Abstract: A method for manufacturing a stiffened panel made from composite material comprising a skin and elongated reinforcing elements wherein some of the elongated reinforcing elements cross each other. The method comprises the steps of (a) laying up a flat laminate comprising stacked plies of composite layers for forming a structure comprising the elongated reinforcing elements of the panel, (b) cutting the flat laminate along intersection lines of planes defining the webs of two crossing reinforcing elements, (c) cutting in the flat laminate the outline of the elongated reinforcing elements, (d) forming the structure comprising the elongated reinforcing elements of the panel, (e) laying up plies of composite layers for forming the skin, and (f) curing the laid-up plies of the skin and the formed structure comprising the elongated reinforcing elements.

Abstract: The present disclosure refers to a method for manufacturing carbon fiber panels stiffened with omega stringers for the construction of aircraft structures, such as fuselage sections, wing panels, etc. One tubular pressure member is provided for each omega stringer of the structure to be manufactured, wherein the tubular pressure member is configured with the shape of the omega stringer. Each tubular pressure member is enclosed between the omega stringer and part of the laminate, and autoclave pressure is applied to the interior of the tubular pressure member, so that the tubular pressure member is used to consolidate the omega stringers and/or part of the laminate from the interior of these two elements, while these two elements are being co-cured or co-bonded in an autoclave. Imperfections on those internal surfaces such as resin wrinkles of the structure are reduced.

Abstract: A system for curing polymer matrix composite parts out of autoclave in manufacturing and repairing processes including: (i) an inner vacuum bag (3) placed over the lay-up to be cured located over a working area (15, 18) with the edges of the inner vacuum bag (3) joined to the working area by first sealing elements (8); (ii) a Perforated bulkhead (4) located over the inner vacuum bag (3), (iii) an Outer vacuum bag (2) over a perforated bulkhead (4) and over the edges of the inner vacuum bag with the edges of the outer vacuum bag (2) joined to the working area (15, 18) by second sealing elements (1), (iv) a vacuum device (5, 7, 6) and (v) a heating source (9, 20).

Abstract: A system for curing polymer matrix composite parts out of autoclave in manufacturing and repairing processes including: (i) an inner vacuum bag (3) placed over the lay-up to be cured located over a working area (15, 18) with the edges of the inner vacuum bag (3) joined to the working area by first sealing elements (8); (ii) a Perforated bulkhead (4) located over the inner vacuum bag (3), (iii) an Outer vacuum bag (2) over a perforated bulkhead (4) and over the edges of the inner vacuum bag with the edges of the outer vacuum bag (2) joined to the working area (15, 18) by second sealing elements (1), (iv) a vacuum device (5, 7, 6) and (v) a heating source (9, 20).

Abstract: An apparatus for manufacturing composite stringers from a composite laminate comprising a male mold, a main caul plate, two fixed female mold halves, two secondary movable caul plates and a retention device. The main caul plate is located below the male mold, the female mold halves are located at both sides of the main caul plate and the secondary movable caul plates, each of them located respectively over each one of the two female mold halves. The male mold, the main caul plate, the two female mold halves and the two secondary caul plates are non-developable surfaces. The retention device keeps each secondary movable caul plate at a predetermined distance of its respective female mold half.

Abstract: A method for manufacturing a stiffened panel made from composite material comprising a skin and elongated reinforcing elements wherein some of the elongated reinforcing elements cross each other. The method comprises the steps of (a) laying up a flat laminate comprising stacked plies of composite layers for forming a structure comprising the elongated reinforcing elements of the panel, (b) cutting the flat laminate along intersection lines of planes defining the webs of two crossing reinforcing elements, (c) cutting in the flat laminate the outline of the elongated reinforcing elements, (d) forming the structure comprising the elongated reinforcing elements of the panel, (e) laying up plies of composite layers for forming the skin, and (f) curing the laid-up plies of the skin and the formed structure comprising the elongated reinforcing elements.

