Turbine vane endwall film cooling considering flow nonuniformities and coolant injection from upstream combustors

Adiabatic film cooling effectiveness of a high-pressure turbine endwall and oncoming flow nonuniformities due to combustor liner coolant flows and geometrical features are investigated in this study. The endwall cooling is accomplished by discrete hole injection within the vane passage. The nonuniformities are generated by using a combustor simulator that features effusion and dilution holes in the liner walls and a step at the combustor- turbine interfacial junction. Coolant flows from louver holes are applied to the step cavity to prevent hot gas injection. The combined and separate effects of the combustor liner geometrical features and coolant flows are examined to document the combustor-turbine interactions. A more specific focus is to investigate secondary-order cooling effects of the combustor liner coolant flows on the subsequential endwall surfaces, which is expected to help reduce coolant consumption for the turbine endwall regions. The results show that total temperature and pressure profiles entering the vane passage are dictated by the combustor geometrical features and aerothermal conditions. The step is beneficial to the endwall film cooling for lower coolant flow rates but the effects are reversed for high coolant flow rates. Furthermore, secondary-order cooling effects from the combustor liner coolant are significant on the downstream endwall, which are found to be more profound than cooling effectiveness of injection from the discrete film holes within the endwall passage. Increasing coolant flow rates from the liner further improves the secondary-order cooling effects. The importance of this study lies in providing guidelines for improved cooling design for the turbine endwall, and for the combustor-turbine integration design regarding thermal issues.