PROCESS FOR THE MANUFACTURE OF PARTS MADE OF COMPOSITE MATERIALS WITH TWO CURING CYCLES

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.

PROCESS AND TOOLS FOR MANUFACTURING COMPOSITE RING FRAMES

The present invention relates to a process for manufacturing composite ring frames for aeronautical fuselages by means of the application of the RTM technology to two preforms with C- and L-shaped sections manufactured using two tools (21, 55) in the following steps: providing the material; hot-forming planar rectangular laminates (41); hot-forming laminates of right angle section (51) on one part of right angle section of the first tool (21), placing an elastic membrane (55) and applying a temperature and vacuum cycle; hot-forming the preforms into a C shape (11) and L shape (13) on a second curved tool (55) by deforming said laminates of right angle section (51) thereon, and applying a temperature and vacuum cycle. The invention also relates to said tools (21, 55) ...

Method and system for designing smooth surfaces for vehicle parts

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.

Preventive defect detection and control process in composite material parts

A preventive defect detection and control process for a mass produced composite material parts automatically obtains representative parameters of backwall and intermediate echo signal results in each cell of a mesh predefined on the parts, stores the parameters in a database and statistically analyzes the stored parameters for detecting isolated and significant alterations in the parts manufacturing process generating porosity defects, detecting slow and permanent alterations in the parts manufacturing process generating porosity defects or detecting areas with a negligible defect generation probability in order to identify areas susceptible of inspection by sampling.

Current insulation device for fuel systems

An electrical insulating device for a linear component forming part of a fuel system subjected to risks of possible external electrical discharges formed by a part ( 13 ) made of insulating material attached to two metal flanges ( 15, 17 ) that allow inserting the device ( 11 ) between two parts of said linear component, said part ( 13 ) being formed by a cylindrical body ( 21 ) with a castellated crown ( 23 ) on its outer side, formed by longitudinal ( 25 ) and transverse ( 27 ) segments, and said metal flanges ( 15, 17 ) having transverse projections ( 31, 33 ) facing the transverse segments ( 27 ) of the castellated crown ( 23 ) for facilitating their attachment by suitable means ( 39 ).

Aircraft tail assembly

An aircraft tail assembly planform comprising curvilinear leading edges (21) and trailing edges (22), with an aircraft tail assembly configuration in which the hinge line (13) is rectilinear and has a non-constant percentage with respect to the chord (50) in each section (51), being the front (11) and rear (12) spars rectilinear with a non-constant percentage with respect to the chord (50) in each section (51), or being these spars (11, 12) curvilinear with a constant or non-constant percentage with respect to the chord (50) in each section (51). Moreover, the invention describes an aircraft tail assembly planform comprising rectilinear leading edges (21) and trailing edges (22), in which the hinge line (13) has a non-constant percentage with respect to the chord (50) in each section (51), being the front (11) and rear (12) spars rectilinear, with a constant or non-constant percentage with respect to the chord (50) in each section (51), or being these spars (11, 12) curvilinear, with a ...

Rotation fitting for the empennage of an aircraft

Rotation fitting secured to a frame (6) of the fuselage of an aircraft and linked to a receiving element secured to the empennage (5) which comprises a central fitting (1) and two side fittings (2,3), the central fitting (1) presenting two lugs (1d,1k) with an opening (1e,1j) each and four side wings (1g,1f,1i,1h), and each one of the side fittings (2,3) presenting a lug (2d,3d) with an opening (2e,3e) each and two side wings (2g,2f,3g,3f), the central fitting (1) being secured to a surface of the frame (6) and the side fittings (2,3) in an opposite surface of the frame (6), such that the frame (6) is arranged between the central fitting (1) and the two side fittings (2,3), such that the lugs (1d,1k) of the central fitting (1) make contact with the lugs (2d,3d) of the side fittings, with all the openings (1e,1j,2e,3e) being aligned in order to house a pin (4) associated with the receiving element secured to the empennage (5).

RIB STRUCTURE FOR TORSION BOXES OF A WING OR HORIZONTAL STABILISER OF AN AIRCRAFT

Rib structure for a torsion box of an aircraft wing or a horizontal stabiliser. The structure comprises a center element which extends between the front spar and the rear spar of the torsion box, a series of vertical stiffening elements arranged between an upper skin and a lower skin of the torsion box. The vertical stiffening element consists of a first lateral wing which extends in the direction of said front spar and a second lateral wing which extends in the direction of said rear spar. These lateral extensions emerge from respective opposite sides of the vertical body and are attached side to side to the center element of the rib.

