Film cooling from a single hole and a row of holes of variable pitch to diameter ratio

An Experimental Convective Heat Transfer Investigation Around a Film-Cooled Gas Turbine Blade

Film cooling on a convex surface with zero pressure gradient flow

An experimental investigation was conducted to model the film-cooling performance for a turbine-vane leading edge using the stagnation region of a cylinder in cross flow. Experiments were conducted with a single row of spanwise-angled coolant holes for a range of the coolant blowing ratio with a freestream-to-wall temperature ratio of about 2.1 and a Reynolds number of 170,000, characteristic of the gas-turbine environment. Data from local heat-flux measurements are presented for coolant-hole injection angles of 25, 35, and 45 deg with the row of holes located at three positions relative to the stagnation line on the cylinder. Results show the spanwise (hole-to-hole) variation of heat-flux reduction due to film cooling and indicate conditions for the optimum film-cooling performance.

Stagnation region gas film cooling for turbine blade leading edge applications

Numerical investigation of film cooling effectiveness on the curved surface

Analysis of the equations governing the equilibrium conditions of a cool gas film injected along a curved wall in a hot gas stream shows that Archimedes type forces may either increase or decrease the cooling efficiency of the film, according to the value of the mass flow rate ratio, the temperature ratio of the flows and the sign of the curvature of the wall. Experiments have been performed to verify these predictions: in the case of a concave wall and relatively high mass flow rate ratios, a gain in performance has been obtained, if compared to the constant pressure flat plate; in the case of a convex wall, the cooling efficiency of the film is less than that of a film on a plate with the same pressure gradient, although it is still higher, for the same mass flow ratio than the efficiency of a film on a constant pressure plate.Copyright © 1974 by ASME

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A. Le Meur

Papers

Curvature Effects on a Turbine Blade Cooling Film