozzle Vane

The present paper investigates the effects of endwall injection of cooling flow on the aerodynamic performance of a nozzle vane cascade with endwall contouring. Tests have been performed on a seven vane cascade with a geometry typical of a real gas turbine nozzle vane. The cooling scheme consists of four rows of cylindrical holes. Tests have been carried out at low speed $(Ma2is=0.2)$ with a low inlet turbulence intensity level (1.0%) and with a coolant to mainstream mass flow ratio varied in the range from 0% (solid endwall) to 2.5%. Energy loss coefficient, secondary vorticity, and outlet angle distributions were computed from five-hole probe measured data. Contoured endwall results, with and without film cooling, were compared with planar endwall data. Endwall contouring was responsible for a significant overall loss decrease, as a result of the reduction in both profile and planar side secondary flows losses; a loss increase on the contoured side was instead observed. Like as for the planar endwall, even for the contoured endwall, coolant injection modifies secondary flows, reducing their intensity, but the relevance of the changes is reduced. Nevertheless, for all the tested injection conditions, secondary losses on the contoured side are always higher than in the planar case, while contoured cascade overall losses are lower. A unique minimum overall loss injection condition was found for both tested geometries, which corresponds to an injected mass flow rate of about 1.0%.