Verfahren zur Herstellung von vorgehärteten Teilen aus Verbundwerkstoff mit in frischem Zustand angewandten Versteifern

Vorrichtung zur Verbesserung der Notausfahreigenschaften eines Fahrwerks

COMPOSITE STRUCTURE WITH EMBEDDED OPTICAL FIBER AND A PROCESS FOR ITS REPAIR

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.

QUALITY CONTROL PROCESS FOR A STRUCTURAL BONDED JOINT

The invention relates to a quality control process for an adhesive joint of two sub-components of a structure comprising the following steps: a) Providing the structure of composite material to be controlled; b) Providing at least one premanufactured testing device (1) representative of one of the sub-components of the structure; c) Bonding said at least one premanufactured testing device (1) to the other sub-component (7) of the structure in conditions similar to those of the real adhesive joint of said sub-components; d) Carrying out at least one mechanical test on said at least one premanufactured testing device (1) which allows assessing the quality of said adhesive joint.

COMPOSITE MATERIAL STRUCTURE FOR AIRCRAFT FUSELAGE AND PROCESS FOR MANUFACTURING IT

The present invention provides a closed composite material structure for aircraft fuselage shaped on a male jig from which it can be separated in a certain direction, said structure comprising a single outer panel and a plurality of inner longitudinal stiffeners integrated in said panel, such that the expansion coefficient of the male jig is greater than the expansion coefficient of the composite material of the structure, thus being able to remove the already manufactured structure, formed by the panel and the integrated inner stiffeners, in a single operation. The present invention further provides a process for manufacturing such a closed structure.

METHOD FOR PRODUCING STRUCTURES FROM COMPOSITE MATERIALS, INCLUDING EMBEDDED PRECURED TOOLS

The present invention relates to a process for manufacturing composite structures formed by two subcomponents of the same material, characterized in that it comprises the following steps: providing a first subcomponent (13), particularly a skin; providing a tool (15) made of a composite and precured for the manufacture of the second subcomponent (17), particularly a stiffener; positioning said tool (15) on said first subcomponent (13); applying preimpregnated composite (21) on said tool (15) so as to form the second subcomponent (17); consolidating the composite structure by means of a process of curing the assembly resulting from the previous steps under suitable pressure and/or temperature conditions.

METHOD FOR PRODUCING PANELS OF COMPOSITE MATERIALS WITH U-SHAPED STIFFENING ELEMENTS

The present invention relates to a process for manufacturing panels (9) for aeronautical structures with U-shaped stiffening members (13) and I-shaped stiffening members (17) between their webs comprising the following steps: providing laminates for shaping the skin (11) on the curing tool (31); providing planar laminates (23) for shaping the stiffening members (13, 17); shaping the U-shaped stiffening members (13) on individual shaping tools (33) and placing the I-shaped stiffening elements (17) in said tools; grouping said individual shaping tools (33) together on an assembly tool (41); placing the group of stiffening members (13, 17) on the skin (11); placing a vacuum bag (55) on the assembly with the aid of profiles (59); consolidating said assembly by means of a curing process under suitable pressure and temperature conditions using external tools (60) to assure verticality of the webs of the stiffening members (13, 17).

CLEANING DEVICE FOR A PREIMPREGNATED CARBON FIBER PLACEMENT MACHINE

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).

COMPOSITE STRUCTURE WITH OPTICAL FIBER EMBEDDED IN ONE OF ITS SURFACE LAYERS AND A PROCESS FOR ITS CONNECTION AND REPAIR

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 a surface layer (13) of said structure (11), and insulation means of the optical fiber (25) areas susceptible to repair with respect to the surface layer (13) in which they are integrated, particularly a protective cover (27) and top and bottom separating films (31, 33), and a process of repairing said areas comprising the following steps: removing the protective cover (27) and the top separating film (31), extracting the area, repairing the optical fiber, relocating the area and providing a new protective cover (27).

