Composite airfoils and methods for forming composite airfoils are provided. For example, a composite airfoil of a gas turbine engine comprises opposite pressure and suction sides extending radially along a span from a root to a tip, which define opposite radial extremities of the airfoil. The composite airfoil further comprises a body section and a tip section, which includes the tip, that each extend radially along the span. The composite airfoil is formed from a composite material comprising fibers disposed in a matrix material. The tip section has a tip fiber volume, and the body section has a body fiber volume that is greater than the tip fiber volume. Another composite airfoil comprises a tip cap applied over the tip that tapers from a first end to a second end such that each of the pressure and suction side walls of the tip cap narrows from a first thickness to a second thickness.

A centrifugal fan includes a housing and an impeller. The housing includes a sidewall, and the sidewall includes a tongue portion. The impeller includes a fan hub and a plurality of blades. The fan hub is rotatably disposed in the housing, and the tongue portion has an inner contour line on a reference plane. The blades connect to the fan hub. Each one of the blades has an end surface facing the sidewall. The end surface has an outer contour line on the section of the blade. Any two adjacent blades have different outer contour lines. The outer contour line of at least one first blade of the plurality of blades is parallel to the inner contour line. The outer contour line of at least one second blade of the plurality of blades is not parallel to the inner contour line.

An abrasive coating system for a substrate of an airfoil in a turbine engine high pressure compressor, comprising a plurality of grit particles adapted to be placed on a top surface of the substrate; a matrix material bonded to the top surface; the matrix material partially surrounds the grit particles, the matrix material consisting of unalloyed chromium and unalloyed aluminum distributed throughout the matrix material, wherein the grit particles extend above the matrix material relative to the top surface; and a film of oxidant resistant coating applied over the plurality of grit particles and the matrix material.

A fan including a hub and a plurality metal blades is provided. Each of the blades extends from the hub and is inclined relative to a radial direction of the hub. Each blade has a distal edge away from the hub, and has a pair of wingtips at the distal edge.

An impeller for a centrifugal turbomachine includes: a hub having a small-diameter portion positioned at a first end portion in an axial direction and a large-diameter portion positioned at a second end portion in the axial direction, the large-diameter portion having a greater diameter than the small-diameter portion; and a blade having a first edge positioned at an axial-directional position of the small-diameter portion and a second edge positioned at an axial-directional position of the large-diameter portion, the blade being disposed on an outer peripheral surface of the hub. The impeller is configured such that, when a first radial-directional cross section is a cross section of the impeller at an axial-directional position passing a tip of the first edge, at least a part of the first radial-directional cross section in a blade-height range of 50% or more is inclined downstream in a rotational direction of the impeller with respect to a radial direction.

The invention relates to a turbomachine rotary-fan blade having a predetermined breaking zone, which extends from the upstream edge along a given length and from the blade-tip edge over a given height. According to the invention, the body is made of a composite material comprising a fibre reinforcement obtained by three-dimensional weaving of warp and weft strands, and a resin matrix in which the fibre reinforcement is embedded, and has, in or in the vicinity of the zone, a discontinuity of at least some of the strands, configured such that the zone partially detaches when there is tangential friction in the thickness direction against the blade-tip edge, the height being less than 3% of the aerodynamic stream height of the blade.

The present invention relates to a method for coating a component, wherein the component has a first and a second surface, and wherein the first and the second surface adjoin each other at an edge, in which method i) first of all, the edge between the first and the second surface is rounded, and ii) subsequently, a coating is applied to the first surface.

A cooling assembly comprises a coolant source chamber inside an airfoil that directs coolant inside the airfoil that extends between a hub end and a tip end that includes a tip body and tip rail along a radial length. A first body cooling chamber and a second body cooling chamber are disposed inside the tip body. The second body cooling chamber is positioned between the tip end and the first body cooling chamber. At least one of the first or second body cooling chambers are fluidly coupled with the coolant source chamber. The coolant source chamber directs the coolant into the first or second body cooling chambers. A rail cooling chamber disposed inside of the tip rail is fluidly coupled with the first or second body cooling chambers. The first or second body cooling chambers directs coolant out of the body cooling chambers and into the rail cooling chamber.

A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft; and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edge, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein a ratio of a leading edge radius of each of the plurality of first rotor blades at 0% span height to a minimum leading edge radius is comprised between 1 and 1.50.

An apparatus and method for an engine component for a turbine engine having a working airflow separated into a cooling airflow and a combustion airflow. The engine component including a wall defining an interior and having an outer surface. A tip wall spanning first and second sides of the wall to close the interior. A tip rail extending from the tip wall and having an inner tip rail surface, which in combination with the tip wall, at least partially bounds a region defining a plenum. A rim formed in at least one of the outer surface and inner tip rail surface.