Conventional heat transfer design methods for high temperature gas turbine airfoils decouple the internal and external flow. Thermal boundary conditions from these decoupled analyses are applied to the blade surfaces to predict turbine life. Typically, the domain for the external flow includes the hot gas path and the film cooling holes while the domain for the internal flow includes the internal flow passages and film cooling holes. The solid blade itself couples the external and internal flow and heat transfer. Since film cooling flow physics can play a significant role on the overall turbine blade heat transfer, there has been increased interest in capturing these effects by including the hole geometry in the solution procedure. Ideally, the complete turbine blade heat transfer analysis would be provided by efficient CFD simulations for the coupled problem including the internal passages, film cooling holes and hot gas path. By prescribing both the external flow and internal flow inflow/outflow boundary conditions, the hole physics can be included in the solution. The current paper presents results obtained for coupled simulations of the NASA C3X vane and VKI rotor which models the internal passages, hole geometries and hot gas path. In both cases, cooling is achieved by rows of pressure-side, leading-edge and suction-side film cooling holes. The rows are independently fed by span-wise, constant area plenums. The former has a total of 152 cylindrical cooling holes whereas the later has a total of 110 cylindrical/shaped holes. In addition, the C3X vane consists of 10 internal radial cooling passages of cylindrical cross-section. The simulations were conducted with the Shear Stress Transport (SST) model on a grid that extended into the viscous sub-layer along all surfaces. The computed surface pressure and external heat transfer coefficient distributions at mid-span are compared to experimental data for both cases. Internal heat transfer predictions are also presented and discussed.

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