Sealing device for a flap track in a belly fairing of an aircraft for a flap-actuating shaft

Sealing device for the flap track for the flap-actuating shaft, in the belly fairing of an aircraft, with a top hollow body 1 and a bottom hollow body 7 which are flexibly compressible, each comprising at least one interior chamber 2, 8 with open sides 2 a, 2 b, a wall 3, 9 with respective external contact surfaces 3 a which are in flexible contact with each other, the outer surfaces 4 a, 10 a of the front walls 4, 10 of the hollow bodies 1, 7 jointly forming an outer sealing surface 11 which covers the flap track 19 , said hollow bodies 1,7 having jointly a vertical compression flexibility such that the contact surfaces 3 a, 9 a retract around the actuating shaft 20 , the interior chambers 2, 8 extending between the front walls 4, 10 and rear walls 5, 11 of the hollow bodies 2, 7.

Klappenbetätigungsmechanismus, insbesondere für Fahrwerksklappen

TOOL AND PROCESS FOR MANUFACTURING PIECES OF COMPOSITE MATERIALS OUTSIDE AN AUTOCLAVE

This invention relates to a tool (9) for manufacturing pieces of composite material outside an autoclave comprising a stacking table (11); a movable head (15) on said stacking table (11) provided with: automatic means for placing tapes (19) of composite material; compacting means for compacting the composite material and microwave emission means (25) to cure the composite material. The invention also relates to a process for manufacturing pieces of composite material outside an autoclave, comprising the following steps: a) Placing composite material in the form of tapes on a tool with the shape of the piece to be manufactured, compacting it and partially curing it after its placement until completing a layer of the piece; b) Repeating step a) until completing the stacking of the piece and c) Curing the last layer of the piece until the required degree of curing.

ELECTRIC INSULATION DEVICE FOR USE IN FUEL LINES

An electrical insulating device for a linear component forming part of a fuel system subjected to risks of possible external electrical discharges formed by a part (13) made of insulating material attached to two metal flanges (15, 17) that allow inserting the device (11) between two parts of said linear component, said part (13) being formed by a cylindrical body (21) with a castellated crown (23) on its outer side, formed by longitudinal (25) and transverse (27) segments, and said metal flanges (15, 17) having transverse projections (31, 33) facing the transverse segments (27) of the castellated crown (23) for facilitating their attachment by suitable means (39).

Metallic local reinforcement for heavy loaded joints of composite components

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.

VERFAHREN ZUR HERSTELLUNG VON GEWICHTSVERRINGERUNGSÖFFNUNGEN IN PLANAREN FLÄCHEN VORIMPRÄGNIERTER VERBUNDTEILE DURCH ZIEHPRESSEN

System zur Kompensation aerodynamischer Messungen in einem Windkanal

Lightning strike protection system for aircraft fuel tanks made of low electrical conductivity composite material

Lightning strike protection system for aircraft fuel tanks made of low electrical conductivity composite material. The system comprises an electrical conductive thin wire mesh ( 1 ) that covers the whole external surface of the tank outer skin (I) made of composite and a thick wire metallic mesh ( 2 ) overlapping the mesh ( 1 ) at a minimum distance of 50 mm to both sides of a row of fasteners (III) joining the said outer skin (I) to one internal part (II) of either composite or metallic material. Both metallic meshes ( 1, 2 ) maintain electrical contact by their installation/assembly and by means of metallic countersunk head washers ( 3 ) connected to bonding points and set to the gap existing between the fastener (III) and the outer skin (I). The corresponding fabrication process is also described.

Reinforced cover for gaps in an aerodynamic contour

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.

Klappenbetätigungsmechanismus, insbesondere für Fahrwerksklappen

Prozess und Werkzeuge zum Reduzieren von thermisch induzierten Restspannungen und Formverzerrungen bei monolithischen Verbundstrukturen

Process for optimizing the structural design of a stiffened composite material panel

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.

Methods and systems for generating suitable meshes for hybrid RANS/LES modelling

A method of generating a mesh of an object (11) that moves through a fluid medium which is used in the design or analysis of the object (11) in connection with a hybrid method combining RANS and LES comprising the following steps: a) creating an inner C topology (21) around the object (11) for a boundary layer description; b) creating an outer C topology (23) covering a space region including the separation region; c) locating the separation region and adapting to it the inner and outer C's topologies (21, 23) so that the outer C topology (23) is adapted to the size of the separated region and the inner C topology (21) is adapted to the size of the boundary layer; d) refining the mesh in the separation region according to the specifications of the RANS/LES method. The invention also provides a system for carrying out the method.

