Abstract: A method and system for designing a smooth surface for a vehicle part, such as the hull of an aircraft, covering an area of the body of the vehicle part organized around an axis X. The method includes providing data indicating a profile, area, and position of a plurality of sectors perpendicular to axis X of the smooth surface, generating a first surface from the data by using a NURBS mathematical methodology, modifying N1 equidistant profiles (Py) of the surface until their area (APy) is consistent with the area determined by the area curve resulting from the data of the first step, and generating a second surface from the profiles (Py) and validating that the second surface adjusts to the area curve by using N2 equidistant profiles of the surface, where N2>N1.

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: Joint arrangement for composite-material structural members in components comprising a skin and a plurality of structural members, applicable to pairs of first and second intersecting structural members (11, 21), such as a beam and a frame in an aircraft fuselage, both members having a configuration which includes webs (31, 41), inner flanges(33, 43) and outer flanges(35, 45), in which, in the zone of intersection between said members (11, 21), the laminations of one or both flanges(33, 35) of the first member (11) include transverse flaps (55, 55; 57, 59) which are joined to one or both flanges(43, 45) of the ends of the segments (23, 25) of the second member (21) which are separated at the intersection, acting as means for transmission of loads between them.

Abstract: A part (5) of composite material having a wedge (13) between two zones (11, 15), the second zone (15) being shorter than the first zone (11), whose structure comprises from its outer surface (21) to its inner surface (23): a first section (31) formed from at least two continuous sheets (41) extending parallel to its outer surface (21), the gradient of the wedge (13) being between 20% and 50%; a wedge (33) in the shape of a triangular prism with its larger surface (27) dimensioned in such a way that it forms a wedge having a gradient of less than 20%; a second section (35) formed by a plurality of continuous sheets (45) extending parallel to the surface bounded by the said first section (31) with the said wedge (33) placed upon it. The invention also relates to a process for manufacture of the part (5).

Abstract: Joining structure for aircraft fairing (7), said fairing (7) being arranged between the horizontal stabilizer and the vertical stabilizer or fuselage section (6) of said aircraft, said structure comprising fittings (2) in the shape of angle, said fittings in turn comprising a vertical surface and a horizontal surface, the fittings (2) being joined via their vertical surface to the fairing (7) by means of removable fixing parts (12), said structure also comprising a deformable resilient element (13) between the fairing (7) and the vertical surface of the fitting (2), such that said resilient element (13) is able to compensate for the necessary deformation play and the manufacturing and assembly tolerances of the fittings (2), the fairing (7), the horizontal stabilizer and the vertical stabilizer or fuselage section (6) by means of different torques for tightening of the above parts.

Abstract: The invention relates to a device (1) for the protection against electrical discharges of a conductive fixing element (4) of an aircraft, comprising a cap (2) made of non-conductive material and a washer (3) made of non-conductive material insulating the fixing element (4) from electrical discharges, said cap being an open receptacle with a cylindrical shape closed at one of its ends and open at the other end, comprising at the lower part of its inner face a thread (15), the washer (3) being retained by the fixing element (4), said washer (3) having a thread (5) on its outer face coupling with the complementary thread (15) in the lower area of the inner face of the cap (2), relative movements between the washer (3) and the cap (2) being prevented.

Abstract: Process for manufacturing monolithic composite structures comprising providing first and second subcomponents (1, 2) of composite material, attaching an expansion compensating tooling to the second subcomponent, placing the second subcomponent along with said tooling on the first subcomponent and bonding it to the latter by means of an uncured structural adhesive, covering the assembly comprising the first and second subcomponents and the tooling with a vacuum bag, performing an autoclave cycle for curing the curable material contained in said assembly under high temperature and pressure conditions, withdrawing said assembly from the curing autoclave and removing the expansion compensating tooling to obtain the desired monolithic composite structure. The tooling for carrying out the process comprises metal L- or I-shaped beams having a rough surface adapted to be applied to the second subcomponent. The invention is useful for manufacturing aircraft parts.

