Abstract: Tooling for manufacturing multi-spar torsion boxes with different web heights, the tooling includes a mandrel module having a hollow beam geometry which comprises a first base and a second base opposite to the first base, and two walls extending between said first base and said second base, and at least one spacer module configured for coupling with the mandrel module, wherein the web height of a multi-spar torsion box is defined by the coupling between the mandrel module and at least one spacer module. A method for manufacturing multi-spar torsion boxes with different web heights.

Abstract: A method for manufacturing an aeronautical sandwich panel with a honeycomb core and results in a core sealed to prevent infused resin from entering into the honeycomb core open cells while improving its mechanical properties, especially for curved or highly curved panels. In further embodiments, the invention proposes the automation of this process.

Abstract: A manufacturing system (1) for manufacturing shaped laminates including: forming tools (2.1, 2.2) extending along a longitudinal direction X, located parallel to one another and configured to receive a laminate (4) between the forming tools; a shaping tool (3) extending along the longitudinal direction X and along a transversal direction Z, and configured to receive the laminate (4) on an external surface of the shaping tool, wherein the forming tools (2.1, 2.2) or the shaping tool (3) are movable in the longitudinal direction X and in the transversal direction Z, relative to the at least one of the shaping tool (3) and the two forming tools (2.1, 2.2).

Abstract: A method for manufacturing an aeronautical sandwich panel with a honeycomb core and results in a core sealed to prevent infused resin from entering into the honeycomb core open cells while improving its mechanical properties, especially for curved or highly curved panels. In further embodiments, the invention proposes the automation of this process.

Abstract: An asymmetric aircraft configuration having a first and a second wing, the first wing having a span larger than the second wing. A first engine is mounted on the first wing, and a second engine mounted on the rear end of the aircraft with its centerline aligned with the aircraft longitudinal axis. The rear end of the aircraft has a T tail, and the second engine is configured to ingest and consume air forming a boundary layer during the flight. A main landing gear assembly includes a first landing gear attached to the first wing, and a second landing gear attached to an area of the fuselage close to the second wing.

Abstract: Tooling for manufacturing multi-spar torsion boxes with different web heights, the tooling includes a mandrel module having a hollow beam geometry which comprises a first base and a second base opposite to the first base, and two walls extending between said first base and said second base, and at least one spacer module configured for coupling with the mandrel module, wherein the web height of a multi-spar torsion box is defined by the coupling between the mandrel module and at least one spacer module. A method for manufacturing multi-spar torsion boxes with different web heights.

Abstract: A fireproof and thermal insulator product, which comprises a material combination of an Earth Silicate material, and alumina, such that the alumina is bonded onto the Earth Silicate material or is diffused inside the Earth Silicate material.

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: 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 material part comprising a web and at least one flange. In a laying up stage, a plurality of plies of different orientations, including 0° oriented plies, are laid up to obtain a flat laminate with edges comprising an outer contour and an inner contour, the flat laminate also comprising at least one section with a radius of curvature of less than 10 m. A forming stage comprises forming a fold along a folding line on the flat laminate to create a flange forming an angle with the web, and applying a stress tension to the 0° oriented plies located at least partially between the folding line and the outer contour of the flat laminate, in the at least one section, to avoid the appearance of wrinkles.

Abstract: An integrated composite trailing edge and its method of manufacturing. The trailing edge comprises an integrated main structure having an upper cover, lower flanges, and a set of ribs extending between the upper cover and the lower flanges, with a set of sandwich type lower panels attachable to the lower flanges of the integrated main structure. The method comprises the steps of providing a set of prepreg laminated preforms over a set of tool modules having a hollow so that each laminated configures a double C-shaped laminated preform having an upper section with a recess, two primary and secondary flanges, the upper section partly forming the upper cover, the two primary flanges partly forming the ribs and the two secondary flanges forming the lower flanges.

Abstract: A method for manufacturing a composite material part comprising a web and at least one flange. In a laying up stage, a plurality of plies of different orientations, including 0° oriented plies, are laid up to obtain a flat laminate with edges comprising an outer contour and an inner contour, the flat laminate also comprising at least one section with a radius of curvature of less than 10 m. A forming stage comprises forming a fold along a folding line on the flat laminate to create a flange forming an angle with the web, and applying a stress tension to the 0° oriented plies located at least partially between the folding line and the outer contour of the flat laminate, in the at least one section, to avoid the appearance of wrinkles.

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: An integrated composite trailing edge and its method of manufacturing. The trailing edge comprises an integrated main structure having an upper cover, lower flanges, and a set of ribs extending between the upper cover and the lower flanges, with a set of sandwich type lower panels attachable to the lower flanges of the integrated main structure. The method comprises the steps of providing a set of prepreg laminated preforms over a set of tool modules having a hollow so that each laminated configures a double C-shaped laminated preform having an upper section with a recess, two primary and secondary flanges, the upper section partly forming the upper cover, the two primary flanges partly forming the ribs and the two secondary flanges forming the lower flanges.