A vane for a turbomachine includes a pressure surface and a suction surface opposite the pressure surface. The pressure surface and the suction surface define a width therebetween. The vane also includes a first end. The first end includes a distal portion, a proximal portion, a pressure surface first portion, and a suction surface first portion. At least one of the pressure surface first portion and the suction surface first portion slope away from the other of the pressure surface first portion and the suction surface first portion such that the width increases from a first end minimum width at the proximal portion to a first end maximum width at the distal portion.

An airfoil includes an airfoil body having a first wall, a second wall, a third wall, a tip surface, and a rib. The first wall radially extends between a root region and a tip region and axially extends between a leading edge and a trailing edge. The second wall radially extends from the tip region towards the root region and axially extends between the leading edge and the trailing edge. The third wall radially extends between the root region and the tip region and axially extends between the leading edge and the trailing edge. The tip surface circumferentially extends between the second wall and the third wall. The rib is radially spaced apart from the tip surface and circumferentially extends between the first wall and the third wall.

An airfoil assembly for a turbine engine, comprising at least one airfoil having a leading edge and a trailing edge, a band having an inner side and an outer side and rigidly coupled to the at least one airfoil along a portion of an interface between the band and the at least one airfoil for providing at least a portion of support for the at least one airfoil, a relief located in the band at the leading edge or the trailing edge of the at least one airfoil and defining a stress relief gap between the band and the leading edge or trailing edge and a closure preventing airflow through the relief.

Gas turbine engine (GTE) airfoils, such as rotor and turbofan blades, having multimodal thickness distributions are provided. In one embodiment, the GTE airfoil includes an airfoil tip, an airfoil root opposite the airfoil tip in a spanwise direction, and first and second airfoil halves extending between the airfoil tip and the airfoil root. The first airfoil half has a first multimodal thickness distribution, as taken in a cross-section plane extending in the spanwise direction and in a thickness direction substantially perpendicular to the spanwise direction. The first multimodal thickness distribution may be defined by multiple locally-thickened airfoil regions, which are interspersed with multiple locally-thinned airfoil regions. The second airfoil half may or may not have a multimodal thickness distribution. By imparting at least one airfoil half with such a multimodal thickness distribution, targeted mechanical properties of the GTE airfoil may be enhanced with relatively little impact on aerodynamic performance.

A rotor assembly having a plurality of scoops disposed in a circumferential array, the scoops extending from an inner surface of the outer wall of the flow path along a radial distance smaller than a radial distance between the inner and outer walls of the flow path. Each of the scoops forms a closed channel from an inlet to an outlet with the inlet and outlet being axially spaced from one another, the outlet being upstream of and adjacent the annular blade path. A gas turbine engine and method of reducing tip vortices in a rotor assembly are also discussed.

A blade or vane has: an airfoil having an inner diameter (ID) end and an outer diameter (OD) end and having a suction side and a pressure side and a leading edge and a trailing edge; and an inner platform and/or attachment root at the ID end and/or an outer platform at the OD end. At least one of the inner platform, root, and/or outer platform comprises one or more pieces of a first alloy. One or more pieces of a second alloy form a leading edge section of the airfoil. One or more pieces of a third alloy form a trailing edge section of the airfoil. One or more pieces of a fourth alloy form a spar of the airfoil between the leading edge section and trailing edge section and extending into said at least one of the inner platform, root, and/or outer platform.

Structures for reducing forward loads in compressors are described. compressor includes inner and outer circumferential support structures positioned concentrically around a central axis, and an aft-most stage including a vane extending radially inward from the outer circumferential support structure. An axial length of the aft-most stage is defined by a spacer arm of the inner circumferential support structure. The vane includes a root, a tip, and a trailing edge extending between the root and the tip. A ratio of a first radial distance between a first point located at an intersection of the tip of the vane and the trailing edge to a radially inner wall of the spacer arm of the inner circumferential support structure to a second radial distance between the first point and a second point located at an intersection of the root of the vane and the trailing edge is between 0.95 and 4.5.

The present disclosure is directed to a method for scaling an airfoil for placement in a turbomachine. The method disclosed herein includes radially scaling a master airfoil to form a scaled airfoil. The method may also include tuning the scaled airfoil. For example, tuning the scaled airfoil may include axially scaling. The scaled airfoil generally has similar characteristics to the master airfoil.

A turbomachine blade of the type having a metal lower coupling root, a metal upper coupling head, and a metal airfoil-shaped oblong member designed to connect the coupling root rigidly to the coupling head; the airfoil-shaped oblong member having a substantially airfoil-shaped main plate-like element connected to the coupling root and to the coupling head, and which is divided into: a lower connecting fin cantilevered from and formed in one piece with the coupling root; an upper connecting fin cantilevered from and formed in one piece with the coupling head; and a center plate-like body, which is located between the lower and upper connecting fins, is shaped/designed to form an extension of the lower and upper connecting fins, and is butt-welded to, to form one piece with, the lower and upper connecting fins.

An assembly for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, an airfoil including a radial end, a first passageway having an outlet at the radial end, and a second passageway having an inlet at the radial end. The assembly further includes a cover having at least one turning cavity configured to direct fluid expelled from the outlet of the first passageway into the inlet of the second passageway.