High pressure turbine rotor stage endwall cooling and passage thermal fields as affected by leakage flow and combustor liner coolant streams

The effects of an engine-representative combustor exit temperature profile and different disc cavity leakage flow rates on endwall adiabatic effectiveness distributions and passage temperature fields in a high pressure turbine rotor stage of a gas turbine are experimentally documented. The measurements are made on a stationary linear blade row cascade with an axisymmetrically-contoured endwall of modern engine geometry and with engine-representative approach flow thermal and fluid mechanics characteristics. The measurements give insight into mixing of coolant emerging as leakage flow and combustor liner coolant mix with hot core gases ahead of the airfoil row. Reported results are thermal fields in the passage, adiabatic wall temperatures and adiabatic effectiveness values in using an engine-representative approach flow temperature profile and with approach flow temperature profiles with 1) no coolant in the approach flow (flat profile) and 2) coolant only within 10% of the span (approach flow profile with a thin thermal boundary layer).The results give insight into mixing between the leakage flow and the mainstream passage flow and its effects on endwall cooling. The results demonstrate that for the conditions studied; much of the endwall cooling is contributed by the coolant in the approach flow. This is an important result that has previously not been well documented.