A method for manufacturing a fiber-reinforced composite material molded article of the present invention includes a step of causing a reinforcing fiber base material to undergo deformation with use of a mold, the reinforcing fiber base material including: reinforcing fibers which are unidirectionally oriented; and a matrix resin composition. The reinforcing fiber base material has a cut in a zone which is to undergo shear deformation and/or compressive deformation, the cut being substantially parallel to an orientation direction of the reinforcing fibers, and has substantially no cut that cuts through a fiber.

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.

This document describes systems and processes for forming improved synthetic molded slabs suitable for use in living or working spaces (e.g., along a countertop, table, floor, or the like).

Some embodiments of the present disclosure relate to an apparatus including an additively manufactured node having a socket. The apparatus includes a panel interconnected with node. The panel includes opposing surface layers and a core between at least a portion of the surface layers. The socket engages an end portion of the panel and shapes the surface layers on the end portion of the panel.

Layers of plies of composite material are laid on a convex tool surface. A first layer is placed with a first ply of a first section having a gap edge adjacent a gap edge of a first ply of a second section, the edges being parallel and a contraction distance from each other. A second layer is placed with a second ply of the first section having a gap edge adjacent a gap edge of a second ply of the second section, the edges being parallel and a contraction distance from each other, the second ply of the second section overlapping onto the first ply of the first section by a splice distance. Consolidation and curing cause contraction of the layers toward the tool, allowing the adjacent gap edges of each layer to be in close proximity or in contact after moving toward each other during the contraction.

An aircraft wing (6) is provided. The aircraft wing (6) comprises at least one structure (62), (64), 66 comprising: a foam core (626), (666); first and second carbon fibre composite layers (624a), (622a), (662a), (664a) respectively attached to top and bottom sides of the foam core to sandwich the foam core; and third and fourth carbon fibre composite layers (624b), (622b), (662b), (664b) respectively disposed adjacent to the first and second carbon fibre composite layers, wherein the total thickness of the structure is between 1 mm and 11 mm. An aircraft having the aircraft wing and a method of manufacturing a structure are also provided.

A method for fabricating a central caisson of an aircraft wing including a frame formed from U-section crossmembers. Formation of at least one crossmember includes forming a stack comprising at least one layer of resin film applied to a layer of dry fibers, pressing this stack to confer on it a U-section form and such as to compress the layers thereof, and pre-polymerization and/or polymerization of the resin of the stack. The manipulated fiber layers are layers of dry fibers, such that manipulation thereof by an automated machine may be effected for a reasonable cost and at a high throughput. Deposition of resin films carried by removable leaves, i.e., of which one of the faces is not adhesive, likewise facilitates manipulation by an automated installation.

A method of fabricating a solid laminate stringer on a composite panel, including unspooling one or more composite layers onto the composite panel; compacting the one or more composite layers unspooled onto the composite panel; cutting the one or more composite layers unspooled onto the composite panel; and curing the one or more composite layers unspooled onto the composite panel, wherein the one or more composite layers are unspooled continuously along a length corresponding to a length of the solid laminate stringer.

Fabrication system and associated methods of fabricating a sandwich panel. In one embodiment, a method includes holding a first skin and a second skin of the sandwich panel with a gap between opposing faces of the first skin and the second skin, and expanding a foamable material between the first skin and the second skin to form a foam core of the sandwich panel.

A method of forming a tool includes applying a first build ply to an initial ply. A second build ply is applied to the first build ply. Successive build plies are applied to build on the second build ply, the successive build plies including a medial section of build plies and a last build ply. A surface defined by the last build ply, the first build ply, and leading edges of the build plies therebetween define a substantially curvature-continuous surface profile extending between a first segment and a second segment. A shape of the substantially curvature-continuous surface profile is determined based on an adjusted ply drop ratio for at least a portion of the successive build plies.