Manufacturing structure using a composite material includes: a tapering step for forming a first tapered surface on a first stiffener in an uncured state; a bending step for bending the stiffener such that a second surface is on the inner side; an arranging step for arranging the first stiffener and a skin at prescribed positions; a vacuum suctioning step for vacuum suctioning the first stiffener and the skin; and a bonding step for curing the first stiffener to bond the first stiffener and the skin. In the tapering step, the angle formed between the first tapered surface and the second surface is an acute angle. In the vacuum suctioning step, the first tapered surface is brought into contact with the skin while maintaining contact between the first surface and the skin, and the first stiffener is deformed so that a second tapered surface is formed on the second surface.

The invention refers to a vehicle rim (1) with a rim body (2) made of fiber composite material, comprising a first subpreform (3) at a first axial end and a third subpreform (5) at the opposite second axial end, wherein a second subpreform (4) is arranged between the first subpreform (3) and the third subpreform (5), wherein the second subpreform (3) engages into a front connecting cavity (6) formed by the first subpreform (3) as well as also into a rear connecting cavity (7) being formed by the third subpreform (5). The invention also refers to a method for the production of a vehicle rim (1) with a rim body (2) made of fiber composite material, wherein an NCF material (19) comprising carbon fibers, glass fibers and/or aramid fibers is created respectively for the formation of a first subpreform (3) for a first axial end of the rim body (2) and of a third subpreform (5) for the opposite axial end of the rim body (2) with a connecting cavity (6, 7), respectively, wherein a second subpreform (4) made of a same or a similar NCF material (19) is inserted into the connecting cavity (6 and/or 7) and is fixed thereat.

In one aspect, a heating tool includes a heat source having a discharge outlet; and a manifold including an intake conduit configured to releasably connect to the discharge outlet of the heat source and a chamber coupled to the intake conduit having a ventilation path. The chamber includes a diverging portion configured to provide uniform airflow through the ventilation path to provide heating to a surface. In an embodiment, the chamber has a first portion and a second portion configured for assembly and disassembly. A method of curing an aircraft component is provided.

Systems and processes that integrate thermoplastic and shape memory alloy materials to form an adaptive composite structure capable of changing its shape. For example, the adaptive composite structure may be designed to serve as a multifunctional adaptive wing flight control surface. Other applications for such adaptive composite structures include in variable area fan nozzles, winglets, fairings, elevators, rudders, or other aircraft components having an aerodynamic surface whose shape is preferably controllable. The material systems can be integrated by means of overbraiding (interwoven) with tows of both thermoplastic and shape memory alloy materials or separate layers of each material can be consolidated (e.g., using induction heating) to make a flight control surface that does not require separate actuation.

Extendible slit tube members and methods for manufacturing extendible slit tube members are provided. In one aspect, an extendible member (10) comprises a laminated shell (2) of plural fibre reinforced layers (P1-P5) constructed and arranged to be configurable between a coiled form and an extended form. In the extended form (12) the shell is resiliently biased in the form of an elongate tube having longitudinal edges (14) defining a slit (3) along its length and wherein the shell can be opened out at the slit to assume a flattened form in which it can be wound about an axis extending transversely to its longitudinal direction to assume its coiled form (11). In the region of one or both longitudinal edges (50), the amount of reinforcing fibre is less than in the region between the edge regions (51),In another aspect, a flexible cord (40) may be attached along the edge of a shell.

A method and apparatus (14) for assembling a reinforcement web (12) for use with a wind turbine blade (10) are provided. A pre-formed flange structure (20) to be integrated with laminate layers (58, 60) to form the reinforcement web (12) is clamped into position against a mould end surface (76) using one or more locating clamps (16). The locating clamps (16) include first and second clamp blocks (80, 82) that are shaped to provide an external profile that avoids resin collection and bridging during resin injection molding, while allowing for clamping to be applied to the flange structure (20) with an easily assembled and disassembled removable engagement of the clamp blocks (80, 82). The locating clamp (16) prevents undesirable dislodgment of the flange structure (20) during the assembly process for the reinforcement web (12), and without necessitating the use of complex or expensive molding equipment or processes.

A composite blade is formed by laying up composite layers in which reinforced fibers are impregnated with resin, and has a blade root and an airfoil extending from the blade root in a longitudinal direction. The composite blade includes a first laminate of the composite layers extending along the longitudinal direction in the airfoil and extending along a first inclination direction inclined toward a direction intersecting the longitudinal direction in the blade root; a second laminate of the composite layers extending along the longitudinal direction and contacting the first laminate in the airfoil, the second laminate extending along a second inclination direction inclined toward a direction opposite to the first inclination direction in the blade root and being separated from the first laminate; and a third laminate of the composite layers provided between the first and second laminates in the blade root.

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.

Manufacturing structure using a composite material includes: a tapering step for forming a first tapered surface on a first stiffener in an uncured state; a bending step for bending the stiffener such that a second surface is on the inner side; an arranging step for arranging the first stiffener and a skin at prescribed positions; a vacuum suctioning step for vacuum suctioning the first stiffener and the skin; and a bonding step for curing the first stiffener to bond the first stiffener and the skin. In the tapering step, the angle formed between the first tapered surface and the second surface is an acute angle. In the vacuum suctioning step, the first tapered surface is brought into contact with the skin while maintaining contact between the first surface and the skin, and the first stiffener is deformed so that a second tapered surface is formed on the second surface.

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.