Abstract: A gas turbine engine airfoil extending from an airfoil base to an airfoil tip at a free end of the airfoil and from a leading edge to a trailing edge of the airfoil has at least one recess extending into and circumferentially completely through the airfoil tip and located inwardly of the leading and trailing edges. The recess is located in an area of the airfoil tip subject to high tip vibratory stress and or rubs and may be a circular scallop having a scallop radius and a maximum depth of a few mills in a range of about 0.005-0.01 inches as measured from a nominal tip edge without the recess. The airfoil may be on a radial blade, a stator vane, or an impeller blade. An annular tip seal may surround a plurality of such airfoil tips with an airfoil tip clearance therebetween.

Abstract: Including a disk, a plurality of circumferentially spaced apart airfoils extending radially outwardly from the disk, a plurality of shroud segments being disposed circumferentially between the plurality of airfoils and configured to attach to a corresponding airfoil of the plurality of airfoils and a plurality of wear surfaces, where at least one wear surface of the plurality of wear surfaces is configured and disposed to prevent excessive wear between the adjacent shroud segments. Each shroud segment of the plurality of shroud segments may be attached to a corresponding airfoil by weld connections to permit clearance during translation friction welding of the airfoil to the disk hub are located to avoid high stress areas.

Abstract: Including a disk, a plurality of circumferentially spaced apart airfoils extending radially outwardly from the disk, a plurality of shroud segments being disposed circumferentially between the plurality of airfoils and configured to attach to a corresponding airfoil of the plurality of airfoils and a plurality of wear surfaces, where at least one wear surface of the plurality of wear surfaces is configured and disposed to prevent excessive wear between the adjacent shroud segments. Each shroud segment of the plurality of shroud segments may be attached to a corresponding airfoil by weld connections to permit clearance during translation friction welding of the airfoil to the disk hub are located to avoid high stress areas.

Abstract: A method enables a rotor blade for a gas turbine engine to be fabricated. The method comprises forming an airfoil including a first side wall and a second side wall that each extend in radial span between an airfoil root and an airfoil tip, and wherein the first and second side walls are connected at a leading edge and at a trailing edge, and forming a winglet that extends outwardly from at least one of the airfoil first side wall and the airfoil second side wall, such that a radius extends between the winglet and at least one of the airfoil first side wall and the second side wall.

Abstract: A fan blade includes perforated anchor tangs, one with inboard corners and the remainder without. The corners are disposed in a complementary seat in a rotor disk for limiting blade pivoting during assembly.

Abstract: A fan outlet guide vane includes integral outer and inner platforms conforming to arcuate sides thereof, and multiple mounting bosses on the platforms. The platform sides are arcuate to conform with the vane sides for improving aerodynamic efficiency. The multiple mounting bosses improve mounting of the vanes to corresponding supporting rings.

Abstract: A fan blade includes an airfoil having a root and an integral platform. The platform includes forward and aft ends and arcuate first and second sides therebetween. The airfoil root is laterally offset closer to the platform first side than the second side. Impact loads with a trailing blade during a blade-out failure are accordingly distributed for reducing peak loading in the trailing blade and damage thereto.