Abstract: A composite structural member with fiber reinforcement comprising a plurality of layers oriented relative to a longitudinal direction of said member in directions including 0°, 90° and +/??, the relative proportion of layers in one of these orientations being variable along a transverse direction of said member so as to spatially adjust the rigidity of the member according to a defined distribution of mechanical stresses along this transverse axis. The rigidity of the member is adapted locally to the stress system so as to dissipate the force flux over the entire volume of said member.

Abstract: The invention seeks to improve the integrity in respect of tolerance to damage following the spread of cracks in failure-mitigating structures of the failsafe type. To do this, the invention proposes to form a composite hybrid structure within the failure-mitigating structure. A failure-mitigating mainframe, particularly of an aircraft fuselage, equipped with a structural element made of composite according to the invention, has profiled longitudinal structural components assembled back to back by fasteners. First sides are longitudinally aligned and bonded to the fuselage skin, and the other sides are longitudinally aligned some distance away from the first sides. Fasteners couple the composite structural element to the external face of the sides of the metal components that are longitudinally aligned some distance away from the first sides.

Abstract: Aircraft fuselage section (32) subjected to impacts of external bodies, the aircraft fuselage having a curved shape with at least a vertical symmetry plane (A-A) and a central longitudinal axis and comprising a skin (35) and a plurality of frames (37) arranged perpendicularly to said longitudinal axis (33), the aircraft fuselage section also comprising at least an inner reticular structure (51, 53) mounted on a supporting structure (41, 43) comprising longitudinal beams (39) attached to the skin (35) and interconnected with said frames (37), said inner reticular structure (51, 53) being arranged for creating at least one closed cell (75) with the skin (35) for improving its resistance and its damage tolerance to said impacts. Said inner reticular structure (51, 53) can be formed by panels (61), rods (65, 65?), cables (67, 67?) or belts (69, 69?).

Abstract: An aircraft structure includes a wing spar (10), with orbital frames (24, 24?, 24?) arranged in transverse planes, right and left horizontal spars (26) connected to the wing spar (10) at the upper panel (14) of the spar respectively to the right and to the left of the spar, and for each right or left horizontal spar (26), a triangular panel (50) that includes a wall (60), a first edge arranged at the horizontal spar (26), a second side arranged at the wing spar (10), and a third side at which a first rectilinear oblique reinforcement (62) connecting the horizontal spar (26) and the wing spar (10) is provided, characterized in that the triangular panel (50) includes at least one second rectilinear oblique reinforcement (64) that connects the horizontal spar (26) and the wing spar (10).

Abstract: A method for producing a central wing box having an upper panel, a lower panel, a front spar and a rear spar, the upper edges of said spars approximately parallel to an axis Y being connected by the upper panel, and the lower edges of said spars approximately parallel to the axis Y being connected by the lower panel. The method includes producing at least one panel and at least one spar from composite material in one piece.

Abstract: Protective shield against ice impacts on aircraft, wherein the shield comprises plies of composite material (1) having microcapsules (2), each microcapsule containing a healing agent (5). When a crack (4) produced on the shield reaches at least a microcapsule (2), the healing agent is spilled in the delaminated area. Some catalyst particles (3) can be included in the material and in that case, the healing agent (5) is polymerized reacting with the catalyst particles (3). If no catalyst particles (3) are included in the material, the healing agent (5) may also actuate when manually heated. Such kinds of materials allow recovering at least partially the impact strength of the shields after impact, which is particularly important in case of ice impacts that can be repetitive during operations in icing conditions.

Abstract: The link for an aircraft comprises at least two portions respectively made of metal and of composite material together forming a common section of a middle segment of the link in a plane perpendicular to a longitudinal direction of the link.

Abstract: Aircraft fuselage section (32) subjected to impacts of external bodies, the aircraft fuselage having a curved shape with at least a vertical symmetry plane (A-A) and a central longitudinal axis and comprising a skin (35) and a plurality of frames (37) arranged perpendicularly to said longitudinal axis (33), the aircraft fuselage section also comprising at least an inner reticular structure (51, 53) mounted on a supporting structure (41, 43) comprising longitudinal beams (39) attached to the skin (35) and interconnected with said frames (37), said inner reticular structure (51, 53) being arranged for creating at least one closed cell (75) with the skin (35) for improving its resistance and its damage tolerance to said impacts. Said inner reticular structure (51, 53) can be formed by panels (61), rods (65, 65?), cables (67, 67?) or belts (69, 69?).

Abstract: The invention seeks to improve the integrity in respect of tolerance to damage following the spread of cracks in failure-mitigating structures of the failsafe type. To do this, the invention proposes to form a composite hybrid structure within the failure-mitigating structure. A failure-mitigating mainframe, particularly of an aircraft fuselage, equipped with a structural element made of composite according to the invention, has profiled longitudinal structural components assembled back to back by fasteners. First sides are longitudinally aligned and bonded to the fuselage skin, and the other sides are longitudinally aligned some distance away from the first sides. Fasteners couple the composite structural element to the external face of the sides of the metal components that are longitudinally aligned some distance away from the first sides.

Abstract: The stiffening frames (14) of an aircraft fuselage (2) remain continuous while passing through an upper panel (5) of the wing box and extend over the ribs (3) situated in that box. The overall mechanical strength of the assembly is improved, despite the notches it is necessary to make in the panel (5).

Abstract: An aircraft structure includes a wing spar (10), with orbital frames (24, 24?, 24?) arranged in transverse planes, right and left horizontal spars (26) connected to the wing spar (10) at the upper panel (14) of the spar respectively to the right and to the left of the spar, and for each right or left horizontal spar (26), a triangular panel (50) that includes a wall (60), a first edge arranged at the horizontal spar (26), a second side arranged at the wing spar (10), and a third side at which a first rectilinear oblique reinforcement (62) connecting the horizontal spar (26) and the wing spar (10) is provided, characterized in that the triangular panel (50) includes at least one second rectilinear oblique reinforcement (64) that connects the horizontal spar (26) and the wing spar (10).

Abstract: The invention relates to a composite structural member with fiber reinforcement and its application in the production of a wing box for aircraft. More specifically, it relates to a composite structural member (230) with fiber reinforcement comprising a plurality of layers oriented relative to a longitudinal direction (2) of said member in directions including 0°, 90° and +/??, the relative proportion of layers in one of these orientations being variable along a transverse direction (3) of said member so as to spatially adjust the rigidity of the member according to a defined distribution of mechanical stresses along this transverse axis (3). Thus the rigidity of the member is adapted locally to the stress system so as to dissipate the force flux over the entire volume of said member.