An airfoil for a gas turbine engine defining a spanwise direction, a root end, a tip end, a leading edge end, and trailing edge end is provided. The airfoil includes: a body extending along the spanwise direction between the root end and the tip end, the body formed of a composite material; and a sculpted leading edge member attached to the body positioned at the leading edge end of the airfoil, the sculped leading edge member formed at least in part of a metal material and defining a non-linear patterned leading edge of the airfoil.

Fans having fan blades with lightweight blade tips are disclosed herein. In one embodiment, for example, each fan blade includes a blade tip attached to an end portion of a main airfoil. The blade tip includes a shaped body that is molded over a support structure to form the shape and exterior surface of the blade tip.

In one embodiment, a gas turbine engine component includes a foam based core and a composite skin member. Both the foam based core and the composite skin member can be used to structurally support the gas turbine engine component. The composite skin member can be a CMC material and is used to partially encapsulate the foam core. The gas turbine engine component can take the form of an airfoil member such as a blade or a vane, a combustor liner, etc. A first portion of the composite skin member includes a first surface extending past an edge of the component creating a step approximate an edge section. In another embodiment, composite skin members can be used to form a continuous shape for the edge section such that the foam core forms part of a gas path surface.

A fan track liner for a fan containment arrangement for a gas turbine includes a cellular impact structure and a supporting sub-laminate integrally formed with each other from fibre-reinforced material. The fan track liner further includes a ballistic barrier comprising a woven reinforcing fibre ply and a layer of reinforcing fibre felt. There is also disclosed an associated fan containment arrangement and a method of manufacturing a fan track liner.

A gas turbine engine has: a first component and a second component defining a respective first gaspath surface and a second gaspath surface of an annular gaspath, the first and second gaspath surfaces axially spaced apart from one another by an annular recess in the first component; a bushing ring disposed within the annular recess and defining stem pockets therein; variable guide vanes pivotable about respective vane axes extending between first and second stems; and a biasing member received within the annular recess and disposed axially between the bushing ring and one of the first component and the second component, the biasing member exerting a force against the bushing ring in an axial direction relative to the central axis and towards the other of the first component and the second component.

A gearbox assembly for a gas turbine engine includes a hollow gear. The hollow gear has one or more reservoirs configured to store a lubricant and one or more passages configured to supply the lubricant to one or more gear meshes in the gearbox assembly. The hollow gear provides a dampening system configured to absorb vibrations and loading experienced in the gearbox assembly and configured to provide impact resistance for the gearbox assembly.

An airfoil for a gas turbine engine defining a spanwise direction, a root end, a tip end, a leading edge end, and trailing edge end is provided. The airfoil includes: a body extending along the spanwise direction between the root end and the tip end, the body formed of a composite material; and a sculpted leading edge member attached to the body positioned at the leading edge end of the airfoil, the sculped leading edge member formed at least in part of a metal material and defining a non-linear patterned leading edge of the airfoil.

A propeller blade comprises a root, a tip distal from the root, a trailing edge extending from the root to the tip, a trailing edge, e.g. foam, insert, a shell forming an outer surface of the propeller blade and a plurality of stitches of yarn extending through two parts of the shell adjacent the trailing edge, wherein the yarns do not extend through the trailing edge insert.

A machine includes a section that defines a target vibrational mode to dampen and a nanocellular foam damper that includes interconnected ligaments in a cellular structure. The interconnected ligaments have an average ligament size defined with respect to a vibrational loss modulus of the nanocellular foam damper and the target vibrational mode. Also disclosed is a method of damping vibration.

An airfoil assembly comprising a central spar extending along the radial direction; a blade skin positioned around the central spar to define an airfoil that has a pressure side and a suction side extending in the axial direction between a leading edge and a trailing edge, wherein at least one cavity is defined between the blade skin and the central spar; and a support structure positioned at least partially within the at least one cavity. The support structure comprises: a foam; and a foam reinforcement structure embedded within the foam.