Patent ID: 8834123
Filing Date: 2014-09-16
Classification: F01D,F05D

Abstract:
1. A compressor blade for a gas turbine engine, comprising: an airfoil extending between a leading edge and a trailing edge and operable to affect a change in total pressure between an upstream side of the airfoil and a downstream side of the airfoil; a stalk having a lower attachment portion and an upper portion defining a platform, the platform having a first side and a second side, a portion of the first side of the platform being coupled to the airfoil, the platform extending between first and second opposite ends, the platform including a leading edge having two corners at the upstream side of the airfoil at the respective first and second opposite ends of the platform, and a trailing edge having two corners at the downstream side of the airfoil at the respective first and second opposite ends of the platform; and only first and second single corner undercuts,