Abstract: A method for forming fiber composite preforms, the preform (1) include a web (2), a flange (3) and a bent part (2.1), and the method includes: laying-up a laminate (4) onto a tooling (5), the laminate (4) comprising lateral and transverse edges (4.1, 4.2) and the tooling (5) comprising a male part (7) comprising a surface (7.1) and a lateral wall (7.2), the web (2) being configured to be located over the surface (7.1) of the male part (7) and the flange (3) being configured to be located over the lateral wall (7.2) of the male part (7); forming the preform (1) over the male part (7); clamping the lateral edges (4.2) of the laminate (4) to the tooling (5) such that the web (2) and the flange (3) of the laid-up laminate (4) are kept under tensional loads, and bending a longitudinal portion of the male part (7).

Abstract: A method for forming fiber composite preforms, the preform (1) include a web (2), a flange (3) and a bent part (2.1), and the method includes: laying-up a laminate (4) onto a tooling (5), the laminate (4) comprising lateral and transverse edges (4.1, 4.2) and the tooling (5) comprising a male part (7) comprising a surface (7.1) and a lateral wall (7.2), the web (2) being configured to be located over the surface (7.1) of the male part (7) and the flange (3) being configured to be located over the lateral wall (7.2) of the male part (7); forming the preform (1) over the male part (7); clamping the lateral edges (4.2) of the laminate (4) to the tooling (5) such that the web (2) and the flange (3) of the laid-up laminate (4) are kept under tensional loads, and bending a longitudinal portion of the male part (7).

Abstract: A tip structure for an aircraft airfoil, such as a control surface (ailerons, flaps, elevators, rudders, etc) and/or a lifting surface (wings, HTP's, VTP's) is a unitary body and includes a tip shell and a metallic material on the outer surface of the tip shell suitable to withstand a lighting strike. The tip shell has been obtained by a single-stage injection molding process using a thermoplastic composite material having fibers dispersed therein, and the metallic material has been integrally formed with the tip shell.

Abstract: A hybrid tool for curing composite structures for aircrafts, such as stringers, torsion boxes, skin panels, wing surfaces, horizontal tail or vertical stabilizers, etc. The hybrid tool comprises a metallic portion and an elastic portion arranged on a surface of the metallic portion. The elastic portion and the metallic portion are permanently joined to each other so that the metallic portion and the elastic portion together define a surface having a shape which copies at least part of a surface of a piece of composite material to be cured. The tool is capable of satisfactorily curing pieces of composite material which have a minimum thickness and/or a very aggressive change of thickness.

Abstract: A leading edge section with laminar flow control includes: a perforated outer skin, a perforated inner skin, and a plurality of suction chambers formed between the outer skin and the inner skin. The leading edge section includes a plurality of stringers span-wise arranged at the leading edge section, and integrally formed with the outer skin, such that the inner skin is joined to the stringers. A method for manufacturing a leading edge section integrating a laminar flow control system is described, wherein a perforated inner skin is joined with a perforated outer skin having a plurality of stringers integrally formed with the outer skin, such that suction chambers are defined by a part of the outer skin, a part of the inner skin and a pair of stringers.

Abstract: This disclosure relates to the manufacturing of a leading edge section with hybrid laminar flow control for an aircraft. A manufacturing method involves: providing an outer hood, a plurality of elongated modules, first and second C-shaped profiles having comprising cavities, and an inner mandrel; assembling an injection molding tool by placing each profile on each end of the inner mandrel, arranging a first extreme of each elongated module in one cavity of the first profile and a second extreme of the module in another cavity of the second profile, both cavities positioned in the same radial direction; and placing the hood on first and second profiles to close the tool. Further, the injection molding tool is closed and filled with an injection compound comprising thermoplastic and short-fiber. Finally, the compound is hardened and demolded.

Abstract: A method for manufacturing a composite part of an aircraft comprising the steps of placing a prepreg preform in a molding die between at least a molding die base part and a molding die upper part, debulking the air trapped in the prepreg, displacing the molding die upper part against the molding die base part in a direction perpendicular to the plane of the prepreg preform such that the molding die upper part presses the laminated perform until it reaches a blocking arrangement defining the position of the upper part relative to the base part for giving the prepreg preform the nominal thickness of the cured part, wherein said displacement is achieved by bolting the different parts of the molding die against each other, providing an elevated temperature to the preform such that it is cured.

Abstract: An aircraft stabilizer (15, 17) with an enlarged area of laminar flow. The lateral skins (45a, 45b; 65a, 65b) of its torsion box (43; 63) include joggled areas (53a, 53b; 73a, 73b) as attachment areas of its leading edge (41; 61) which are arranged in a rearward position with respect to the forward most spar of the stabilizer (15, 17).

Abstract: A tip structure for an aircraft airfoil, such as a control surface (ailerons, flaps, elevators, rudders, etc) and/or a lifting surface (wings, HTP's, VTP's) is a unitary body and includes a tip shell and a metallic material on the outer surface of the tip shell suitable to withstand a lighting strike. The tip shell has been obtained by a single-stage injection molding process using a thermoplastic composite material having fibers dispersed therein, and the metallic material has been integrally formed with the tip shell.

Abstract: An aircraft aerodynamic surface includes a torsion box having an upper skin, a lower skin, and a front spar, and a leading edge having an external shell and an impact resisting structure. The external shell may be shaped with an aerodynamic leading edge profile, being configured to provide Laminar Flow Control (LFC) to the leading edge. The impact resisting structure is spanwise arranged between the external shell and the front spar, and is configured for absorbing a bird strike to prevent damage in the front spar. Also, at least one of the external shell and the impact resisting structure is fitted with the upper and lower skins of the torsion box to thereby facilitate leading edge exchange.

