A fan case assembly adapted for use with a gas turbine engine includes a fan case and a fan track liner. The fan case that extends circumferentially about an axis. The fan track liner extends circumferentially at least partway about the axis and is coupled with the fan case. The fan track liner includes an abradable section and a core section located radially outward of the abradable section. The core section defines a triply periodic minimal surface geometry.

The invention relates to a hub (2) of an intermediate casing (1) for a bypass turbomachine comprising: a bleed stream duct (18), a bleed valve, comprising a mobile door at the inlet orifice to the bleed stream duct (18), a set of bleed vanes (22) which are mounted with the ability to rotate about a pivot (26) in the bleed stream duct (18) between an open configuration in which a flow of air coming from the inlet orifice (4) passes between the bleed vanes (22) and a closed configuration, the pivot (26) for each bleed vane (22) being closer to its leading edge (BA) than to its trailing edge (BF).

A flame resistant shield includes two opposing walls between which are positioned at least: a first insulating layer, the first layer being capable of distributing heat in the plane formed by the first layer and being insulating across its thickness, a second insulating layer, one of the opposing walls which covers the first layer being produced from a refractory antioxidant material or having at least the surface intended to be exposed to flames covered with a material preventing this surface from being oxidized, the other wall being a support.

The invention relates to a fan blade (3) of a turbomachine (15) comprising a structure (4) made of composite material having, within the root (5) and/or the stilt (6) of the blade, a first layer (11) comprising the lower face (16) and having a first thickness (E1) of between 10% and 25% of the total thickness (E), a second layer (12) comprising the suction face (17) and having a second thickness (E2) of between 10% and 25% of the total thickness (E), and a central layer (13) extending between the first layer (11) and the second layer (12), the shortening of the warp strands (9) in the first and second layers (11, 12) being greater than the shortening of the warp strands (9) in the central layer (13).

A thermally conductive lightweight elastomeric seal including a stator substrate having an external surface; a casing coupled to the external surface, the casing including radial walls extending orthogonal radially from the external surface; an abradable material disposed within the casing, the abradable material comprises an elastomer material with imbedded metal-coated hollow microspheres, wherein the abradable material comprises a density of 0.5 to 0.6 grams/cubic centimeter; the abradable material and the casing being coupled together.

A gas turbine engine including a component having a first and a second component that are attached or integrally formed with each other. The first and the second component define an enclosed void. An insert is positioned within the enclosed void which is formed of a material having a relatively low Young's modulus. The insert may prevent resin from filling the enclosed void during assembly/manufacture of the component, while also reducing a weight of the component and an amount of stress on the component. The component may be a fan platform positioned between a plurality of fan blades, an outlet guide vane, or another component of the gas turbine engine.

A turbocharger assembly can include a center housing; a turbine housing; and an insert disposed between the center housing and the turbine housing where a surface of the turbine housing and a first surface of the insert define a volute and where a surface of the center housing and a second surface of the insert define a chamber.

The present invention relates to a labyrinth seal for a turbine engine, in particular of an aircraft, comprising a rotor element and a stator element extending around the rotor element, the rotor element being suitable for rotating relative to the stator element about an axis of rotation having an axial direction (DA), the rotor element comprising an annular lip having an outer radial end extending towards an abradable element (57) carried by the stator element, the outer radial end of the annular lip having a corrugation in the axial direction (DA) and a non-zero axial expanse (E.sub.5) associated with the corrugation, the abradable element (57) comprising a plurality of cells (50a, 50b) arranged adjacent to one another along the axial direction (DA) and an ortho-radial direction (O), the cells (50a, 50b) comprising walls which extend in an essentially radial direction, the cells being distributed with a first cell density in a first densified annular zone (Z.sub.51) of the abradable element, said densified annular zone (Z.sub.51) being located opposite the radial end of the lip, said densified annular zone having an axial expanse less than or equal to the axial expanse of the outer radial end of the lip, the cells being distributed according to a reference density of cells outside said first zone, the first density being greater than the reference density.

An electric turbo compressor including a housing, an electric motor, and a rotating body driven to rotate by the electric motor. The rotating body includes a rotary shaft, a first impeller, and a second impeller. The first impeller includes a first hub fixed to the rotary shaft, and a plurality of first blades arranged on the first hub, and the second impeller includes a second hub fixed to the rotary shaft, and a plurality of second blades arranged on the second hub. In the rotating body, mass of part of the rotating body from one ends to the other ends of the first blades along the axial direction is greater than mass of part of the rotating body from one ends to the other ends of the second blades along the axial direction.