Platform for an aircraft turbo machine fan rotor, the platform being configured to be secured to a fan disc between two adjacent fan blades. The platform further including a longitudinal wall defining an aerodynamic external face. The wall includes a honeycomb structure interposed between two skins which are respectively an internal skin and an external skin, with the external skin defining the aerodynamic external face.

A fan track liner for a fan containment arrangement for a gas turbine engine comprises a cellular impact structure and a supporting sub-laminate integrally formed with each other from a fibre-reinforced polymer material.

An airfoil assembly for a gas turbine engine includes an airfoil extending radially relative to a central reference axis. A spar is located within the airfoil and spaced from the airfoil at all radial locations along the airfoil such that a gap is maintained between the airfoil and the spar. A foam is located between the airfoil and the spar.

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.

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 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 yam extending through two parts of the shell adjacent the trailing edge, wherein the yarns do not extend through the trailing edge insert.

An air intake structure for a nacelle of an aircraft is disclosed having an air intake lip with a U-shaped section open towards the rear, an acoustic panel that extends the air intake lip towards the rear and on an inner side, an outer panel extending the air intake lip towards the rear and on an outer side, and a rear strengthening frame fixed between the outer panel and the acoustic panel, where the rear strengthening frame partly defines an inner volume that is immediately in front of the rear strengthening frame. The air intake structure comprises an impact absorber element positioned in the inner volume and the impact absorber element assumes the form of a structure that is deformable during an impact and that is at least partially filled with a fluid. The impact absorber element is a flexible shroud filled with a pressurized gas.

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.

Provided is a method of shaping a composite blade made of a composite material by curing prepreg in which reinforcing fibers are impregnated with resin. A foaming agent disposed in an internal space of the composite blade contains a plurality of foaming bodies and foaming agent resin. The foaming bodies foam by being heated. The foaming agent resin cures by being heated. The foaming bodies include low-temperature side foaming bodies and high-temperature side foaming bodies. The low-temperature side foaming bodies foam in a low temperature range during a curing step. The high-temperature side foaming bodies foam in a high temperature range corresponding to temperatures higher than the low temperature range during the curing step.

A rotor blade for a gas turbine engine including an inner ply layer group includes a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter.