The influence of film cooling and inlet temperature profile on heat transfer for the vane row of a 1-1/2 stage transonic high-pressure turbine

The current study focuses on determination of the local heat flux for the airfoil and endwall surfaces of the vane row of a fully-cooled turbine stage. The measurements that are essential to this study were performed at the Ohio State University Gas Turbine Laboratory using the Turbine Test Facility operating in blowdown mode. The full-scale rotating turbine stage used consists of a high-pressure vane, a high-pressure rotor, and a low-pressure vane. Temperature, pressure, and heat-flux measurements are obtained at the proper corrected engine design conditions, such as the Flow Function (FF), the corrected speed, the stage Pressure Ratio (PR), and the temperature ratios of gas to wall and gas to coolant. The measurements are repeated for different vane inlet temperature profiles and different vane cooling flows in order to establish an in-depth understanding of the influence of film cooling on local heat transfer, and thus on cooling effectiveness. Double-sided Kapton heat-flux gauges are used for heat-flux measurements at different span locations along the airfoil surfaces and along the inner endwall. Film cooling is managed via numerous cooling holes located on the inner and outer endwalls, at the airfoil leading edge with a showerhead arrangement, at numerous locations on the airfoil pressure and suction surfaces, and at the vane trailing edge, which results in a fully-cooled first stage vane.;This is a unique data set in that the measurements were performed not only at the design corrected conditions for a rotating turbine stage, but also for a fully-cooled vane environment. It is the first time that heat transfer data obtained in such an environment has become available for a fully-cooled vane endwall. The unique film-cooled endwall heat transfer data demonstrated in contour plots reveals insight to the complex flow behavior that is dominant in this region, which becomes even more complicated with the addition of coolant.;Addition of cooling resulted in notable reductions in heat transfer levels, but the percent variation in heat transfer caused by the temperature profiles were still comparable to that observed in an un-cooled environment. The variations between the profiles and the cooling levels are found to be comparable on the airfoil surface, as well as to those observed between the spans. The differences between the cooling levels were more clearly observed on the airfoil pressure surface than the suction surface, and coolant had more effect in reducing heat transfer at the inner spans. At the endwall region, the profile effects are more significant than the cooling effects, resulting in larger differences in heat transfer levels. Within the range of coolant variation studied, an increase in the coolant mass flow served to smooth out the large gradients due to flow complexity in the endwall heat transfer rather than increasing the cooling effectiveness even further.;The combined trailing edge and outer endwall cooling results in significant reduction in heat transfer at all surfaces, in a growing fashion towards the trailing edge, and at the endwall exit, while the purge flow through the wheel-space cavity does not have an influence on the vane heat transfer. The reduction achieved by the vane outer cooling is comparable to the reduction obtained by the highest cooling level studied.;The hot streak inlet profiles were performed with different alignments at varying magnitudes. Alignment with vane leading edge lowers heat transfer compared to the alignment with mid-passage both at the mid-span suction surface and through the endwall passage, and increases it at the endwall exit, while the pressure surface is found to be insensitive to this switch. When the magnitude of the hot streak is increased, no observable difference is observed at the endwall.;A comparison of the current results with those obtained from a previous research program with the un-cooled version of the vane with the same geometry at similar non-dimensional experimental design operating conditions gives good comparison on the pressure surface and at the endwall, but significantly lower heat transfer on the airfoil suction surface as would be anticipated due to the ingestion occurring through the cooling holes filling the plenum and being ejected onto the suction surface.;The goal of this research was to establish an extensive database for typical engine hardware with a film-cooled first stage vane, which represents the foundation for future turbomachinery film cooling modeling and component heat transfer studies. Until this time, such a database was not available within the gas turbine industry.