Auxiliary power unit intake duct with aero-acoustic guide vanes

An auxiliary power unit intake duct for an aircraft, comprising an internal straight portion 1a, an internal first bend portion 1b and an internal second bend portion 1c, extending between an air inlet 2 and a plenum entry 3 of the auxiliary power unit, a plurality of acoustic guides vanes 4,5 located in the straight portion 1a of the intake duct 1, extending horizontally between opposite side portions of the straight portion 1a; and a plurality of curved aerodynamic guide vanes 6,7 located in the second bend portion 1c proximate to said plenum inlet 3, extending vertically between a top portion and a bottom portion of the second bend portion 1c, each of the curved aerodynamic guide vanes 6,7 being concentric with said bend curvature of the second bend portion 1c.

Vorrichtung zum Übertragen des Druckes zwischen einem abnehmbaren und einem festen Teil eines Modells in einem aerodynamischen Windkanal

Vorrichtung zur Verhinderung von Ruderflattern bei aerodynamischen Steuerflächen

Ein Verfahren für die abmontierbare und/oder feste Zugkompressionsverbindung zur Anwendung bei Verbundmaterialien

Ultrasonic head for pulse-echo multichannel inspection

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).

Sicherheitsbefestigung eines Stellantriebs

Ein Verfahren für die abmontierbare und/oder feste Zugkompressionsverbindung zur Anwendung bei Verbundmaterialien

Verfahren zur Herstellung von vorgehärteten Teilen aus Verbundwerkstoff mit in frischem Zustand angewandten Versteifern

Verfahren zur Überwachung von Strukturschäden und ihr Fortschritt in monolitischen Verbundstrukturen mittels Faser-Bragg-Gitter-Sensoren

Prozess und Werkzeuge zum reduzieren von thermisch induzierten Restspannungen und Formverzerrungen bei monolithischen Verbundstrukturen

Non-destructive method for detecting zones with non conductive materials in a composite part

Non-destructive method for detecting zones with non-conductive materials, such as materials that include glass fibers, in a part made of a conductive composite, such as a composite whose reinforcing fibers are carbon fibers, provided with an organic coating, that includes the following stages: a) providing a device for applying an electric potential on the surface of said part; b) determining the dielectric breakdown potential Pr corresponding to the thickness E of the coating; c) applying said dielectric breakdown potential Pr with said device to the part for the purpose of identifying those zones that have non-conductive materials when dielectric breakdown does not occur in them.

Schutzsystem gegen Blitzeinschlag für Flugzeugbrennstofftanks aus Compositematerial mit geringer elektrischer Leitfähigkeit

Schutzsystem gegen Blitzeinschlag für Flugzeugbrennstofftanks aus Compositematerial mit geringer elektrischer Leitfähigkeit

Computer-aided method for a cost-optimized calculation of aerodynamic forces on an aircraft

A computer-aided method suitable for assisting in the design of an aircraft by providing the values of dimensional variables, dependant of a predefined set of parameters, for the complete aircraft or an aircraft component, comprising the following steps: a) Defining a parametric space grid; b) Obtaining a suitable Reduced Order Model (ROM) model, particularly a Proper Orthogonal Decomposition (POD) model, for calculating said variables for whatever point over the parametric space through an iterative process. Computer Fluid Dynamics (CFD) is used to calculate said variables for an appropriately selected set of points over the parametric space, which are used to approximate, via ROM and ad hoc interpolation, the variables in any other point over the parametric space. The method minimizes the required number of CFD calculations (to minimize the computational cost, which dramatically depends on this number) for a given level of error.

Vorrichtung zum Dürchführen von Einschlägen mit niedriger Geschwindigkeit auf mechanische Strukturen

Halterungssystem für Hilfsaggregaten

Tubular components for aeronautical fuselages and processes and jigs for its manufacturing

The present invention relates to a tubular-shaped component (19) for an aeronautical fuselage made of composite materials comprising an outer skin (1) without joints having a cylindrical or cylindrical-conical shape, a plurality of stiffeners or stringers (2) longitudinally arranged in said outer skin (1) and a plurality of anchoring elements (3) for anchoring other elements to subsequently be incorporated transversally arranged between said stiffeners or stringers (2). The invention also comprises a jig (18) comprising a support body (12) and a plurality of panels (9) that can shift between an extended position (10) and a retracted position (10') in both directions and a process for manufacturing said component (19) in the last step of which once the tubular component (19) is formed and cured, it is separated from the jig (18) by means of a longitudinal movement after the shifting of those elements of the jig (18) preventing it, and particularly the panels (9) that can be shifted.