Abstract: The present disclosure refers to a method for manufacturing carbon fiber panels stiffened with omega stringers for the construction of aircraft structures, such as fuselage sections, wing panels, etc. One tubular pressure member is provided for each omega stringer of the structure to be manufactured, wherein the tubular pressure member is configured with the shape of the omega stringer. Each tubular pressure member is enclosed between the omega stringer and part of the laminate, and autoclave pressure is applied to the interior of the tubular pressure member, so that the tubular pressure member is used to consolidate the omega stringers and/or part of the laminate from the interior of these two elements, while these two elements are being co-cured or co-bonded in an autoclave. Imperfections on those internal surfaces such as resin wrinkles of the structure are reduced.

Abstract: An apparatus for manufacturing composite stringers from a composite laminate comprising a male mould, a main caul plate, two fixed female mould halves, two secondary movable caul plates and a retention device. The main caul plate is located below the male mould, the female mould halves are located at both sides of the main caul plate and the secondary movable caul plates, each of them located respectively over each one of the two female mould halves. The male mould, the main caul plate, the two female mould halves and the two secondary caul plates are non-developable surfaces. The retention device keeps each secondary movable caul plate at a predetermined distance of its respective female mould half.

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: A composite skin (11) with swept angle and dihedral for an aircraft mid-box lifting surface having a left side (5), a right side (7) and a central transition area (9) where both sides meet, that is designed and manufactured as a single part from left tip to right tip following a common ply structure for the whole part defined in relation to a single orientation rosette (25). The invention also refers to a composite aircraft mid-box lifting surface including an upper and lower panel with said composite skin (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: This invention relates to a process for manufacturing a piece of composite material made with a polymer resin and fiber reinforcement from at least two subcomponents comprising the following steps: a) providing the first subcomponent partially cured in a curing cycle in an autoclave at a maximum temperature T1, comprised between the resin gelling temperature GT and the resin curing temperature CT, applied for a predetermined time PT1 such that the exothermal component is released from the first subcomponent in a degree exceeding 50%; b) providing the second subcomponent in a fresh or cured state; c) assembling the two subcomponents and then joining them to one another in a curing cycle in an autoclave at a maximum temperature T2 comprised between 90% and 100% of the resin curing temperature CT, applied for a predetermined time PT2.

Abstract: A composite skin (11) with swept angle and dihedral for an aircraft mid-box lifting surface having a left side (5), a right side (7) and a central transition area (9) where both sides meet, that is designed and manufactured as a single part from left tip to right tip following a common ply structure for the whole part defined in relation to a single orientation rosette (25). The invention also refers to a composite aircraft mid-box lifting surface including an upper and lower panel with said composite skin (11).

Abstract: A process for manufacturing stiffened structures (11) in composite materials formed by an outer skin (13) and a plurality of stiffeners or stringers (15) the cross-section of which has a closed form delimiting an inner opening (23), comprising the following steps: a) Providing a shaping tool (31); b) Providing auxiliary male tools (37) of a material capable of varying its volume due to the effect of temperature and/or pressure, covered with membranes suitable for curing c) Providing stiffeners or stringers (15); d) Arranging the stiffeners or stringers (15) in the tool (31) and the auxiliary male tools (37) in their inner openings (23); e) Laminating the outer skin (13); f) Curing the stiffened structure (11) with high temperature and pressure; g) Removing the auxiliary male tools (37) after reducing their volume; h) Separating the cured stiffened structure (11) from the tool (31). The invention also relates to the auxiliary male tools.

Abstract: A device for manufacturing pieces of composite material by means of laying, cutting and hot forming operations comprising two same tables (11, 13) configurable with a planar upper surface (41) or with an upper surface (43) with the shape of the piece to be manufactured and means for generating a vacuum in the closed space on it; an automatic laying head (15) supported by means (19) allowing its displacement on said tables (11, 13) and a cutting head (17) associated to the same; and a tool (21), also displaceable on said tables (11, 13) including a forming membrane (25) and heating means (27). The invention also comprises a process for manufacturing pieces of composite material by means of laying, cutting and hot forming operations carried out on a same device.