A tooling to enable variation in radius of molded continuous fiber reinforced polymer curved components which includes a first tooling surface formed on a first tooling member which may be formed of sprung material, a second tooling surface formed on a second tooling member which may be formed of sprung material. The first tooling surface and the second tooling surface are positioned to lie one over the other to form a gap therebetween for receiving a supply of continuous fiber reinforced polymer. The first tooling surface and the second tooling surface are movable towards and away from each other to modify the size of the gap therebetween and thereby modify the amount of compression applied by the first and second tooling surfaces on the continuous fiber reinforced polymer positioned therebetween. A mechanism is engaged with the first tooling member to modify a curvature of the first tooling surface.

A molding device for producing a shell element reinforced with supporting elements and composed of fiber composite material, including a mold part with a mold surface including a cavity for receiving a supporting element of the shell element, and a mold core to be arranged in the cavity and to support the supporting element on the shell element when the supporting element of the shell element is arranged in the cavity. The molding device enables prevention of the mold cores from falling out in an uncontrolled manner when the shell element is removed from the mold as the mold core includes a holding device to engage with the supporting element of the shell element to be produced located in the cavity and to hold the mold core on the supporting element when the shell element to be produced is released from the mold part.

Systems are provided for fabricating a preform for a fuselage section of an aircraft. The system includes advancing a series of arcuate mandrel sections in a process direction through an assembly line, laying up fiber reinforced material onto the arcuate mandrel sections via layup stations, uniting the series of arcuate mandrel sections into a combined mandrel; and splicing the fiber reinforced material laid-up onto the arcuate mandrel sections.

A method for welding at least two parts including a thermoplastic material and having respective surfaces to be welded, including: inserting an insert between the surfaces to be welded of the two parts; generating heat via the insert; wherein the insert moves in relation to the parts to be welded in a welding direction. Also, an installation adapted for implementation of this method.

A shell arrangement for a fuselage of an aircraft includes a first shell portion extending in a curved manner in a shell peripheral direction and a second shell portion extending in a curved manner in the shell peripheral direction, wherein, in an overlapping portion of the shell arrangement, a first end region of the first shell portion and a second end region of the second shell portion overlap each other in the shell peripheral direction and a first outer face of the first shell portion and a second inner face of the second shell portion are adhesively bonded to each other by an adhesive layer, and wherein an edge of the first shell portion which terminates the first end region with respect to the shell peripheral direction extends in the overlapping portion in an undulating manner in a shell longitudinal direction which extends transversely relative to the shell peripheral direction.

A system for fabricating a composite structure includes a ply carrier including a ply support surface configured to support at least one composite ply. The system includes a carrier transfer device configured to convey the ply carrier. The system includes a lamination system configured to selectively apply the at least one composite ply to the ply support surface of the ply carrier. The system includes a transfer system configured to remove the ply carrier from the carrier transfer device and to apply the at least one composite ply to at least a portion of a forming surface of a forming tool. The system includes a forming system configured to form the at least one composite ply over the at least a portion of the forming surface of the forming tool.

This invention relates to a method for forming a panel or skin for a hollow core door, a method of manufacturing a said door, a hollow core door, a panel, and an apparatus for forming a panel and skin for a hollow core door. The method comprises providing a substrate comprising a patterned groove therein; applying a laminating material on the substrate with an adhesive. The method then comprises sealing the substrate with the laminating material thereon in a chamber and providing at least one deformable membrane adjacent the laminating material; removing air from the chamber wherein under vacuum, the deformable membrane deforms into the groove thereby urging the laminating material into the groove; heating the laminating material to facilitate bonding thereof to the substrate; and allowing the substrate with the laminating material thereon to cool so as to form a panel with the laminating material bonded thereto.

A reinforcing component for a structure of an aircraft or spacecraft has a first component region for reinforcing at least one other element, and at least one second component region, which is permanently connected to the first component region. The reinforcing component is formed using a thermoplastic plastics material in each of the first component region and the second component region. In the first component region, the thermoplastic plastics material forms a matrix in which continuous reinforcing fibres are embedded. In the second component region, the reinforcing component comprises discontinuous reinforcing fibres or is free of reinforcing fibres.

A method for producing a fuselage portion, in particular for an aircraft or spacecraft, has the following method steps: welding a skin portion containing a thermoplastic material with a former containing a thermoplastic material in the region of a predetermined welding zone; and connecting an attachment element, configured as a crack stopper, to the skin portion and to the former in the region of the welding zone.

A method for integrating a backing-structure assembly in a structure of an aircraft or spacecraft is described. In this case, a plurality of individual elements is joined to form the backing-structure assembly. The individual elements for the backing-structure assembly and a skin portion for the structure are provided. The elements are arranged on a pre-assembly device which comprises retaining devices, which are each configured to hold one of the elements so as to be adjustable with respect to the position and/or location thereof. Some or all of the elements are connected to the skin portion. In the method, by adjusting the retaining devices for tolerance compensation, gaps between joint regions of the elements and the skin portion are eliminated or adjusted.