An experimental study is carried out in a stationary linear cascade which simulates a turbine rotor to compare the thermal performance of two new axisymmetric endwall contour geometries. Measurements of endwall adiabatic film cooling effectiveness and near-endwall passage temperature fields are made for this purpose. In addition to documenting endwall contouring effects, a range of disc cavity leakage flow rates is investigated. This information is meant to quantify, over the range tested, the benefits and penalties of introducing leakage flow into the passage using the designated endwall contouring. Special attention is paid to determine whether the endwall curvature has any effect on the interaction between mainstream and secondary flows within the passage. Results indicate improved thermal performance when strong endwall curvature exists near the blade leading edge. The strong curvature causes cavity leakage flow to remain closer to the endwall, thereby increasing cooling effectiveness.

No abstract

No abstract

The effects of a representative combustor exit temperature profile on leakage flow film cooling effectiveness were experimentally documented. This was done in a stationary, linear blade row cascade with an axisymmetric blade platform of dolphin-nose-shape. Endwall adiabatic film cooling effectiveness distributions and near-endwall passage thermal fields are documented. Results from the case with a representative combustor exit temperature profile are compared to those with other combustor exit temperature profiles including a base case with a uniform temperature distribution. All cases were done in the same facility over a range of disk cavity leakage flow rates. This study quantifies the sensitivity of endwall film cooling due to coolant in the leakage flow and in the approach flow to the shape of the combustor exit temperature profile. The results indicate that leakage flow film cooling effectiveness is significantly lower with a well-mixed (uniform temperature) combustor exit temperature profile than in cases in which the combustor exit temperature distribution is strongly variable. That is, it is demonstrated that combustor cooling flow aids endwall protection considerably. It is also shown that leakage flow has only a mild influence over the endwall cooling that can be attributed to coolant in the approach flow.

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.