An Experimental Study of Turbine Vane Heat Transfer With Leading Edge and Downstream Film Cooling

Abstract

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.

A Review of Hole Geometry and Coolant Density Effect on Film Cooling

“…In that case, lower heat-transfer levels are expected. Hylton et al (1988) and Nirmalan and Hylton (1990) studied the effects of parameters such as Mach number, Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio on the C3X vane film cooling. The leading edge has a showerhead array of five equally spaced rows.…”

Impact of Different Film-Cooling Modes at Leading Edge on the Aerodynamic and Heat Transfer Performance of Heavy Duty Gas Turbine

“…At present, cooling technology has been widely used in aviation turbine and ground heavy duty gas turbine. The vast majority of turbine blade cooling system is very similar regardless of the specific model of the structure [1].A lot of researches on the internal and external cooling have been conducted so far [2][3][4][5][6].…”

“…Sai Shrinicas Sreeharam [12] stated that the downstream cooling effect decreased immediately due to the air mixing with the mainstream in the low blowing ratio conditions. Hilton [13] found that the different blowing ratio at leading edge had little impact on the distribution of surface pressure. Goldstein [14] found that the position of cooling holes was closely related to the jet stagnation line by the research of cylindrical holes model at leading edge.…”

The Influence of the Boundary Layer State and Reynolds Number on Film Cooling and Heat Transfer on a Cooled Nozzle Guide Vane

“…As has been shown before by Drost and Bölcs (1999) the film cooling effectiveness is strongly influenced by the nature of the incoming boundary layer. Nirmalan and Hylton (1990) performed heat transfer experiments on a first-vane model equipped with downstream film cooling, with and without leading edge cooling over a wide range of main flow and injection conditions. They reported that high leading edge blowing can actually increase heat transfer over the entire pressure side and associated this to increased turbulence levels caused by the leading edge injection.…”