The results of an experimental study of aerodynamic (surface velocity) and heat transfer distributions over the surfaces of two different, highly loaded, low-solidity contemporary turbine vane designs are presented. The aerodynamic configurations of the two vanes were carefully selected to emphasize fundamental differences in the character of the suction surface pressure distributions and the consequent effect on surface heat transfer distributions. The experimental measurements were made in moderate-temperature, three-vane cascades under steady-state conditions. The principal independent parameters (Mach number, Reynolds number, turbulence intensity, and wall-to-gas temperature ratio) were varied over ranges consistent with actual engine operation, and the test matrix was structured to provide an assessment of the independent influence of each parameter. These measurements are intended to serve as verification data for a parallel analytical code development effort. The results of this parallel effort are briefly reviewed, and the principal conclusions to date are summarized.

This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine vane external heat transfer. Steady-state experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters—Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio—were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The vane external heat transfer data obtained in this program indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil. The data presented in this paper illustrate the interaction of these variables and should provide the airfoil designer and computational analyst with the information required to improve heat transfer design capabilities for film-cooled turbine airfoils.

The purpose of this experimental investigation was to produce a data base of end-wall heat transfer data under conditions that simulate those in the passage of the first-stage stator in advanced turbine engines. The data base is intended to be sufficiently complete to provide verification data for refined computational models, and to provide a basis for advanced core engine endwall cooling designs. A linear, two-dimensional cascade was used to generate the data base. The test plan provided data to examine the effects of exit Mach number, exit Reynolds number, inlet boundary layer thickness, gas-to-wall temperature ratio, inlet pressure gradients, and inlet temperature gradients. The data generated consist of inlet, intrapassage, and exit aerodynamic data plus intrapassage endwall heat flux, adiabatic wall temperature measurements, and inlet turbulence data.

This paper presents data showing the improvement in cooling effectiveness of turbine vanes through the application of water-air cooling technology in an industrial/utility engine application. The technique utilizes a finely dispersed water-in-air mixture that impinges on the internal surfaces of turbine airfoils to produce very high cooling rates. An airfoil was designed to contain a standard impingement tube which distributes the water-air mixture over the inner surface of the airfoil. The water flash vaporizes off the airfoil inner wall. The resulting mixture of air-steam-water droplets is then routed through a pin fin array in the trailing edge region of the airfoil where additional water is vaporized. The mixture then exits the airfoil into the gas path through trailing edge slots. Experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters — Mach number, Reynolds number, wall-to-gas temperature ratio, and coolant-to-gas mass flow ratio — were maintained over ranges consistent with typical engine conditions. Five impingement tubes were utilized to study geometry scaling, impingement tube-to-airfoil wall gap spacing, impingement tube hole diameter, and impingement tube hole patterns. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, gas-to-coolant temperature ratio, water- and air-coolant-to-gas mass flow ratios, and impingement tube geometry. Heat transfer effectiveness data obtained in this program demonstrated that overall cooling levels typical for air cooled vanes could be achieved with the water-air cooling technique with reductions of cooling air flow of significantly more than 50%.

This paper presents the effects of downstream film cooling, with and without leading edge showerhead film cooling, on turbine vane external heat transfer. Steady state experimental measurements were made in a three-vane, linear, two-dimensional cascade. The principal independent parameters — Mach number, Reynolds number, turbulence, wall-to-gas temperature ratio, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio — were maintained over ranges consistent with actual engine conditions. The test matrix was structured to provide an assessment of the independent influence of parameters of interest, namely, exit Mach number, exit Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio. The vane external heat transfer data obtained in this program indicate that considerable cooling benefits can be achieved by utilizing downstream film cooling. The downstream film cooling process was shown to be a complex interaction of two competing mechanisms. The thermal dilution effect, associated with the injection of relatively cold fluid, results in a decrease in the heat transfer to the airfoil. Conversely, the turbulence augmentation, produced by the injection process, results in increased heat transfer to the airfoil. The data presented in this paper illustrate the interaction of these variables and should provide the airfoil designer and computational analyst the information required to improve heat transfer design capabilities for film cooled turbine airfoils.