Abstract: A panel for an aircraft wing, having a sandwich structure in which there is arranged a reinforcing part. When the panel is attached to a bracket so as to be articulated to another element, the reinforcing part is provided in order to distribute, over its entire surface, a force transmitted by the bracket to the panel.

Abstract: A panel for an aircraft wing, having a sandwich structure in which there is arranged a reinforcing part. When the panel is attached to a bracket so as to be articulated to another element, the reinforcing part is provided in order to distribute, over its entire surface, a force transmitted by the bracket to the panel.

Abstract: The disclosed embodiments offer a solution for controlling the thickness of a finished structural composite part when precise tolerances must be complied with for the quality of the assembly and in order to ease the assembly operations of this part. In order to implement the method according to the disclosed embodiments, an additional layer of plies made out of a non-structural composite material, called calibrating plies, is applied onto the structural plies of the structural part, at least in an area where the thickness must be maintained within close tolerances and, after the composite material has been cured, the calibrating plies are machined in order to obtain the thickness required for the finished part. The thickness of the calibrating plies applied is determined according to the possible thicknesses of structural plies, taking into account the manufacturing spreads so that the machining operation for the calibrating plies does not affect the structural plies.

Abstract: An aircraft fuselage including at least two sections kept assembled at a circumferential join, each section having at least one panel. The circumferential join includes at least one inner ring, situated on one face, called the inner face, of the fuselage. The inner ring has a width li, partially overlaps each of the two panels and is fixed to the two panels by fasteners. The fuselage also includes an outer ring, situated on one face, called the outer face, of the fuselage, opposite to the inner face. The outer ring has a width le, partially overlaps the two panels, and is fixed to the two panels and to the inner ring by some or all of the fasteners fixing the inner ring to the two panels.

Abstract: A stiff panel (1) comprises a skin (2) and at least one stiffener (3) of composite material on said skin. Free surface (311) of web (31) of stiffener (3) on the stiff panel according to the invention comprises a deformable core (33) and a covering (34) which covers core (33) and joins said core with structural part (35) of web (31). The width of core (33) is equal to or greater than the width of structural part (35) of web (31) of stiffener (3), depending on the shock energy against which panel (1) must be protected.

Abstract: The disclosed embodiments offer a solution for controlling the thickness of a finished structural composite part when precise tolerances must be complied with for the quality of the assembly and in order to ease the assembly operations of this part. In order to implement the method according to the disclosed embodiments, an additional layer of plies made out of a non-structural composite material, called calibrating plies, is applied onto the structural plies of the structural part, at least in an area where the thickness must be maintained within close tolerances and, after the composite material has been cured, the calibrating plies are machined in order to obtain the thickness required for the finished part. The thickness of the calibrating plies applied is determined according to the possible thicknesses of structural plies, taking into account the manufacturing spreads so that the machining operation for the calibrating plies does not affect the structural plies.

Abstract: An aircraft fuselage including at least two sections kept assembled at a circumferential join, each section having at least one panel. The circumferential join includes at least one inner ring, situated on one face, called the inner face, of the fuselage. The inner ring has a width li, partially overlaps each of the two panels and is fixed to the two panels by fasteners. The fuselage also includes an outer ring, situated on one face, called the outer face, of the fuselage, opposite to the inner face. The outer ring has a width le, partially overlaps the two panels, and is fixed to the two panels and to the inner ring by some or all of the fasteners fixing the inner ring to the two panels.

Abstract: A stiff panel (1) comprises a skin (2) and at least one stiffener (3) of composite material on said skin. Free surface (311) of web (31) of stiffener (3) on the stiff panel according to the invention comprises a deformable core (33) and a covering (34) which covers core (33) and joins said core with structural part (35) of web (31). The width of core (33) is equal to or greater than the width of structural part (35) of web (31) of stiffener (3), depending on the shock energy against which panel (1) must be protected.