The flow and thermal fields exiting gas turbine combustors dictate the overall performance of the downstream turbine. The goal of this work was to investigate the effects of engine representative combustor exit profiles on high pressure turbine vane aerodynamics and heat transfer. The various profiles were produced using a nonreacting turbine inlet profile generator in the Turbine Research Facility (TRF) located at the Air Force Research Laboratory (AFRL). This paper reports how the pressure loading and heat transfer along the vane surface was affected by different turbine inlet pressure and temperature profiles at different span locations. The results indicate that the inlet total pressure profiles affected the aerodynamic loading by as much as 10%. The results also reveal that the combination of different total pressure and total temperature profiles significantly affected the vane heat transfer relative to a baseline test with uniform inlet total pressure and total temperature. Near the inner diameter endwall, the baseline heat transfer was reduced 30-40% over the majority of the vane surface. Near the outer dimeter endwall, it was found that certain inlet profiles could increase the baseline heat transfer by 10- 20%, while other profiles resulted in a decrease in the baseline heat transfer by 25-35%. This study also shows the importance of knowing an accurate prediction of the local flow driving temperature when determining vane surface heat transfer.
All Science Journal Classification (ASJC) codes
- Mechanical Engineering