Composite structure with optical fiber embedded in one of its surface layers and a process for its connection and repair

The invention relates to a composite structure (11) formed by a plurality of layers (13, 15, 15, 19, 21, 23) including an optical fiber (25) for structural monitoring purposes which is at least partly embedded in a surface layer (13) of said structure (11), and insulation means of the optical fiber (25) areas susceptible to repair with respect to the surface layer (13) in which they are integrated, particularly a protective cover (27) and top and bottom separating films (31, 33), and a process of repairing said areas comprising the following steps: removing the protective cover (27) and the top separating film (31), extracting the area, repairing the optical fiber, relocating the area and providing a new protective cover (27).

Control surface of aircraft

Aircraft control surface (1), in particular for an aircraft lifting surface (2), that comprises a primary control surface (6) that comprises a hinge axis (10), and a secondary control surface (7) that comprises a hinge axis (11), with the secondary control surface (7) rotating by means of its hinge axis (11) relative to the primary control surface (6), said secondary control surface (7) only partially occupying the span of the primary control surface (6), the length of the secondary control surface (7) along its hinge axis (11) being significantly less than the length of the primary control surface (6) along its hinge axis (10), and moreover the width or chord of said secondary control surface (7) along the direction of its hinge axis (11) narrows significantly towards the tip of the lifting surface (2) according to a law of narrowing designed expressly for adapting the distribution of torsional stiffness along the span of the lifting surface (2) to the distribution of aerodynamic load ...

Method for simulating the behavior of a bonded joint of two parts

A method for simulating the behavior of a bonded joint of two composite material parts with an adhesive material layer. The method includes the steps of a) providing a Finite Element Model with all the relevant information for analyzing the structural joint; b) providing a calculation model for calculating the deformations of the adhesive material layer when it is subjected to tensile/peel and shear stresses, including a plastic behavior of the adhesive of a linear type under tension/peeling and of a non-linear type under shearing; and c) simulating the behavior of the bonded joint by applying the calculation model to obtain the failure indices of each failure mode. A system which is useful as an aid in the design of the structural joint using a computer-implemented Finite Element Model of the structural joint.

VERFAHREN UND WERKZEUGE ZUR AUSKLEIDUNG VON FORMWERKZEUGEN FÜR GEKRÜMMTE TEILE

Aircraft control device

The invention relates to a control surface (3) for an aerodynamic lifting surface (2) of aircraft comprising a main closing rib (9) located at one end of the control surface (3) to which a main torsion bar (8) is joined, the mentioned torsion bar (8) being joined at its other end to a lever system (14) on which at least one actuator (15) acts, such that it is possible to act on the rotation of the mentioned control surface (3) during the flight of the aircraft. The invention also relates to an actuation method for actuating such a control surface (3).

Integrated multispar torsion box of composite material

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.

Method for determining the thickness of a coating on a composite material

The invention relates to a method for determining the thickness of a coating on a composite material by applying induced currents wherein the composite material together with its coating is placed on a conducting material, in which circulating electric currents induced by an alternating magnetic field are created, an indirect measurement being subsequently carried out based on the measurement of the lift-off effect that the coating, together with the composite material, offers in induced currents against the mentioned conducting material.

Integrated protective device against electrical discharges in fixing elements

The invention relates to an integrated device (1) for the protection against electrical discharges of conductive fixing elements (2) between one another or of a fixing element and another conductive element of aircraft, comprising a cap (3) and an o-ring gasket (4) made of non-conductive insulating material, the mentioned integrated device (1) comprising a modified fixing element (2) working together with the cap (3) and the o-ring gasket (4), the fixing element (2) surrounding the cap (3) and the o-ring gasket (4) in a non-conductive manner, thus preventing the possibility of electrical discharges such that in the device (1) the number of necessary components is minimized, the impact in weight for the aircraft being minimum, the modified fixing element (2) further comprising a thread (5) on the outer face of the area that has to be protected.

Composite structure with embedded optical fiber and a process for its repair

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.

