Abstract: The manufacturing method object of the invention comprises the following stages: A) stacking strips of prepreg material on a laminating tool (14), so that an angled laminated part (2) is obtained comprising a central section (2a) contained in a first plane (12), at least one side section (2b) contained in a second plane (13), and at least one bending axis (5) between the central section (2a) and the at least one side section (2b), so that, the first plane and the second plane form an angle ?; B) forming of the angled laminated part (2) comprising bending along the bending axis (5) the, at least, one side section with respect to the central section (2a), obtaining a formed part (6); C) curing, of the formed part (6).

Abstract: A method of manufacturing composite material parts is provided. The method includes tape laying by means of fiber placement on the outer surface of a tape laying tool, such that at least one laminate with a geometry different from the end geometry of the part is obtained and forming the at least one laminate obtained in tape laying by means of a forming tool, such that the laminate acquires the end geometry of the part. The method also includes curing the at least one part.

Abstract: The invention discloses a method for prevention of porosity in composite material parts (1), which is applied when the curing stage of the part (1) is being carried out on a female tool (2). In the curing stage, an adhesive (4) is placed on at least one of the areas of the radii (5) of the composite material part (1), so that the possible subsequent appearance of porosity in said areas of the composite material part (1) is prevented. Thus, the adhesive (4) acquires the fluidization temperature before the resin of the composite material and it is displaced filling the holes generated during the placement of the part (1) on the female tool (2), in the area of the radii (5) of the part (1), during the coupling between the part (1) and the female tool (2).

Abstract: The manufacturing method object of the invention comprises the following stages: A) stacking strips of prepreg material on a laminating tool (14), so that an angled laminated part (2) is obtained comprising a central section (2a) contained in a first plane (12), at least one side section (2b) contained in a second plane (13), and at least one bending axis (5) between the central section (2a) and the at least one side section (2b), so that, the first plane and the second plane form an angle ?; B) forming of the angled laminated part (2) comprising bending along the bending axis (5) the, at least, one side section with respect to the central section (2a), obtaining a formed part (6); C) curing, of the formed part (6).

Abstract: A defect recognition procedure in prepreg materials (1) draws a first transversal cross line (4b) at the beginning boundary (3b) of a defective area (2) in a prepreg material (1). A second transversal cross line (4e) at the end boundary (3e) of a defective area (2) is drawn as well. The cross lines (4b, 4e) form an angle (?) with respect to the prepreg material (1) motion direction (5). Each transversal cross line (4b, 4e) delimiting the beginning and the end of a defective area (2) has identification codes (Bi, Ei).

Abstract: A method of manufacturing composite material parts is provided. The method includes tape laying by means of fiber placement on the outer surface of a tape laying tool , such that at least one laminate with a geometry different from the end geometry of the part is obtained and forming the at least one laminate obtained in tape laying by means of a forming tool, such that the laminate acquires the end geometry of the part. The method also includes curing the at least one part.

Abstract: The invention discloses a method for prevention of porosity in composite material parts (1), which is applied when the curing stage of the part (1) is being carried out on a female tool (2). In the curing stage, an adhesive (4) is placed on at least one of the areas of the radii (5) of the composite material part (1), so that the possible subsequent appearance of porosity in said areas of the composite material part (1) is prevented. Thus, the adhesive (4) acquires the fluidization temperature before the resin of the composite material and it is displaced filling the holes generated during the placement of the part (1) on the female tool (2), in the area of the radii (5) of the part (1), during the coupling between the part (1) and the female tool (2).

Abstract: A defect recognition procedure in prepreg materials (1) draws a first transversal cross line (4b) at the beginning boundary (3b) of a defective area (2) in a prepreg material (1). A second transversal cross line (4e) at the end boundary (3e) of a defective area (2) is drawn as well. The cross lines (4b, 4e) form an angle (?) with respect to the prepreg material (1) motion direction (5). Each transversal cross line (4b, 4e) delimiting the beginning and the end of a defective area (2) has identification codes (Bi, Ei).

Abstract: A prepreg peel ply material (21) to be used for improving the surface roughness of a composite laminate (13) comprising a backing paper (23) and a base fabric (27) impregnated with a thermosetting resin, wherein the adherence between the backing paper (23) and the base fabric (27), measured by a peeling off parameter P is comprised between 2.9-3.5 min/N/cm2 and the adherence between the base fabric (27) and the composite material, measured by said parameter P is comprised between 130-160 min/N/cm2. The invention also refers to a method for determining the adaptability to an ATL machine of a prepreg peel ply material (21) including the obtention of said parameter P and to a method for manufacturing a composite component comprising an step of laying-up a peel ply base fabric (27) with an ATL machine as the last layer of said component and even as the first layer.

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 an integrated multispar torsion box structure of composite material for aircraft, comprising a lower skin (12), an upper skin (11), several spars (9), each of which comprises in turn a chord (13) and a web (14), several stringers (10) in the lower skin (12) and several stringers (10) in the upper skin (11), characterized in that the mentioned integrated torsion box structure is achieved by means of joining unitary U-shaped structural elements (15), unitary U-shaped structural elements (16) with a flap and unitary C-shaped structural elements (17) with a flap. The invention also relates to a method for manufacturing an integrated multispar torsion box structure of composite material for aircraft.

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 an integrated multispar torsion box structure of composite material for aircraft, comprising a lower skin (12), an upper skin (11), several spars (9), each of which comprises in turn a chord (13) and a web (14), several stringers (10) in the lower skin (12) and several stringers (10) in the upper skin (11), characterized in that the mentioned integrated torsion box structure is achieved by means of joining unitary U-shaped structural elements (15), unitary U-shaped structural elements (16) with a flap and unitary C-shaped structural elements (17) with a flap. The invention also relates to a method for manufacturing an integrated multispar torsion box structure of composite material for aircraft.