Experimental investigation of rotor tip film cooling at an axial turbine with swirling inflow conditions using pressure sensitive paint

Authors

Manuel Wilhelm - Institute of Gasturbines and Aerospace Propulsion
Heinz-Peter Schiffer - Institute of Gasturbines and Aerospace Propulsion

Abstract

Rotor tip film cooling is investigated at the Large Scale Turbine Rig, which is a 1.5-stage axial turbine rig operating at low speeds. Using pressure sensitive paint, the film cooling effectiveness η at a squealer-type blade tip with cylindrical pressure-side film cooling holes is obtained. The effect of turbine inlet swirl on η is examined in comparison to an axial inflow baseline case. Coolant-to-mainstream injection ratios are varied between 0.45% and 1.74% for an engine-realistic coolant-to-mainstream density ratio of 1.5. It is shown that inlet swirl causes a reduction in η for low injection ratios by up to 26%, with the trailing edge being especially susceptible to swirl. For injection ratios greater than 0.93% however, η is increased by up to 11% for swirling inflow, while for axial inflow a further increase in coolant injection does not transfer into a gain in η.