Abstract: The invention relates to a fairing system for a horizontal stabiliser of an aircraft and to a process for installing said system, such that the fairing system for the sealing of the horizontal stabiliser (2) and the fuselage (1) of the aircraft comprises a primary fairing (50, 51) arranged on the horizontal stabiliser (2) of the aircraft, the mentioned primary fairing (50, 51) being solidly connected to the mentioned horizontal stabiliser (2) which is in turn trimmable with respect to the fuselage (1), the fairing system in turn comprising a secondary fairing (80, 81) such that the primary fairing (50, 51) is coupled to the horizontal stabiliser (2) through supports (30) arranged in the direction of the chord of the primary fairing (50, 51), said supports (30) being located in the outer part of said primary fairing (50, 51) with respect to the fuselage (1), the secondary fairing (80, 81) being arranged in the outer part of the supports (30), the purpose of said secondary fairing (80, 81) being to cover the me

Abstract: Arrangement for mounting a propulsive system on an aircraft airframe by an external pylon which connects the engine with the airframe, the pylon 4comprising a plurality of spars intruding into the airframe and units for attaching the spars to the internal airframe structure, the cited units being dampers, shock absorbers or active actuators with characteristics chosen so that the units are particularly suited for the efficient reduction of attachment and internal dynamic loads arising mainly but not only from hard landing events and aircraft flight manoeuvres, and for damping sustained vibrations of the propulsive system-airframe assembly, the arrangement allowing the easy removal of the pylon 4 from the rest of the airframe and the interchangeability of the pylon 4 between different aircrafts, being also the cited arrangement tolerant to damage or complete failure of at least one of the spars.

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: Fitting (30) for trimming the horizontal stabilizer (1a) of an aircraft, made of composite material, said fitting (30) comprising side walls (2a and 2b) of a torsion box as well as joining elements (3a, 3b) which join the fitting (30) to the frames of the tail fuselage (1) of the aircraft, said side walls (2a, 2b) being joined together by means of a central element (4) which comprises a first end part (4a) joined to the first side wall (2a) of the fitting (30), a second end part (4b) joined to the second side wall (2b) of the fitting (30) and a central part (4c) which joins together said end parts (4a, 4b), the fitting (30) moreover comprising end elements (5a, 5b) which are joined to said side walls (2a, 2b) on their outer face, said fitting (30) having, owing to its greater rigidity in response to side load and vertical load stresses which tend to close the side walls (2a, 2b), its greater integration and the simplicity of the load path, an optimum structural behaviour in response to the aircraft stresses.

Abstract: The invention relates to a composite structure (11) formed by a plurality of layers (13, 15, 17, 19, 21, 23) including an optical fiber (25) for structural monitoring purposes which is at least partly embedded in said structure (11), incorporating a protective cover (27) in those areas of its embedded part susceptible to needing repair, and to a process for repairing said embedded optical fiber comprising the following steps: identifying the optical fiber area in need of repair, removing material until reaching the cover (27), extracting said area, removing the protective cover (27), repairing the optical fiber (25), relocating the repaired area in the structure and returning the removed material.

Abstract: An electrical insulation system for a linear element forming part of a fluid system subjected to risks of possible external electrical discharges, such as an aircraft fuel system, using an insulating insert (12) between two parts (14, 16) of said linear element, consisting of providing said parts (14, 16) with grooved edges and in forming the insulating insert (12) between them preferably by an injection technique, such that the insulating insert (12) provides an inner duct for the passage of fluid between said parts (14, 16) and covers said grooved edges, assuring the leak-tight joint of the insulating insert (12) with said parts (14, 16).

Abstract: A mechanically fastened permanent or detachable joint between two or more parts (21, 23, 25, 41), at least one of them made of a composite material, preferably carbon fibers impregnated with thermoset resin, with continuous plies from the full-composite area along the joint area (15) and including a metal reinforcement, preferably made of titanium, in the joint area (15), the joint area (15) thickness being increased in the thickness of said reinforcement plus the eventual additional composite plies laid only on the reinforced area. The metal reinforcement may consist of one or more metal foils or sheets (31) interleaved between the composite plies or one or more metal plates (33) embedded between two of the composite plies.

