Turbine Blade Platform Impingement Cooling With Effects of Film Extraction and Roughness

With the movement of combustion systems to flatter exit temperature profiles for the purpose of reduced emissions, vane endwalls and blade platforms are experiencing higher gas temperatures than previously. While vane endwalls may be cooled by conventional methods extended from the airfoil techniques, blade platforms require simple but robust cooling methods that are amenable to the high rotational loads. This study presents a platform cooling solution that utilizes a single impingement jet directed to the underside of the platform from the forward shank face. Detailed local distributions of heat transfer coefficients are presented for the cooled platform region formed by the cavity between the pressure side and suction side of two adjacent blade castings. Fundamental understanding is provided for impingement into this confined cavity region, with two differing jet positions. The effects of jet impingement distance to the surface from Z/D of 1.5 to 8, and jet Reynolds number from 65,000 to 155,000 are investigated. Furthermore, to augment the cooled surface heat transfer coefficients, surface roughness is applied by a patented method appropriate to actual turbine hardware. The single cooling jet produces a platform cooling distribution that may be treated as two distinct regions. The primary impingement region is seen to behave in line with expectations from prior literature, with some cavity geometry effects. Impingement heat transfer coefficients are increased by as much as 50% through the use of an applied randomly roughened surface that models an actual braze alloy method. This roughness benefit is however lost if the impingement cooling jet is not relatively close to the platform surface. The non-impingement heat transfer regions of the platform are affected only by the total cooling flow rate supplied by the impingement jet.