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.

According to one implementation, a composite material structure includes a corrugated stringer and a panel. The corrugated stringer has a corrugated structure including portions each having hat-shaped cross section. The corrugated stringer is made of a composite material. The panel is integrated with the corrugated stringer. The panel is made of a composite material. Further, according to one implementation, a manufacturing method of a composite material structure includes: setting a textile on a laminated body of prepregs; and producing the composite material structure by covering the laminated body with a bagging film, forming a vacuum state in a space covered with the bagging film, impregnating the textile with the resin, and thermal curing of the laminated body of the prepregs. The laminated body is a panel before curing. The textile has a structure corresponding to a corrugated stringer.

A fuselage assembly for an aircraft including a stringer extending longitudinally along the fuselage. The stringer defines a fluid channel. The fluid channel is configured to form at least part of a distribution duct of a cabin air distribution system.

A fuselage assembly for an aircraft including a stringer extending longitudinally along the fuselage. The stringer defines a fluid channel. The fluid channel is configured to form at least part of a distribution duct of a cabin air distribution system.

A joining method for connecting at least two thermoplastic workpieces is provided to permit the joining even of non-transparent carbon fiber reinforced plastics parts by means of laser welding, in which a splice is produced at the edge regions of the workpieces and the workpieces are subsequently positioned relative to one another in such a manner that the opposite splice regions bound a seam region. Connecting bodies are then inserted into the seam region and heated by means of local heat input by laser beam such that a fixed integrally bonded connection forms between the workpieces and the connecting bodies.

A joining method for connecting at least two thermoplastic workpieces is provided to permit the joining even of non-transparent carbon fiber reinforced plastics parts by means of laser welding, in which a splice is produced at the edge regions of the workpieces and the workpieces are subsequently positioned relative to one another in such a manner that the opposite splice regions bound a seam region. Connecting bodies are then inserted into the seam region and heated by means of local heat input by laser beam such that a fixed integrally bonded connection forms between the workpieces and the connecting bodies.

An aircraft fuselage assembly includes a primary structure with first and second rows of support elements extending parallel to a longitudinal axis and into an interior space of the fuselage where they have a free end. The free ends of the first and second rows are spaced from one another to from a free space below a common plane of the free ends and extending to the primary structure. A cargo floor assembly with at least one first cargo floor element is fixedly connected to the free ends of the first row in the common plane and at least one second cargo floor element fixedly connected to the free ends of the second row in the common plane. At least one central cargo floor element is arranged between the first and second lines and is releasably connected to the support elements of the first and second rows.

An aircraft fuselage assembly includes a primary structure with first and second rows of support elements extending parallel to a longitudinal axis and into an interior space of the fuselage where they have a free end. The free ends of the first and second rows are spaced from one another to from a free space below a common plane of the free ends and extending to the primary structure. A cargo floor assembly with at least one first cargo floor element is fixedly connected to the free ends of the first row in the common plane and at least one second cargo floor element fixedly connected to the free ends of the second row in the common plane. At least one central cargo floor element is arranged between the first and second lines and is releasably connected to the support elements of the first and second rows.

A skin panel configured for attachment to an aircraft structure is constructed of a composite material and has a continuous, smooth outer surface; a continuous, smooth inner surface that is configured for attachment to an aircraft structure; a peripheral edge that extends entirely around the skin panel and has a continuous, smooth surface and a plurality of cavities inside the material of the skin panel. The cavities reduce the weight of the skin panel without significantly detracting from the compression strength and tensile strength of the skin panel.

A skin panel configured for attachment to an aircraft structure is constructed of a composite material and has a continuous, smooth outer surface; a continuous, smooth inner surface that is configured for attachment to an aircraft structure; a peripheral edge that extends entirely around the skin panel and has a continuous, smooth surface and a plurality of cavities inside the material of the skin panel. The cavities reduce the weight of the skin panel without significantly detracting from the compression strength and tensile strength of the skin panel.