Methods of bonding first structures to second structures are disclosed wherein the first and second structures are fabricated materials having different physical characteristics. For example, the first structure may be a composite fan blade and the second structure may be a composite or metallic rotor, both for use in gas turbine engines. The method includes providing the first and second structures and plating or otherwise coating a portion of the first structure with a metal to provide a metal-coated portion. The method includes applying at least one intermediate material onto the metal-coated portion of the first structure. The method further includes bonding the metal-coated portion of the first structure and the intermediate material to the second structure. The bonding is carried out using a relatively low-temperature process, such as liquid phase bonding, including TLP and PTLP bonding. Brazing is also a suitable technique, depending on the materials chosen for the first and second structures.

The invention relates to a method for manufacturing a composite platform (30) for a fan of an aircraft turbine engine, wherein said platform comprises an elongate wall (32) and is configured to extend between two fan blades (3), said wall comprising an outer aerodynamic surface (32a) and an inner surface (32b), on which a fastening tab is located (34), said fastening tab being configured to be fixed to a fan disc (2). The invention is characterized in that it comprises the steps of: a) preparing fabrics or sheets which are pre-impregnated with a resin, b) depositing the fabrics or sheets in a mold, c) positioning a metal reinforcement (36) in the mold on the fabrics or sheets, the reinforcement being integrally formed with said fastening tab, d) depositing the fabrics or sheets on a part of the reinforcement, and e) closing and heating the mold for solidification of the assembly formed by the fabrics or sheets and the reinforcement.

The invention relates to a method for manufacturing a composite platform (30) for a fan of an aircraft turbine engine, wherein said platform comprises an elongate wall (32) and is configured to extend between two fan blades (3), said wall comprising an outer aerodynamic surface (32a) and an inner surface (32b), on which a fastening tab is located (34), said fastening tab being configured to be fixed to a fan disc (2). The invention is characterized in that it comprises the steps of: a) preparing fabrics or sheets which are pre-impregnated with a resin, b) depositing the fabrics or sheets in a mold, c) positioning a metal reinforcement (36) in the mold on the fabrics or sheets, the reinforcement being integrally formed with said fastening tab, d) depositing the fabrics or sheets on a part of the reinforcement, and e) closing and heating the mold for solidification of the assembly formed by the fabrics or sheets and the reinforcement.

The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.

A composite platform for a fan of an aircraft turbine engine. The platform includes an elongate wall and is configured to extend between two fan blades. The wall includes an aerodynamic outer surface and an inner surface, on which a fastening tab is located, wherein the fastening tab is configured to be attached to a fan disc. The fastening tab is integrally formed with a metal reinforcement which has a plate having an elongate shape and which extends over more than 50% of the longitudinal extent of the wall, the wall being produced by overmolding a resin on the plate so as to be integrated into the wall.

An impeller includes a metallic inducer portion and a polymeric exducer portion connected to the metallic inducer portion. The metallic inducer portion includes an inducer hub, inducer blades attached to the inducer hub, and an inducer coupling on an end of the inducer hub. The polymeric exducer portion includes an exducer hub, exducer blades attached to the exducer hub, and an exducer coupling on an end of the exducer hub. The exducer coupling connects to the inducer coupling.

An airfoil member (100) comprising has a substrate (120) along at least a portion of an airfoil (102) of the airfoil member. A sheath (122) has a channel (144) receiving a portion (160) of the substrate. A plurality of separate spacers (320; 380; 400) are between the sheath and the substrate and have a plurality of gaps between the spacers.

A method of fabricating a fan case for a gas turbine engine defines a metallic ring including an outer surface and an inner surface. A first composite material is assembled about the outer surface of the metallic ring. A second composite material is assembled about the first composite material. The first composite material and the second material are cured about the metallic ring within a tool to form a first subassembly. The first subassembly is removed from the tool. A fan case assembly for a gas turbine engine and a gas turbine engine are also disclosed.

A compressor wheel that can be employed in devices such as turbochargers. The compressor wheel includes an axially extending hub having an inlet end, a shaft bore extending from the inlet end and an arcuate outer surface opposed to the shaft bore. The axially extending hub is composed of a metal and has a porous region located proximate to the arcuate outer surface of the axially extending hub. The compressor wheel also includes a blade array disposed on the arcuate outer surface of the axially extending hub. The blade array has an outer surface and an inner region. The blade array comprises a plurality of circumferentially-spaced, radially and axially extending blades disposed thereon and is composed, at least in part of a polymeric material. Polymeric material located in the inner region of the blade array extends into the porous region defined in the axially extending hub.

An impeller includes a metallic inducer portion and a polymeric exducer portion connected to the metallic inducer portion. The metallic inducer portion includes an inducer hub, inducer blades attached to the inducer hub, and an inducer coupling on an end of the inducer hub. The polymeric exducer portion includes an exducer hub, exducer blades attached to the exducer hub, and an exducer coupling on an end of the exducer hub. The exducer coupling connects to the inducer coupling.