The present invention relates to a fan blade (3) having a structure made from a composite material, comprising a fibrous reinforcement (5), which is obtained by three-dimensional weaving of warp strands and weft strands, and a matrix in which the fibrous reinforcement (5) is embedded, wherein—the fibrous reinforcement (5) comprises a first portion (14) forming the trailing edge (9) of the structure made from a composite material and a second portion (15) forming its leading edge (8), and wherein—the warp strands of the fibrous reinforcement (5) comprise first strands (12) having a predetermined stiffness and second strands (13) having a greater stiffness than that of the first strands (12), the first portion (14) comprising all or part of the first strands (12) and being devoid of second strands (13) while the second portion (15) comprises all or part of the second strands (13).

A method of monitoring turbine blade creep in a gas turbine engine is provided. The method includes: receiving stereo images of a turbine blade of a row of turbine blades, the images having been obtained using a stereo borescope located in the engine adjacent the row of turbine blades; identifying same features of the blade in each of the stereo images; mapping each of the identified features by triangulation onto a 3D space to produce a 3D depth map of at least part of the blade; providing a 3D reference model of the blade; and comparing the 3D reference model with the 3D depth map to measure one or more deviations in shape of the blade to determine an amount of creep-induced distortion of the blade.

A blade element (4, 8) for a comminution device (1) to comminute fiber material has at least one comminution section (22) with comminution parts (20, 24, 25, 26) and free spaces (21) therebetween, and at least one feed section (23) extending at least partly in a longitudinal direction (X) of the blade element (4, 8), each feed section (23) intended to feed fiber material to the respective comminution section (22). The comminution parts have a first dimension (d20a, d20b, d20c, d24a, d24b, d24c) extending in a circumferential direction (C) of the blade element and a second dimension (e20a, e20b, e20c, e24a, e25a, e26a) extending in the longitudinal direction (X) of the blade element. At the same longitudinal (X) position in the blade element (4, 8) the first dimension of the comminution parts is arranged to increase in the circumferential direction (C) of the blade element toward the feed section.

A rotor blade includes an airfoil having an airfoil shape. The airfoil shape has a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances expressed in a unit of distance by multiplying the Cartesian coordinate values of X, Y and Z by a scaling factor of the airfoil in the unit of distance. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each Z value. The airfoil profile sections at Z values are joined smoothly with one another to form a complete airfoil shape.

Proposed is an impeller-type exhaust apparatus for blocking transfer of heat including a first disc facing a high-temperature gas, a second disc spaced apart from the first disc to form a heat blocking space, a plurality of blades formed along the circumferences of the first disc and the second disc, and a motor which provides rotary power to the first disc and the second disc, wherein heat can be blocked through the heat blocking space.

A turbine blade tip shroud has a first cutter tooth extending from a tip rail from one of the upstream side and the downstream side of the tip rail and adjacent the leading edge of the body. The tip shroud also includes a second cutter tooth extending from the tip rail from the other side of the tip rail at a position axially distant from the first cutter tooth. The cutter teeth are thus axially offset. The tip shroud can be initially manufactured with this shape or may be modified from a used tip shroud having, for example, opposing cutter teeth near a leading edge of a body of the tip shroud. Various tip shroud surface profiles, which are expressed in terms of Cartesian coordinates, are also provided.

A component for a turbine engine is provided. The component includes a main body portion having a flow path surface; and a protective layer formed of a chopped fiber material, the protective layer cohered to at least a portion of the flow path surface of the main body portion.

An airfoil for a gas turbine engine defining a radial direction and an axial direction, the airfoil including: a flared portion extending from a suction surface of the airfoil at an outer edge of the airfoil along the radial direction; and a plenum disposed at the outer edge of the airfoil, the plenum having a suction-side sidewall and a pressure-side sidewall, wherein the suction-side sidewall has a first sidewall portion adjacent to a second sidewall portion disposed outside of the first sidewall portion along the radial direction, wherein the first sidewall portion defines a first angle, wherein the second sidewall portion defines a second angle, and wherein the first angle is greater than the second angle.

A set of compression stage(s) of a turbomachine, forming an annular fluid passage and comprising at least one annular stator platform and/or at least one annular rotor platform having an outer longitudinal profile inclined (I.sub.Si/I.sub.Ri) inwards and upstream with respect to a nominal profile of the fluid stream, where the inclination (I.sub.Si/I.sub.Ri) of the outer longitudinal profile of the or each of the annular platforms, relative to the nominal profile of the fluid stream, oscillates along the circumference of the annular platform or platforms, between a maximum value in front of the blades of the annular platform and a minimum value between each pair of adjacent blades of the annular platform.

The invention relates to a composite blade (5) for a turbine engine rotor, for example, an unducted propeller, comprising a skin (6) made of woven fibres forming the outer profile of the blade and an attachment (11) with cylindrical geometry in the direction of the span emerging from the blade root (9), which attachment is intended to retain the blade (5) on a hub of the rotor, characterised in that it further comprises a spar (16) having a hollow tube structure made of braided carbon fibres, fixed to the attachment (11) and extending inside the skin (6) over at least part of the span of the blade (5). The invention also relates to a propeller comprising said blade and to a method for manufacturing said blade.