Abstract: A stiffened panel including at least one stiffener, known as a longitudinal stiffener, extending substantially from one extremity of the panel to another and whose trace on a surface of the panel follows a path having non-zero geodesic curvature between the two extremities of the stiffener. This configuration makes it possible to optimize both a stiffness of the panel and its resistance to the force flow to which the panel is subjected to during operation by minimizing its mass through an optimum definition of an orientation of the stiffeners relative to the force flow.

Abstract: The disclosed embodiments concern an aircraft fuselage section made of composite material with an internal skeleton having frames that form the radial structure of the fuselage and stringers that form the longitudinal structure of said fuselage, and a skin made of composite material, surrounding the internal skeleton, with said skin having an evolving thickness along a longitudinal axis X, with an approximately constant, regular interior profile. The disclosed embodiments also concern a process for manufacturing the skin of an aircraft fuselage section made of composite material, having a draping operation for strips of fibers precoated with resin around a mold, with the number of strips of fibers varying longitudinally depending on the area of the fuselage being considered so as to create a skin with an evolving thickness.

Abstract: A fuselage element including a first fuselage segment including a skin, and a junction device for connecting the first segment to a second adjacent fuselage segment is disclosed. The first segment extending along the longitudinal axis of the fuselage element includes stiffening elements extending along the axis. The end of at least one stiffening element extends beyond a free edge of the skin by a predetermined length.

Abstract: A method for manufacturing a tool equipment used in molding a panel for a fuselage of an aircraft from a composite material is provided. The method includes manufacturing a mold such that a fixed part includes a hub and an intermediary member. A plurality of removable modules which are able to be mounted on this fixed part are also manufactured and then fastened to the fixed part so as to modify an external surface upon which composite material is laid-up or draped during the molding of the panel.

Abstract: The disclosed embodiments relate to a method for manufacturing an aircraft fuselage section of composite material that comprises the following operations manufacturing a one-piece section; cutting the section longitudinally; spreading apart two edges located on the two sides of the cut; reassembling the section and making a longitudinal joint with the two edges overlapping. The disclosed embodiments also relate to a section made in accordance with this method.

Abstract: The invention relates to a fuselage element (1) comprising a first fuselage segment (2a) comprising a skin (3a), and junction means (4) for connecting the first segment (2a) to a second adjacent fuselage segment (2b), the first segment (2a) extending along the longitudinal axis (X) of the fuselage, wherein the fuselage element (1) includes stiffening elements (5a) extending along the axis (X). The end (5a1) of at least one stiffening element (5a) extends beyond a free edge (7) of the skin (3a) by a predetermined length (L).

Abstract: A tool equipment for manufacturing a panel in a composite material, in particular for an aircraft fuselage. The tool equipment (1) comprises a mould (2) including a fixed part (4) and a plurality of removable modules (6) which are able to be mounted on this fixed part (4).

Abstract: The invention relates to the field of structural panels, more specifically, but not exclusively, to the field of aircraft fuselage panels. The invention concerns a stiffened panel (10) comprising at least one stiffener (4), known as a longitudinal stiffener, extending substantially from one extremity of the panel to the other and whose trace on the surface of said panel's skin (1) follows a path having non-zero geodesic curvature between the two extremities of said stiffener. This configuration makes it possible to optimize both the panel's stiffness and its resistance to the force flow to which it is subjected in operation by minimizing its mass through an optimum definition of the orientation of the stiffeners relative to said force flow.

Abstract: The disclosed embodiments concern an aircraft fuselage section made of composite material with an internal skeleton having frames that form the radial structure of the fuselage and stringers that form the longitudinal structure of said fuselage, and a skin made of composite material, surrounding the internal skeleton, with said skin having an evolving thickness along a longitudinal axis X, with an approximately constant, regular interior profile. The disclosed embodiments also concern a process for manufacturing the skin of an aircraft fuselage section made of composite material, having a draping operation for strips of fibers precoated with resin around a mold, with the number of strips of fibers varying longitudinally depending on the area of the fuselage being considered so as to create a skin with an evolving thickness.

Abstract: The disclosed embodiments relate to a method for manufacturing an aircraft fuselage section of composite material that comprises the following operations manufacturing a one-piece section; cutting the section longitudinally; spreading apart two edges located on the two sides of the cut; reassembling the section and making a longitudinal joint with the two edges overlapping. The disclosed embodiments also relate to a section made in accordance with this method.

Abstract: An elongated part, of composite, metal matrix material, respectively comprises 35 to 45 vol. % of a matrix based on aluminum or magnesium alloy and 65 to 55 vol. % of continuous carbon fibers, arranged in sheets parallel to the length thereof. At least approximately 90% of the carbon fibers are ultra-high modulus fibers. In 25 to 60% of the sheets, said fibers are oriented at 0%±5° with respect to the longitudinal direction of the part, when the matrix is based on aluminum. In the other sheets, the fibers are then oriented between ±20 and ±40° with respect to said direction. When the matrix is based on magnesium, the ultra-high modulus fibers are oriented at 0°±5° in at least 90% of the sheets. This gives a high rigidity and high stability in the indicated direction, which favors applications in the space industry.