Endwall film cooling performance for a first-stage guide vane with upstream combustor walls and inlet injection

Effects of an upstream combustor wall on turbine nozzle endwall film cooling performance are numerically examined in a linear cascade in this paper. Film cooling is by two rows of cooling holes at 20% of the axial chord length upstream of the vane leading edge (LE) plane. The combustor walls are modeled as flat plates with square trailing edges (TE) positioned upstream of the endwall film cooling holes. A combustor wall is in line with the LE of every second vane. The influence of the combustor wall, when shifted in the axial and tangential directions, is investigated to determine effects on passage endwall cooling for three representative film cooling blowing ratios. The results show how shed vortices from the combustor wall greatly alter the flow field near the cooling holes and inside the vane passage. Film cooling distribution patterns, particularly in the entry region and along the pressure side of the passage, are affected. The combustor wall leads to an imbalance in film cooling distribution over the endwalls for adjacent vane passages. Results show a larger effect of tangential shift of the combustor wall on endwall cooling effectiveness than the effect of an equal axial shift. The study provides guidance regarding design of combustor-to-turbine transition ducts.