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A full thermal experimental assessment of a novel dendritic cooling scheme for highpressure turbine vanes has been conducted and is presented in this paper, including a comparison to the current state-of-the-art cooling arrangement for these components. The dendritic cooling system consists of cooling holes with multiple interna! branches that enhance internal heat transfer and reduce the blowing ratio at hole exit. Three sets of measurements are presented, which describe, first, the local internaI heat transfer coefficient of these structures and, secondly, the cooling flow capacity requirements and overall cooling effectiveness of a highly engine-representative dendritic geometiy. Fullcoverage surface maps of overall cooling effectiveness were acquired for both dendritic and baseline vanes in the Annular Sector Heat Transfer Facility, where scaled nearengine conditions of Mach number, Reynolds number, inlet turbulence intensity, and coolant-to-mainstream pressure ratio (or momentum flux ratio) are achieved. Engine hardware was used, with laser-sintered metal counterparts for the novel cooling geometry (their detailed configuration, design, and manufacture are discussed). The dendritic system will be shown to offer improved overall cooling effectiveness at a reduced cooling mass flow rate due to a more uniform film cooling effectiveness, a decreased tendency for films to lift off in regions of low externa! cross flow, improved through-wall heat transfer and internal cooling efficiency, increased internal wetted surface area of the cooling holes, and the enhanced turbulence induced in them. [ABSTRACT FROM AUTHOR]

A new experimental technique for the accurate measurement of steady-state metal temperature surface distributions of modern heavily film-cooled turbine vanes has been developed and is described in this paper. The technique is analogous to the thermal paint test but has been designed for fundamental research. The experimental facility consists of an annular sector cascade of high pressure (HP) turbine vanes from a current production engine. Flow conditioning is achieved by using an annular sector of deswirl vanes downstream of the test section, being both connected by a three-dimensionally contoured duct. As a result, a transonic and periodic flow, highly representative of the engine aerodynamic field, is established: Inlet turbulence levels, mainstream Mach and Reynolds numbers, and coolant-to-mainstream total pressure ratio are matched. Since the fully three-dimensional nozzle guide vane (NGV) geometry is used, the correct radial pressure gradient and secondary flow development are simulated and the cooling flow redistribution is engine-realistic. To allow heat transfer measurements to be performed, a mainstream-to-coolant temperature difference (up to 33.5°C) is generated by using two steel-wire mesh heaters, operated in series. NGV surface metal temperatures are measured (between 20°C and 40°C) by wide-band thermochromic liquid crystals. These are calibrated in situ and on a per-pixel basis against vane surface thermocouples, in a heating process that spans the entire color play and during which the turbine vanes can be assumed to slowly follow a succession of isothermal states. Experimental surface distributions of overall cooling effectiveness are presented in this paper. By employing resin vanes of the same geometry and cooling configuration (to implement adiabatic wall thermal boundary conditions) and the transient liquid crystal technique, surface distributions of external heat transfer coefficient and film cooling effectiveness can be acquired. By combining these measurements with those from the metal vanes, the results can be scaled to engine conditions with a good level of accuracy. [ABSTRACT FROM AUTHOR]