A method for manufacturing a composite barrel structure includes fabricating a first plurality of composite panels that are assemblable into a first partial composite barrel section. The fabricating includes assembling a first layup of composite material and, concurrently, assembling at least one additional layup. The fabricating further includes heating the first layup with the at least one additional layup. A system for fabricating a plurality of panels that are assemblable into partial barrel sections includes a first workstation for fabricating a first plurality of composite panels that are assemblable into a first partial composite barrel section. The first workstation includes a first assembly station configured to concurrently assemble a first layup of composite material and at least one additional layup and a first heating station configured to heat the first layup concurrently with the at least one additional layup to yield the first plurality of composite panels.

Items may be packaged for shipping or storage using additive manufacturing techniques, also known as three dimensional (3-D) printing. Packages made by such processes may be referred to as 3-D printed packages and may include packing material printed at least partially around the item(s) and/or an outer cover printed about at least a portion of an exterior of the packing material and/or the item(s). A packaging system may include a 3-D printer and a computing device communicatively coupled to the 3-D printer. The computing device may obtain a packaging model describing a package for one or more items. A print module of the computing device may include instructions to print the package at least partially about the item(s) according to the packaging model.

A thermoplastic composite article comprising a porous core layer and an open cell skin disposed on a first surface of the core layer is described. The composite article comprises a noise reduction coefficient of at least 0.5 as tested by ASTM C423-17 and a flame spread index of less than 25 and a smoke development index of less than 150 as tested by ASTM E84 dated 2009.

The invention relates to a method for producing a shapable core (10) from a rigid panel (12), the plane of the panel being defined by the axes X and Y and the height H being oriented in the direction Z of an orthonormal reference frame, for producing composite material products, consisting in cutting the panel (12) to form core elements (16). According to the invention, the method consists in making the cuts (14, 34) along the axis Z, producing hooking means (17) on each of the core elements (16) cut in this way, so as to allow the core elements (16) to be connected to each other and to produce a hinge connection (22) with retention between the core elements (16) in the plane XY.

A vehicle outer plate panel includes: a transparent resin plate that is formed into a prescribed shape through pressure molding and is attached to a vehicle so as to constitute an outer wall of a body; and a decorative layer that is formed on at least a part of the transparent resin plate through screen printing such that the transparent resin plate is decorated therewith. A method of producing an vehicle outer plate panel is also disclosed.

An embodiment system for molding a plurality of plastic composite material panels to be assembled on a vehicle body includes a coextrusion unit configured to manufacture a multi-layered sheet in which a plurality of resin layers are laminated, a thermoforming unit configured to manufacture a forming sheet having a plurality of panel forming portions partitioned by a forming connection portion by thermoforming the multi-layered sheet, a reinforcing source spraying unit configured to spray a reinforcing source in which a fiber material and a polyurethane compound are mixed toward a preset region on the forming sheet, and a press-molding unit configured to press-mold the forming sheet applied with the reinforcing source to manufacture a composite material panel sheet in which a reinforcing layer is formed on the forming sheet.

In one aspect, there is a composite skin for a tiltrotor aircraft including a first skin having a periphery defined by a forward edge, an aft edge, and outboard ends; a second skin; and a honeycomb core disposed between the first skin and the second skin, the honeycomb core comprised of a plurality of honeycomb panels positioned along the longitudinal axis of the first skin, the plurality of honeycomb panels having an array of large cells, each cell having a width of at least 1 cm; wherein the second skin and the honeycomb core have an outer perimeter within the periphery of the first skin.

A method for forming a fiber-reinforced thermoplastic control surface may comprise: stacking plies of thermoplastic composite sheets to a first desired thickness to form a first skin; stacking plies of thermoplastic composite sheets to a second desired thickness to form a second skin; forming the first skin in a first contour; forming the second skin in a second contour; forming a stiffening member including a thermoplastic resin, the stiffening member including a shape having a plurality of peaks and troughs; assembling the stiffening member between the first skin and the second skin; and joining the stiffening member to the first skin and the second skin.

A system for manufacturing a panel includes a forming assembly having opposing press plates and at least one platen assembly arranged between the opposing press plates. The platen assembly includes first and second platens connected together via at least one elastic deformable member. Further, the forming assembly is operable in a heating mode and a cooling mode. Moreover, the first platen is maintained at a predetermined temperature range during each of the heating and cooling modes. During the heating mode, the elastic deformable member(s) is compressed such that the first and second platens contact each other. As such, one or more layers of material to be consolidated are held by the forming assembly as the forming assembly applies heat and pressure to the layer(s), thereby consolidating the panel.

Disclosed is a panel for an aircraft structure, including a laminated skin of layers of metal of which in each case two are joined together by a fibre-reinforced adhesive layer, as well as a laminated doubler of layers of metal and at least one fibre-reinforced adhesive layer in each case between two layers of metal, wherein the doubler has a smaller size of perimeter than the skin and is bonded to an outermost layer of metal of the laminated skin by a fibre-reinforced supplementary adhesive layer, with the feature that at least one part of the perimeter of the fibre-reinforced supplementary adhesive layer is staggered inwards relative to the corresponding perimeter of the doubler and that that part of the inwards-staggered perimeter of the fibre-reinforced supplementary adhesive layer is delimited by a glued edge. Also disclosed is a method for manufacturing such a panel.