Abstract: The invention provides a hybrid laminate and skin panels of hybrid laminate structure that are suitable for a supersonic civilian aircraft. The hybrid laminates include layups of layers of titanium alloy foil and composite plies, that are optimally oriented to counteract forces encountered in use, that are bonded to a central core structure, such as titanium alloy honeycomb. The reinforcing fibers of the composite plies are selected from carbon and boron, and the fibers are continuous and parallel oriented within each ply. However, some plies may be oriented at angles to other plies. Nevertheless, in a preferred embodiment of the invention, a substantial majority of, or all of, the fibers of the hybrid laminates are oriented in a common direction. The outer surfaces of the laminates include a layer of titanium foil to protect the underlying composite-containing structure from the environment, and attack by solvents, and the like.

Abstract: The invention provides a hybrid laminate and skin panels of hybrid laminate structure that are suitable for a supersonic civilian aircraft. The hybrid laminates include layups of layers of titanium alloy foil and composite plies, that are optimally oriented to counteract forces encountered in use, that are bonded to a central core structure, such as titanium alloy honeycomb. The reinforcing fibers of the composite plies are selected from carbon and boron, and the fibers are continuous and parallel oriented within each ply. However, some plies may be oriented at angles to other plies. Nevertheless, in a preferred embodiment of the invention, a substantial majority of, or all of, the fibers of the hybrid laminates are oriented in a common direction. The outer surfaces of the laminates include a layer of titanium foil to protect the underlying composite-containing structure from the environment, and attack by solvents, and the like.

Abstract: A method for forming a thermoplastic laminate having a desired laminate thickness is disclosed. Layers of composite material using precured or partially precured thermoplastic resin laminate layers are layed-up on a forming tool. A relatively constant, uniform pressure is applied to the layers to produce a preconsolidated laminate having a thickness approximately equal to the ultimate desired laminate thickness. The laminate layers are free to slide relative to one another during debulking because they are either cured or partially cured. Thus, wrinkles are avoided. The debulked laminae are then consolidated in a conventional autoclave under elevated temperature and pressure.

Abstract: A method of consolidating fiber reinforced thermoplastic poly(amide-imide) components is disclosed that is suitable for use in fabricating large parts such as aircraft wing skins and the like. The method includes enclosing the part in a vacuum bag and placing the part in the bag in an autoclave. The method includes elevating the temperatures of the part to that where volatile solvent is substantially evaporated. Temperature of the part is then further increased until resin viscosity is reduced to a minimum. A pressure differential on the order of at least 150 psig is then applied at a high rate acting upon the article, causing the resin to flow to consolidate the part, resulting in a void-free composite.