A method of assembly of a carbon fiber space frame structure eliminates the need for gusset plates at the joints of the space frame structure. The elimination of the gusset plates at the joints of the space frame structure reduces the weight of the space frame structure and eliminates stresses that occur at the joints due to gusset plates attached at the joints.

Aircraft stringers having carbon fiber reinforced plastic (CFRP) material reinforced flanges are disclosed. An example stringer to be coupled to a skin of an aircraft comprises a flange. The flange includes a first portion of a first stiffening segment. The flange further includes a first portion of a second stiffening segment coupled to the first portion of the first stiffening segment. The flange further includes a CFRP reinforcement segment coupled to the first portion of the first stiffening segment and to the first portion of the second stiffening segment. The CFRP reinforcement segment strengthens the first portion of the first stiffening segment and the first portion of the second stiffening segment.

A wind turbine blade comprising first and second adjacent blade sections arranged end to end along the length of the blade. Each section comprises an aerodynamic fairing and a spar. Each spar comprises a shear web extending across the fairing and a pair of spar caps, one at either end of the shear web. Each spar cap in the first section has a different cross-sectional shape and/or material from the respective spar cap in the second section and wherein the spar cap in the first section is joined to the respective spar cap in the second section via a connection piece. Each connection piece is a pre-cured component extending along the length of the blade from a first inclined end configured to connect to a first complimentary inclined end of a spar cap of the first blade section and a second inclined end.

A method for bonding composite structures which includes providing a first and second composite substrate and coupling a co-cure prepreg tape having chemically protected polymerizable functional groups onto a surface of both the first and second composite substrates. The first and second composite substrates are then cured to the co-cure prepreg tape at a first temperature to form a co-cure prepreg tape portion where the first and second composite substrates are fully cured and the co-cure prepreg tape is partially cured. The co-cure prepreg tape portion of the first composite substrate is then coupled to the co-cure prepreg tape portion of the second composite substrate and a deprotection initiator is applied to facilitate deprotection of the chemically protected polymerizable functional groups and cure the co-cure prepreg tape portion of the first and second composite substrates to form a single covalently bonded composite structure.

System includes a first object having an energy-assisted bonding (EAB) mechanism along a surface of the first object. The EAB mechanism includes a heat-activatable adhesive layer and a carbon-filled (CF) sheet material. The CF sheet material is electrically conductive for resistive heating. A control sub-system is configured to control a coupling actuator to drive an actuator body toward the first object, wherein the actuator body and the first object engage each other. The coupling actuator is configured to apply pressure to the EAB mechanism along the surface of the first object. The control sub-system is also configured to control the power source to apply a current through the CF sheet material of the EAB mechanism to provide thermal energy through resistive heating that activates the adhesive layer along the interface.

One example method for producing a component from organic sheets may comprise placing a first organic sheet and a second organic sheet next to one another to form a component preform, forming at least one overlapping joining zone by tacking the first and second organic sheets together with a connecting part in the form of a third organic sheet, transferring the component preform to a joining tool, using the joining tool to form a joined component by connecting the organic sheets through melting and compression in the overlapping joining zone, and consolidating the joined component at least in the zone of the overlapping joining zone.

A method for assembling a box structure includes elementary parts assembled along an understructure of stiffeners and skins. The understructure and skins are made of composite material with a polymer matrix. The method includes sizing the box structure for the loads to which it is subjected and for a glued assembly. A map of the loads on the structure is obtained and a first load limit is defined depending on the probability of the structure being damaged. The understructure and the skins are assembled by gluing them. An additional layer is applied that covers the assembled elementary parts to areas of the assembled box structure where the first load limit is reached.

The present invention relates to a process for producing a fiber-foam composite (FSV1), wherein a first fiber material (FM1) is applied to a first foam body (SK1) to give a first structured fiber surface (FO1) to which a second foam body (SK2) is subsequently applied to give the fiber-foam composite (FSV1).

Systems and methods of constructing a thermoplastic component include joining a thermoplastic skin to a thermoplastic core by introducing heat at an interface of each of the thermoplastic skin and the thermoplastic core to at least partially melt the thermoplastic skin to a thermoplastic core at the interfaces, applying pressure to at least one of the thermoplastic skin and the thermoplastic core to sandwich the elements, and cooling the interfaces below the melting point of each of the thermoplastic skin and the thermoplastic core to consolidate the thermoplastic skin and the thermoplastic core into a unitary thermoplastic component. An optional thermoplastic film may be disposed between the thermoplastic skin and the thermoplastic core. The thermoplastic skin may be joined to the thermoplastic core to form an intermediate thermoplastic component prior to joining the thermoplastic skin to the intermediate component.

A composite structure (1) for an aircraft, having at least one insert (2) for receiving attachment devices, each insert (2) includes a core (3) having a major dimension and containing at least one through-hole (4), and a composite strip arrangement formed by a first section (5) surrounding the core (3) and attached to said core (3) by an adhesive polymeric layer, and a second section (6) including at least one free end (6a). The first (5) and the second portion (6) of the composite strip arrangement are disposed over a first surface (1a) of the composite structure (1), such that the major dimension of the core (3) is positioned transversal to said first surface (1 a). The at least one insert (2) is co-cured with the composite structure (1).