A method for manufacturing a composite assembly with a continuous skin for a rear end of an aircraft by obtaining an upper part of the rear end by composite tooling. The upper part comprises a multi-spar vertical tail plane. The spars of the vertical tail plane comprise widening roots that form an upper shell of the rear end and an upper skin. Furthermore, a lower part comprises a lower shell of the rear end including semi-complete frames and stringers and a lower skin. The upper and lower parts are assembled with a joining procedure. The upper and lower skins are joined to obtain the composite assembly with the continuous skin.

An aircraft structure (11) for flow control including a perforated panel (13) having an inner surface (15) directed to a structure interior (17), an outer surface (19) in contact with an ambient flow (21), and a plurality of micro pores (23) connecting the inner and outer surfaces (15, 19). Elongate crack stopper elements (25) are attached to the inner surface (15) of the perforated panel (13). The crack stopper elements (25) are configured to inhibit crack propagation within the perforated panel (13).

A method and apparatus for forming a composite structure. A plurality of consolidated overbraided thermoplastic preforms are co-consolidated in a circumferential stackup that is circumferentially constrained. Fibers of the plurality of consolidated overbraided thermoplastic preforms are tensioned during co-consolidation.

A system for manufacturing modular aircraft includes at least a common tooling component, wherein the at least a common tooling component is configured to manufacture at least a flight component. The system includes at least a modular tooling component, wherein the at least modular tooling component wherein the at least a modular tooling component is configured to manufacture at least a fuselage component and a collar component. The system includes at least a tooling interface component, wherein the at least a tooling interface component is configured to mechanically connect the at least a common tooling component at a first end to the at least a modular tooling component at a second end.

A lightning strike protection material for an aircraft includes an electrically-conductive grid with grid-forming members and nodes where grid-forming members overlap or intersect. A plurality of the grid-forming members and/or nodes include an outward-facing surface, and at least a portion of the outward-facing surface is concave.

Systems and methods for mating component parts are described. In some embodiments, a system for mating components may include a first component, a second component, a first adhesive, and a second adhesive. The first adhesive may be configured to set in less than 5 minutes after the first component and the second component are placed in a mating position, and the second adhesive may require greater than 5 minutes to set. The first adhesive, after setting, may have sufficient strength to maintain the first component and the second component in the mating position while the second adhesive sets.

A method for fabricating a composite wing structural component for an aircraft is described. The method comprises extruding a filler material into each mold channel of a plurality of mold channels of a die to form a plurality of filler segments, removing the plurality of filler segments from the plurality of mold channels of the die, and arranging the plurality of filler segments in a space in the composite structural component, the space being defined by a radius of the composite structural component, such that the filler segments are in end-to-end contact. The method further comprises curing the plurality of filler segments in the space to fuse the plurality of filler segments.

A system and method for forming a composite part. The apparatus comprises a sleeve that molds a composite material. The sleeve has a first face and a second face. The second face has features to mold the composite material. The first face comprises a first inclined surface having an angle less than about 90 degrees and greater than about 0 degrees.

A panel for an active laminar flow control arrangement may comprise a longitudinal wall and one or more division walls extending from the longitudinal wall and extending between the first end and the second end. The panel may be coupled to a nacelle outer skin via a plurality of adhesive fasteners pre-installed onto an inner surface of the outer skin and onto stiffeners and/or stringers associated with the outer skin.

A fuselage structure splice may include a panel having an edge and a strap connected to the first panel along the panel edge. A strap surface not in contact with the panel tapers toward the panel with distance from the panel edge. A stringer is mounted on the panel and extends away from the edge of the panel and has a flange mounted to the panel. A fitting has a stringer base portion and a strap base portion. The stringer base portion is connected to the stringer flange and extends along a first line extending in a plane normal to the panel edge. The strap base portion of the fitting is mounted on the strap surface and extends along a second line in the plane. The second line is transverse to the first line and the strap base portion of the fitting has a constant thickness along the second line.