Patent ID: 8292574
Filing Date: 2012-10-23
Classification: F01D,F04D,F05D,Y02T

Abstract:
1. A compression stage for a gas turbine engine comprising: a rotor stage having a plurality of rotor blades spaced circumferentially around a rotor hub with a longitudinal centerline axis; a stator stage having a plurality of stator vanes spaced in a circumferential direction coaxial with the longitudinal centerline axis; each stator vane having pressure and suction sides extending between leading and trailing edges, and longitudinally between a root and a tip, the trailing edge having an exit swirl angle distribution profile from the root to the tip such that the exit swirl angle increases between a first radius location and an intermediate radius location between the first radius location and the tip wherein the exit swirl angle has a maximum value at the intermediate radius location and remains substantially constant between the intermediate radius location and the tip; and each rotor blade having a blade leading edge shape adapted to receive the flow from a stator stage with the exit swirl angle distribution profile.