Abstract: The present invention relates to a cleaning device (5) to be used in a preimpregnated carbon fiber placement machine including at least one guiding roller (47) for guiding a band (21) of tows to an application head (25), comprising two filtering elements (11, 13) arranged such that a band (21) of tows supported by two plates (15, 17) joined with adjustable clamping means (29) for said filtering elements (11, 13) and means of fixing/separating them to/from said machine, circulates between them. The invention also relates to a preimpregnated carbon fiber placement machine with at least one cleaning device (5).

Abstract: A hatch-actuating mechanism, particularly for aircraft landing gear. The mechanism includes a lateral closing panel (1) and a secondary closing panel (4) articulated to one another. A lateral pivoting fitting (5) is joined to the lateral panel (1) and pivots about a fixed point on the structure (7). A secondary pivoting fitting (8) is joined to the secondary panel (4). A lateral actuating cylinder (10) controls the movement of the lateral panel. An intermediate actuating cylinder controls the movement of the secondary panel, the intermediate actuating cylinder (13) being actuatable independently from the lateral actuating cylinder (10).

Abstract: Reinforced cover for gaps in an aerodynamic contour of a vehicle, with a first attachment section 1a attachable to a structural element 2, 5 of the vehicle 11, on a first axial plane I; a second elastic section 1c which covers a gap 3 which is located between a fixed part 5 and a moving part 6, 12 of the vehicle 11, and provided with a low-friction coating 8 on its inner surface 1e, an outer surface 1k with a fiberglass layer 7a; a main internal body 9 of an elastic material, a flexible transition section 1g between the first and the second section, 1a, 1c; the outer surface 1k of the second section 1c flushing with the outer surface 5a and extending towards the free end 1d as a continuation of the outer surface 5a of the fixed part 5.

Abstract: The present invention relates to a computer-aided process for carrying out the structural design of a stiffened panel (9) made of a composite material, optimizing a target variable, comprising a first preparation phase (21) in which a Simulation Model (25) of the stiffened panel (9) with all the relevant information for the structural analysis thereof is obtained from a General Finite Element Model (23) and the modifiable variables and the restrictions (27) are defined, and a second simulation phase (51) in which the design variables are iteratively modified for the purpose of optimizing the target variable, taking into account in each iteration the load distribution changes resulting from the previous iteration, and verifying that the restrictive conditions and the safety margins associated to the pre-established failure modes are met using for that purpose simulation expert modules (55) of families of failure modes.

Abstract: Ultrasonic head for pulse-echo multichannel inspection applied to flat and curved surfaces, which is connected to a displacement system (3), and which comprises a chassis (1) housing a tilting frame (4) joined to said chassis via a first point of rotation (1a) and a second point of rotation (1b) which define a main tilting axis (1c) around which the tilting frame (4) rotates, said tilting frame (4) presenting a set of feeler-carriers (2) in its central opening (4h) which carry an array of ultrasonic feelers (12), said array of feel-carriers (2) being able to rotate around at least one axis of rotation (5, 5a, 5b) transverse to the main tilting axis (1c) in an independent movement to that made by the tilting frame (4), the tilting frame (4) being able to be divided into two secondary frames (6, 8) which rotate independently of the rotation of the tilting frame (4).

Abstract: Fitting with torsion box produced from a plastic material reinforced with carbon fiber, comprising a torsion box (2) with side walls and individual members for attaching the box to respective frames of the tail fuselage (1), along with a stiffening element (4) which joins the side walls (2a, 2b) and which comprises individual end parts (4a, 4b) and a central part (4c); where each end part (4a, 4b) comprises first sections (4d, 4e) which are extended parallel to the side walls (2a, 2b) to both sides of the central part (4c), second sections (4f, 4g) which are extended towards the corresponding side wall (2a, 2b) and third sections (4h, 4i) each one backing onto and attached to the side wall (2a, 2b); the first sections (4e, 4d) of the end parts (4a, 4b) and the side walls (2a, 2b) of the torsion box include individual primary passage holes (5a, 5b, 5c, 5d) aligned in order to create the first fastening.