An example method for vane classification includes scanning, using a structured light scanner, a vane for a turbine engine to capture three-dimensional (3D) data about the vane. The method further includes generating a point cloud from the 3D data about the vane. The method further includes connecting, using a processing system, points of the point cloud to generate a mesh surface. The method further includes determining, using the processing system, an airflow for an airfoil of the vane based at least in part on the mesh surface. The method further includes constructing the turbine engine based at least in part on the airflow for the airfoil of the vane without reference to an adjacent airfoil of the vane.

Fan or propeller vane (1) for an aircraft turbomachine, the vane being made from a composite material and comprising a blade (2) and a base (3), the base being formed by a longitudinal end (41) of a spar (4) which is formed by a fibrous reinforcement formed from threads woven in three dimensions and a portion (42) of which extends inside the blade (2), the blade (2) having an aerodynamic profile which is defined by a skin (5) which is formed by woven threads and which surrounds the portion of the spar, the spar (4) and the skin (5) being embedded in a polymerised resin, characterised in that the portion (42) of the spar comprises projecting longitudinal stiffening members (6) which together delimit spaces (8) for receiving longitudinal inserts (7) which are formed from a honeycomb material.

A method for manufacturing a turbine blade comprising designing a turbine blade includes receiving initial geometrical and aerodynamic information of the turbine blade, obtaining the maximum amount of stress within a determined area of maximum stress, and obtaining a safety factor by dividing material yield stress of the turbine blade by the maximum amount of stress. The method further includes performing a first plurality of operations responsive to the safety factor being less than 1.5 and the determined area of maximum stress occurring at the junction of the blade airfoil and the blade root. The first plurality of operations includes creating a fillet at the junction of the blade airfoil and the blade root and increasing respective thickness of each airfoil slice of the plurality of airfoil slices with a distance from the junction of the blade airfoil and the blade root of no more than 15% of the blade airfoil length.

An airfoil includes an airfoil body that has a geometrically segmented coating section. The geometrically segmented coating section includes a wall having an outer side. The outer side has an array of cells, and there is a coating disposed in the array of cells.

A component for a gas turbine engine, including: at least one internal cavity extending through the component, the internal cavity having at least one inlet opening and at least one outlet opening each being in fluid communication with the at least one internal cavity; a plurality of cooling features extending from a surface of the at least one internal cavity, the plurality of cooling features are formed in accordance with at least one of the following groups: i) a plurality of airfoil shaped features that extend upwardly from the surface of the at least one internal cavity and a plurality of wedge shaped features each having a triangular base that has an upstream portion and a downstream portion, the upstream portion extending further from the surface than the downstream portion; ii) a plurality features having a curved or J shaped base that extends upwardly from the surface, a plurality features having a double curved or symmetrically J shaped base that extends upwardly from the surface, and a plurality features having a base that extends upwardly from the surface with a curved portion that defines an opening therethrough; iii) a first plurality of pins with a plurality of grooves that are formed into a peripheral surface of each of the first plurality of pins and a second plurality of pins with a plurality of grooves that are formed into a peripheral surface of each of the second plurality of pins the plurality of grooves formed in the peripheral surface of each of the second plurality of pins are configured such that V shapes or inverted V shapes are formed in the peripheral surface of each of the second plurality of pins; and iv) a plurality of chevron shaped trip strips that are located in a channel, the plurality of chevron shaped trip strips are spaced from each other such that a U shaped passage is formed therebetween and each chevron shaped trip strip has a top portion that curls inwardly towards the channel and a plurality of pairs of features that each extend from a surface of another channel towards each other where a gap is located between distal ends of the plurality of pairs of features.

An airfoil assembly for a turbine engine comprising an outer band, an inner band radially spaced inwardly from the outer band to define an annular region, and multiple airfoils circumferentially spaced within the annular region. Each corresponding airfoil of the multiple airfoils can project from a surface at a root and can further include an outer wall defining a pressure side and a suction side. A projection can extend upwardly from the surface on the pressure side and a valley can extend into the surface on the suction side to define a contour in the surface.

An airfoil includes an endwall section and an airfoil section that defines, at least in part, an airfoil profile. At least one of the airfoil section or the endwall section includes a seal cavity, and a seal is disposed in the seal cavity.

A method for producing a metal bladed element of a turbine engine, in particular of an aircraft, includes steps of producing the bladed element, depositing a coating made of wear-proof material on at least one portion of the bladed element and verifying, preferably visually, the conformity of the bladed element. Verifying the conformity of the bladed element includes implementing a verification element on the bladed element. The bladed element is configured according to a conformity threshold value to conceal a non-conformity of the coating, if the non-conformity has at least one dimension less than the threshold value, and to show at least one portion of this non-conformity if the at least one dimension is greater than the threshold value.

A hollow turbine airfoil or a hollow turbine casting including a cooling passage partition. The cooling passage partition is formed from a single crystal grain structure nickel based super alloy, a cobalt based super alloy, a nickel-aluminum based alloy, or a coated refractory metal.

A process for manufacturing a turbomachine blade of the type having at least one 3D cavity, wherein the blade is produced by a succession of depositions and selective consolidations of layers of a metal additive manufacturing powder based on an alloy of copper and nickel, the alloy including from 2% to 7% of nickel. It also relates to a turbomachine blade, wherein it is manufactured by metal additive manufacturing using the process.