Rising combustor exit temperatures in gas turbine engines necessitate active cooling for the downstream turbine section to avoid thermal failure. Film-cooling has long been an integral part of turbine cooling schemes. Cooling the endwall of a turbine airfoil is particularly difficult as much of the coolant is swept off the endwall by vortical flow patterns that develop in the passage. Although film-cooling has potential cooling benefits, this cooling method also leads to increased aerodynamic penalties in terms of total pressures losses in the turbine stage. This study investigated the trade-off between the cooling benefit and aerodynamic penalties associated with cooling the turbine endwall region for two different cooling hole shapes. Two commonly used film-cooling hole geometries were investigated; cylindrical holes and shaped holes. Compared to the case without any film-cooling, results showed that film-cooling with either hole geometry increased the aerodynamic losses through the turbine stage. Shaped film-cooling holes generated less total pressure losses through the turbine vane passage than cylindrical holes as a result of the separation from cylindrical hole injection having increased mixing losses. Shaped holes also provided better cooling to the endwall region than cylindrical holes, making them more effective both aerodynamically and thermally.
|Number of pages
|International Journal of Heat and Fluid Flow
|Published - Jun 2007
All Science Journal Classification (ASJC) codes
- Condensed Matter Physics
- Mechanical Engineering
- Fluid Flow and Transfer Processes