Abstract: Aircraft component, such as an aircraft wing, that comprises at least one panel (11) of a composite material formed by a skin (13) and at least a stiffening stringer (15) configured by a web (17) and a foot (19) bonded to said skin (13); the stringer (15) having a run-out zone inside said panel (11) subjected to a high load level; the stringer (15) having a web (17) of decreasing height in said run-out zone and a foot (19) having a first section (31) of variable width from an initial value W1 to a final value W2 and a second section (33) with a width W2 in said run-out zone; the foot (19) and the web (17) of said stringer (15) having a decreasing thickness in said run-out zone for improving the load transfer from the stringer (15) to the skin (13).

Abstract: The invention relates to a reinforcing frame (1) of the fuselage (2) of an aircraft, said fuselage (2) comprising a skin (3) and stringers (20), the mentioned frame (1) comprising a structural member (4) and a closed, internally hollow honeycomb member (5), the structural member (4) comprising a lower base (6) which is arranged on the inner face of the skin (3) of the fuselage (2), and through which the stringers (20) connecting the frames (1) to the rest of the structure of the fuselage (2) pass, and upper reinforcing members (7) on which the mentioned honeycomb members (5) are arranged, such that the honeycomb members (5) increase the inertia and the stiffness of the frame (1) without adding weight thereto, the frame (1) having high stiffness in the transverse direction, in addition to being intrinsically stable to local buckling.

Abstract: The invention relates to a process for making swaged lighting holes (1) in planar areas of parts (5) made of preimpregnated composite, by means of which the part (5) is arranged between a female tooling (11) with the final shape of the bottom side of the swaged part and a caul plate (19) with the final shape of the top side of the swaged part, except in an area (3) which must be swaged, pressure P1 is applied on the caul plate (19) in the area of the part (5) that must be maintained substantially planar, and pressure P2 is applied by means of a male tooling (17) on the area (3) of the part (5) which must be swaged at a forward speed allowing the flow of the resin during the slippage of the fibers of the material so that it may deform, adapting to the final shape of the female tooling (11).

Abstract: The invention relates to a reinforcing frame (1) of the fuselage (2) of an aircraft, said fuselage (2) comprising a skin (3) and stringers (20), the mentioned frame (1) comprising a structural member (4) and a closed, internally hollow honeycomb member (5), the structural member (4) comprising a lower base (6) which is arranged on the inner face of the skin (3) of the fuselage (2), and through which the stringers (20) connecting the frames (1) to the rest of the structure of the fuselage (2) pass, and upper reinforcing members (7) on which the mentioned honeycomb members (5) are arranged, such that the honeycomb members (5) increase the inertia and the stiffness of the frame (1) without adding weight thereto, the frame (1) having high stiffness in the transverse direction, in addition to being intrinsically stable to local buckling.

Abstract: The invention relates to a piece made of composite material with areas of different thickness manufactured from a stack of composite material fabrics comprising at least two adjacent areas (11, 15) of different thickness and a transition area (13) between both configured with a single slope (17), in which the stack is structured by: a first and a fourth section formed by at least two continuous fabrics (21, 21?; 29, 29?) extending along the three mentioned areas (11, 13, 15); a second section formed by one or more symmetrical and balanced fabric packets (23) ending in the transition area (13) placed among one or more continuous fabrics (25); a third section formed by a symmetrical and balanced fabric packet (27) extending along the three mentioned areas (11, 13, 15) placed among one or more continuous fabrics (25).

Abstract: The invention relates to a piece made of composite material with areas of different thickness manufactured from a stack of composite material fabrics comprising at least two adjacent areas (11, 15) of different thickness and a transition area (13) between both configured with a single slope (17), in which the stack is structured by: a first and a fourth section formed by at least two continuous fabrics (21, 21?; 29, 29?) extending along the three mentioned areas (11, 13, 15); a second section formed by one or more symmetrical and balanced fabric packets (23) ending in the transition area (13) placed among one or more continuous fabrics (25); a third section formed by a symmetrical and balanced fabric packet (27) extending along the three mentioned areas (11, 13, 15) placed among one or more continuous fabrics (25).

Abstract: The present invention relates to a process for manufacturing composite ring frames for aeronautical fuselages by means of the application of the RTM technology to two preforms with C- and L-shaped sections manufactured using two tools (21, 55) in the following steps: providing the material; hot-forming planar rectangular laminates (41); hot-forming laminates of right angle section (51) on one part of right angle section of the first tool (21), placing an elastic membrane (55) and applying a temperature and vacuum cycle; hot-forming the preforms into a C shape (11) and L shape (13) on a second curved tool (55) by deforming said laminates of right angle section (51) thereon, and applying a temperature and vacuum cycle. The invention also relates to said tools (21, 55).

Abstract: The present invention relates to a process for manufacturing composite ring frames for aeronautical fuselages by means of the application of the RTM technology to two preforms with C- and L-shaped sections manufactured using two tools (21, 55) in the following steps: providing the material; hot-forming planar rectangular laminates (41); hot-forming laminates of right angle section (51) on one part of right angle section of the first tool (21), placing an elastic membrane (55) and applying a temperature and vacuum cycle; hot-forming the preforms into a C shape (11) and L shape (13) on a second curved tool (55) by deforming said laminates of right angle section (51) thereon, and applying a temperature and vacuum cycle.

Abstract: The invention relates to a process for making swaged lighting holes (1) in planar areas of parts (5) made of preimpregnated composite, by means of which the part (5) is arranged between a female tooling (11) with the final shape of the bottom side of the swaged part and a caul plate (19) with the final shape of the top side of the swaged part, except in an area (3) which must be swaged, pressure P1 is applied on the caul plate (19) in the area of the part (5) that must be maintained substantially planar, and pressure P2 is applied by means of a male tooling (17) on the area (3) of the part (5) which must be swaged at a forward speed allowing the flow of the resin during the slippage of the fibers of the material so that it may deform, adapting to the final shape of the female tooling (11).

Abstract: Method for manufacturing elements made from composite material using the co-bonding technique, in which uncured elements (preformed beams (2)) are bonded onto another cured element (the skin (3)) with multiple thickness changes, using a rigid tooling made of invar used for support and positioning during the curing process. Each element is made with preimpregnated material using automated tape laying. The preform of the beams (J-section) is obtained by hot forming of flat laminates. The final curing and bonding to the precured skin (co-bonding) is performed using a direct vacuum bag in an autoclave. A flat development of the vacuum bag is performed, it is traced with a numerical control machine and it is made prior to being placed on the tool. For large surfaces with difficult access the final adjustment is performed with the tool and the part in a vertical position, due to the ergonomic difficulties involved in working on certain areas of same. The invention is applicable to the field of aeronautics.