A composite laminate for an airframe lifting surface including: at least two sides and one ramp area defined by a decreasing staggered laminate extended along a ramp direction, wherein the composite laminate includes: first plies formed by tapes arranged parallel to the ramp direction, second plies formed by tapes arranged orthogonal to the ramp direction, third plies formed by tapes arranged in a first laying up direction, being the first laying up direction different from the ramp direction and the direction orthogonal to the ramp direction, and fourth plies formed by tapes arranged in a second laying up direction, being the second laying up direction different from the ramp direction, the direction orthogonal to the ramp direction and the first laying up direction; wherein in the ramp area, the tapes forming the third and/or fourth plies are extended from one laminate side to another laminate side.

A composite structure including a multi-layer laminate having a plurality of regions each oriented out of plane relative to an adjacent region, and each extending in a longitudinal dimension of the multi-layer laminate, wherein the multi-layer laminate includes at least one ply layer having an oblique fiber orientation relative to the longitudinal dimension, and wherein the at least one ply layer includes a first section and a second section each having a side edge, the first and second sections aligned side-by-side such that a first interface defined between the respective side edges is oriented to extend in the longitudinal dimension.

A bonded joint for use in bonding composite materials is provided and includes a composite rib having electrically conductive properties and a composite structure having electrically conductive properties. An electrically conductive preform is provided that facilitates a bond between the composite rib and the composite structure. A mesh composition that bonds the composite rib to the preform and that bonds the preform to the composite structure is provided and is electrically conductive to conduct current between the composite rib and the composite structure.

A system for fusing thermoplastic composite structures includes a skin and a substructure on an inner surface of the skin. The system also includes a shaping surface of a tool, with the skin laid up on the shaping surface. The shaping surface is configured to maintain the shape of an outer mold line. The system further includes at least one insulation layer applied over a flange of the substructure and over exposed portions of the inner surface of the skin not in contact with the substructure, and a vacuum bag at least partly enclosing the skin and the substructure. Heat can be applied to the shaping surface to fuse the substructure to the skin such that the skin exceeds its melting point and at least a portion of a raised segment of the substructure does not exceed its melting point.

An aircraft fuselage includes an element forming a skin and a load-carrying structure supporting the element forming the skin. The load-carrying structure includes a plurality of elements forming a spar disposed parallel to an axial direction defined by the fuselage and a plurality of elements forming a frame disposed spaced apart along the axial direction. Each element forming a frame being arranged substantially perpendicularly to the elements forming a spar, where the element forming the skin is fastened on an external perimeter of each element forming a frame by means of fastening elements configured to keep the element forming the skin away from the external perimeter and where the element forming the skin is made of a transparent material.

A method for manufacturing a structural element for a fuselage of an aircraft. To improve the manufacture of structural elements, a method includes laying up textile material members on a mandrel to form a plurality of structural element preforms that are space apart along an extended direction of the mandrel. The structural element preforms form closed loops and are subsequently cured to obtain annular structural elements. The annular structural elements are used as basic building blocks for stiffening panel members or are directly used as structural frame elements reinforcing cut-outs in a fuselage for windows and/or doors.

Provided are methods and devices for supporting of variety of different pre-cured composite stringers after forming and prior to curing. A post-forming processing device comprises a base with a channel for receiving hat portions of different stringers. The device also comprises a support structure, at least partially extending within the channel. The support structure is configured to conform to different hat portions and to retain the shape of these hat portions. For example, the support structure is made from a flexible material, which conforms to any shape variations. In some examples, the support structure is made from a jamming material that is reshaped together with each of the pre-cured composite stringers. A post-forming processing device is used for supporting different pre-cured composite stringers while various operations are performed on these stringers, such as stringer trimming, inspection, installation of bladders and noodles, and the like.

An anti-icing honeycomb core composite manufactured by forming an electromagnetic wave absorption layer by using dielectric fiber, molding the electromagnetic wave absorption layer into a honeycomb core structure by using a molded part including a first base, a second base, and an inner block, hardening the honeycomb core structure, and removing the molded part. The molding step includes first stacking, on the first base including a plurality of grooves in which the inner blocks each having a hexagonal column shape are able to be seated, a plurality of the inner blocks and a plurality of the electromagnetic wave absorption layers as the honeycomb core structure so that the electromagnetic wave absorption layer is disposed between the plurality of inner blocks, and second stacking covering the inner blocks and the electromagnetic wave absorption layers stacked on the first base with the second base having the same shape as the first base.

A fuselage component for an aircraft, including a shell part composed of a fiber composite material and a stiffening structure which is connected to an inner surface of the shell part and has shaped parts composed of a fiber composite material, wherein the shaped parts each have a flat base section which extends along the inner surface of the shell part, and a step-shaped profile edge which extends in an encircling manner around the base section. The shaped parts are arranged in the form of parquetry with overlapping profile edges. As a result, the overlapping profile edges form stiffening profiles intersecting at an intersection. Furthermore, an aircraft which has at least one such fuselage component and a method for producing a fuselage component are described.

The present disclosure is directed to an exterior surface protective layer for protecting a composite structure from environmental conditions including: at least one curable film; such as at least two curable films, and an electrically conductive material, wherein the electrically conductive material comprises a wire-free electrically conductive material, and/or an electrically conductive polymer weave, wherein the at least one curable film includes at least one of polyurethane, polyimide, polyester or epoxy upon curing, and wherein a weight of the exterior surface protective layer ranges from about 0.02 pounds per square foot to about 0.1 pounds per square foot. An exterior surface protected composite structure and methods of forming an exterior surface protected composite structure are also provided.