Abstract: This disclosure relates to the manufacturing of a leading edge section with hybrid laminar flow control for an aircraft. A manufacturing method involves: providing an outer hood, a plurality of elongated modules, first and second C-shaped profiles having comprising cavities, and an inner mandrel; assembling an injection moulding tool by placing each profile on each end of the inner mandrel, arranging a first extreme of each elongated module in one cavity of the first profile and a second extreme of the module in another cavity of the second profile, both cavities positioned in the same radial direction; and placing the hood on first and second profiles to close the tool. Further, the injection moulding tool is closed and filled with an injection compound comprising thermoplastic and short-fiber. Finally, the compound is hardened and demoulded.

Abstract: A leading edge section with laminar flow control includes: a perforated outer skin, a perforated inner skin, and a plurality of suction chambers formed between the outer skin and the inner skin. The leading edge section includes a plurality of stringers span-wise arranged at the leading edge section, and integrally formed with the outer skin, such that the inner skin is joined to the stringers. A method for manufacturing a leading edge section integrating a laminar flow control system is described, wherein a perforated inner skin is joined with a perforated outer skin having a plurality of stringers integrally formed with the outer skin, such that suction chambers are defined by a part of the outer skin, a part of the inner skin and a pair of stringers.

Abstract: Highly integrated infused box made of composite material with two skins (3), several ribs (4), several stringers (5), a front spar and a rear spar, comprising a first semibox (1) and a second semibox (2) joined by connecting means, in which the first semibox (1) comprises one skin (3) and the ribs (4), and the second semibox (2) comprises one skin (3), the front spar, the rear spar and the stringers (5). A manufacturing method is also provided, which comprises forming processes for the first semibox (1), the second semibox (2) and an assembly process of the first semibox (1) with the second semibox (2).

Abstract: Method for manufacturing a base structure (8) of an aeronautical torsion box (1) for an aircraft (11) characterized in that it comprises the steps of: a.—providing at least a fresh skin (3), at least one fresh stringer (4), at least a fresh front spar (5) and fresh rear spar (6), b.—positioning the fresh skin (3), the at least one fresh stringer, the fresh spars (5, 6) in a curing tool in a configuration corresponding to that of a base structure (8), c.—subjecting the structure (8) to a single curing cycle, obtaining a cured base structure (8).

Abstract: A Method for manufacturing of a CFRP part includes laying out one or more pre-preg or composite plies, on a mould comprising a marking tool, forming an uncured laminate with a guiding mark on the uncured laminate using curing the laminate giving the part the final shape with the guiding mark and trimming and/or drilling the CFRP part taking as reference the guiding mark. There are also provided a mould for moulding and curing a CFRP part comprising a marking tool adapted to perform a guiding mark on a fresh CFRP part, and a system for manufacturing of a moulded CFRP part.

Abstract: Method for manufacturing a base structure (8) of an aeronautical torsion box (1) for an aircraft (11) characterized in that it comprises the steps of: a.—providing at least a fresh skin (3), at least one fresh stringer (4), at least a fresh front spar (5) and fresh rear spar (6), b.—positioning the fresh skin (3), the at least one fresh stringer, the fresh spars (5, 6) in a curing tool in a configuration corresponding to that of a base structure (8), c.—subjecting the structure (8) to a single curing cycle, obtaining a cured base structure (8).

Abstract: A method for manufacturing a composite part of an aircraft comprising the steps of placing a prepreg preform in a molding die between at least a molding die base part and a molding die upper part, debulking the air trapped in the prepreg, displacing the molding die upper part against the molding die base part in a direction perpendicular to the plane of the prepreg preform such that the molding die upper part presses the laminated perform until it reaches a blocking arrangement defining the position of the upper part relative to the base part for giving the prepreg preform the nominal thickness of the cured part, wherein said displacement is achieved by bolting the different parts of the molding die against each other, providing an elevated temperature to the preform such that it is cured.

Abstract: Highly integrated infused box made of composite material with two skins (3), several ribs (4), several stringers (5), a front spar and a rear spar, comprising a first semibox (1) and a second semibox (2) joined by connecting means, in which the first semibox (1) comprises one skin (3) and the ribs (4), and the second semibox (2) comprises one skin (3), the front spar, the rear spar and the stringers (5). A manufacturing method is also provided, which comprises forming processes for the first semibox (1), the second semibox (2) and an assembly process of the first semibox (1) with the second semibox (2).

Abstract: A laminate for joining parts of composite material includes a first layer of adhesive material, a second layer of preimpregnated composite material, adjacent to the first layer, and a third layer of adhesive material, adjacent to the second layer. A method for joining at least two parts of composite material with a laminate includes the steps of providing already cured composite material parts, positioning the laminate for joining between two cured parts, and applying pressure and temperature to the whole structure including the at least two parts and the laminate, in such a way that the laminate is cured and the at least two parts are joined together in a resulting structure.

Abstract: A hybrid tool for curing composite structures for aircrafts, such as stringers, torsion boxes, skin panels, wing surfaces, horizontal tail or vertical stabilizers, etc. The hybrid tool comprises a metallic portion and an elastic portion arranged on a surface of the metallic portion. The elastic portion and the metallic portion are permanently joined to each other so that the metallic portion and the elastic portion together define a surface having a shape which copies at least part of a surface of a piece of composite material to be cured. The tool is capable of satisfactorily curing pieces of composite material which have a minimum thickness and/or a very aggressive change of thickness.