Patent ID: 6629817
Filing Date: 2003-10-07
Classification: F01D,Y10T

Abstract:
A method for contouring an airfoil for a gas turbine engine to facilitate improving film cooling effectiveness of the airfoil, the airfoil including a leading edge, a trailing edge, a first sidewall, and a second sidewall, the first and second sidewalls connected chordwise at the leading and trailing edges to define a cavity, the first and second sidewalls extending radially between an airfoil root to an airfoil tip, said method comprising the steps of:forming an inflection in an outer surface of at least one of the airfoil first sidewall and the airfoil second sidewall, such that the inflection extends a distance radially between the airfoil root and the airfoil tip and such that the inflection is a distance downstream from the stagnation line of the leading edge of the airfoil, wherein the inflection is defined between a concave surface and a convex surface; and forming at least one opening immediately adjacent the inflection and through at least one of the convex surface and the concave surface immediately upstream of the inflection point of the inflection for receiving cooling fluid therethrough from the airfoil cavity to the airfoil outer surface wherein said opening through the airfoil sidewall is at an injection angle measured with respect to the airfoil outer surface that is less than about 16 degrees.