A manufacturing method of a reinforced structure includes a step of preparing a stringer that has a fiber exposed surface, a step of disposing an electrically conductive protection member on the fiber exposed surface to cover at least a part of the fiber exposed surface, a step of arranging the stringer on a skin having an uncured resin component, and a step of curing the skin after the arranging. The electrically conductive protection member includes a composite material made of combination of a resin component and an electrically conductive fiber component. The step of curing includes curing the skin and the electrically conductive protection member simultaneously. The manufacturing method can prevent edge glow without increasing the number of production steps.

Methods, systems, and apparatuses are disclosed for the manufacture of composite components having incorporated reinforcing structures machined into composite material substrates, and composite components manufactured according to disclosed methods, and assemblies and larger structures comprising the composite material components.

A bonded structure is described. The bonded structure includes an outer structure including a close tolerance hole associated with a first accuracy level. The bonded structure includes an interior structure comprising an oversized hole associated with a second accuracy level that is different from the first accuracy level. The bonded structure includes an elastomeric grommet disposed in the oversized hole. The bonded structure includes one or more spacers between the outer structure and the interior structure providing a space between the outer structure and the interior structure. The bonded structure includes a fastener positioned in the close tolerance hole and in the oversized hole. The bonded structure includes a bonding media disposed in the space between the outer structure and the interior structure via one or more channels of the fastener. The bonding media, elastomeric grommet, and oversized hole collectively position the interior structure relative to the outer structure.

An ultraviolet (UV) curing system is configured to cure a composite structure. The UV curing system includes a plurality of UV light assemblies that are configured to adaptively conform to a shape of the composite structure. A UV curing method is configured to cure a composite structure. The UV curing method includes positioning a UV curing system on the composite structure, and adaptively conforming a plurality of UV light assemblies of the UV curing system to a shape of the composite structure.

One aspect of a method of manufacturing a honeycomb core includes positioning a first thermoplastic columnar cell adjacent a second thermoplastic columnar cell, modifying a thermoplastic property of the first thermoplastic columnar cell, wherein the modified thermoplastic property permits joining a circumferential surface of the first thermoplastic columnar cell to a circumferential surface of the second thermoplastic columnar cell. The method also includes joining the circumferential surface of the first thermoplastic columnar cell having the modified thermoplastic property to the circumferential surface of the second thermoplastic columnar cell resulting in the honeycomb core.

A method for producing an aerodynamic element for an aircraft, including a wall which is covered, at least partially, by a printed film including a plurality of ribs and/or grooves, the method including: providing a raw film made of a deformable material which is devoid of grooves and ribs; providing a pressure plate which includes a face provided with ribs and/or grooves complementary to the ribs and/or grooves of the printed film; placing said raw film on the wall of the element; positioning the pressure plate on the raw film, the printed side of the pressure plate facing the raw film; and a forming step during which the raw film is bonded with the wall of the component and during which the raw film is shaped by cooperation with the pressure plate to obtain the printed film including the ribs and/or the grooves.

Methods of forming a curvature into a composite stiffener and a forming tool are presented. The composite stiffener is positioned in a forming region created by a base assembly and an upper assembly. A first airflow is applied through the material of the base assembly. A second airflow is applied through the material of the upper assembly. A plurality of brackets of the upper assembly is moved relative to each other to change a curvature of the forming region along a length of the forming region to form the curvature into the composite stiffener.

Embodiments are directed to systems and methods for two or more cured composite assemblies that are bonded together to form a tubular composite structure, wherein each of the cured composite assemblies do not have a tubular shape. The tubular composite structure may form a spar for an aerodynamic component, for example. The two or more cured composite assemblies may comprise carbon or fiberglass composite materials or a combination of materials. Each of the cured composite assemblies may further comprise axial edges that are configured to be bonded to another of the cured composite assemblies, wherein the axial edges have a sloped shape. An adhesive agent may be applied on the axial edges for bonding two cured composite assemblies. Alternatively, or additionally, one or more fasteners may be used to attach the axial edges of at least two cured composite assemblies.

An aerofoil shaped body includes a plurality of longitudinal spars, an upper aerofoil cover, and a lower aerofoil cover. The spars and the covers are made of composite laminate material. One of the spars is integrally formed with one of the covers to form a spar-cover such that the composite laminate material of the spar extends continuously into the cover through a fold region created between the spar and the cover. The fold region has a fold axis extending substantially in the longitudinal direction, and the fold axis projected onto two orthogonal planes has curvature in both those planes.

A leading edge slat of a wing element of an aircraft. The aircraft defining a mark including a main fuselage axis x and a spanwise axis y. The wing procuring a lift along an axis z. The wing element having a skin forming the leading edge slat, a spar linked to the skin and a stiffening structure linked on the leading edge side to the spar and to the skin. The stiffening structure being formed from a formed sheet metal having a plurality of bosses distributed according to the length of the leading edge. The bosses extending between the spar and the inner face of the skin.