TURBINE VANE ENDWALL FILM COOLING WITH SLASHFACE LEAKAGE AND DISCRETE HOLE CONFIGURATION
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Turbine vanes are typically assembled as a section containing single or double airfoil units in an annular pattern. First stage guide vane assembly results in two common mating interfaces a gap between combustor and vane endwall and another resulted from the adjacent sections, called slashface. High pressure coolant could leak through these gaps to reduce the ingestion of hot gas and achieve certain cooling benefit. As vane endwall region flow field is already very complicated due to highly three-dimensional secondary flows, then a significant influence on endwall cooling can be expected due to the gap leakage flows. To determine the effect of leakage flows from those gaps, film cooling effectiveness distributions were measured using Pressure Sensitive Paint (PSP) technique on the endwall of a scaled up, mid-range industrial turbine vane geometry with the multiple rows of discrete film cooling holes inside the passages. Experiments were performed in a blow-down wind tunnel cascade facility at the exit Mach number of 0.5 corresponding to Reynolds number of 3.8 105 based on inlet conditions and axial chord length. Passive turbulence grid was used to generate freestream turbulence level about 19% with an integral length scale of 1.7 cm. Two parameters, coolant-to-mainstream mass flow ratio and density ratio were studied. The results are presented as two-dimensional film cooling effectiveness distribution on the vane endwall surface and the corresponding spanwise averaged values along the axial direction are also demonstrated.
