Abstract: A gas turbine engine rotor assembly including a rotor having a radially outer rim with an outer surface shaped to reduce circumferential rim stress concentration between each blade and the rim. Additionally, the shape of the outer surface directs air flow away from an interface between a blade and the rim to reduce aerodynamic performance losses between the rim and blades. In an exemplary embodiment, the outer surface of the rim has a concave shape between adjacent blades with apexes located at interfaces between the blades and the rim.

Abstract: A gas turbine engine includes a compressor rotor assembly which directs air at a sufficent pressure and temperature to a compressor rotor bore. The compressor assembly includes a compressor including a plurality of rotors joined with a plurality of couplings. Each of the rotors includes a radially outer rim, a radially inner hub, and a web extending therebetween. The web includes a flange including a plurality of slots which define a plurality of radial vanes that are airfoil-shaped. The vanes effectively deswirl bleed air and re-direct the airflow to a compressor bore.

Abstract: A combustor liner includes a wall having an outboard surface and an opposite inboard surface. A plurality of first holes are inclined through the wall in a multihole pattern to channel cooling fluid therethrough to form a cooling film layer along the inboard surface. A second hole extends perpendicularly through the wall within the multihole pattern to form a shadow along the inboard surface devoid of the first holes. A transition hole extends through the wall in the shadow at a greater inclination than the first holes for cooling the wall at the shadow.