The Effects of Injection Angle and Hole Exit Shape on Turbine Nozzle Pressure Side Film Cooling

Abstract

Using the pressure sensitive paint (PSP) technique, film cooling effectiveness was measured on a turbine vane pressure surface, with a four-row showerhead cooling hole configuration and a single row of holes on the pressure side. Nitrogen gas was used to simulate film cooling flow providing an oxygen concentration map corresponding to an effectiveness map by the mass transfer analogy. Three showerhead coolant injection angles (45°, 90°, and 135°) were studied and two pressure side injection angles (20° and 40°) for cylindrical holes and a 40° angle for shaped hole were studied. In addition, studies were performed on three combinations of shower head and pressure side injections. Film effectiveness was measured for each of the cases at three blowing ratios. The pressure sensitive paint (PSP) was used to indicate oxygen concentration and was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the airfoil surface. The results indicate that 45° spanwise angle injection provides best film coverage for the shower head injections. For pressure side injections, the 20° cylindrical hole injection results in the highest effectiveness values and the shaped hole improves film effectiveness immediately downstream from the injection point. The film effectiveness for three combined injections and the interaction between showerhead injection and the pressure side injection are also presented and discussed.

Simulation of Mist Film Cooling on Rotating Gas Turbine Blades

“…There were numerous studies that focused on film cooling over flat surfaces with streamwise coolant injection in the past decades, eg. [1][2][3]; others studied film cooling in airfoil cascade environments to better simulate the flow and heat transfer mechanisms at engine conditions [4][5][6]. While most of the above studies were conducted at the stationary cascade blades, studies on rotating turbine are also abundant.…”

Experimental investigation of wall thickness and hole shape variation effects on full-coverage film cooling performance for a gas turbine vane

“…They found that adiabatic effectiveness increased for fan-shaped holes up to blowing ratio 1.5. Furthermore, Zhang et al [17] also showed that adiabatic effectiveness decreased for fan-shaped holes on the pressure side if blowing ratio increased above 1.5. Nathan et al [18] investigated the showerhead and one additional row of cooling on the suction and the pressure surface of a turbine vane in terms of overall cooling effectiveness.…”

Film cooling sensitivity of laidback fanshape holes to variations in exit configuration and mainstream turbulence intensity

“…[11][12][13][14][15][16]. Guo et al [11] found that fanshape holes produced higher cooling effectiveness and lower heat transfer coefficient than cylindrical holes on both suction and pressure side of a nozzle vane model in a transonic annular cascade.…”

“…Guo et al [11] found that fanshape holes produced higher cooling effectiveness and lower heat transfer coefficient than cylindrical holes on both suction and pressure side of a nozzle vane model in a transonic annular cascade. Zhang et al [12,13] investigated the influence of blowing ratio on the fanshape hole film cooling on a vane blade model. The results showed that when the blowing ratio was less than 1.5 the cooling effectiveness of fanshape holes increased with the increasing of blowing ratio.…”