Hatch-actuating mechanism, particularly for an aircraft landing gear

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).

Optimized configuration of engines for aircraft

The invention relates to a configuration of engines (3) for aircraft located in the rear part of the fuselage (2) of said aircraft, the engines (3) being attached in a fixed manner by pylons (5) to the structure of the aircraft, said structure comprising a torsion box (14) which traverses the fuselage (2) and is used to attach the pylons (5), the fuselage (2) comprising an opening (4) allowing the passage of the suspension pylons (5) for the engines (3), said configuration further comprising a pivoting area (8), an actuator (7) and a fitting (6) through which the actuator (7) is attached to the suspension pylons (5) and to the torsion box (14) of the aircraft, such that the assembly formed by the actuator (7) and the fitting (6) allow balancing the pylon (5) and engine (3) assembly of the aircraft through the pivoting area (8), thus achieving controllable and optimal thrust vectoring of the aircraft for each flight phase.

Cleaning device for a preimpregnated carbon fiber placement machine

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).

Arrangement for mounting an engine on the airframe of an aircraft

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.

Process of manufacturing composite structures with embedded precured tools

The present invention relates to a process for manufacturing composite structures formed by two subcomponents of the same material, characterized in that it comprises the following steps: providing a first subcomponent (13), particularly a skin; providing a tool (15) made of a composite and precured for the manufacture of the second subcomponent (17), particularly a stiffener; positioning said tool (15) on said first subcomponent (13); applying preimpregnated composite (21) on said tool (15) so as to form the second subcomponent (17); consolidating the composite structure by means of a process of curing the assembly resulting from the previous steps under suitable pressure and/or temperature conditions.

Process and tooling for reducing thermally induced residual stresses and shape distortions in monolithic composite structures

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.

Current insulation system for fluid systems

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).

Process and tooling for lining molds of curved pieces

The invention relates to a process and tooling for the lining of molds for curved pieces, the process comprising the application of a lining ( 15 ) of width A 2 on the mold ( 3 ) in several successive passes, exerting pressure thereon so that it adapts to the curvature of the piece, and the number of passes being the minimum required so that the width A 2 of the tape ( 15 ) allows achieving such adaptation, and the tooling comprising a drum ( 25 ) for applying the lining ( 15 ), a means of attaching its end to the mold ( 3 ) at the beginning of each pass, a conical roller ( 25 ), to exert pressure on the lining ( 15 ) uniformly along its entire width, and means ( 26, 29 ) for its movement through side guides ( 27, 28 ) arranged on the mold ( 3 ) parallel to the curved lines ( 5, 7 ) demarcating the piece.

Process for making swaged lighting holes in planar areas of preimpregnated composite parts

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).

Method for manufacturing elements of composite materials by the co-bonding technique

Method for manufacturing elements made from composite material using the co-bonding technique, in which uncured elements (preformed beams (2)) are bonded onto another cured element (the skin (3)) with multiple thickness changes, using a rigid tooling made of invar used for support and positioning during the curing process. Each element is made with preimpregnated material using automated tape laying. The preform of the beams (J-section) is obtained by hot forming of flat laminates. The final curing and bonding to the precured skin (co-bonding) is performed using a direct vacuum bag in an autoclave. A flat development of the vacuum bag is performed, it is traced with a numerical control machine and it is made prior to being placed on the tool. For large surfaces with difficult access the final adjustment is performed with the tool and the part in a vertical position, due to the ergonomic difficulties involved in working on certain areas of same. The invention is applicable to the field of ...

Optimized aircraft manhole

An aircraft manhole including a skin panel of an aircraft structure, wherein an outer door and an inner door of the aircraft manhole are positioned on the skin panel. The aircraft manhole further includes stringers installed on the skin panel positioned on and attached to the skin, and fixing edges partially projecting through the manhole on the side for positioning the outer door and the inner door. The aircraft manhole also includes reinforcements arranged between the outer door and the inner door of the manhole that are attached to the skin, the reinforcements forming the fixing edges on one of its sides, and a cavity arranged in the skin to allow installing the outer door, thus maintaining the aerodynamic surface of the assembly as well as its tightness.

Protective device against electrical discharges in fixing elements

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.

Method to monitor structural damage occurrence and progression in monolithic composite structures using fibre Bragg grating sensors

A method of monitoring structural damage in a composite structure manufactured by co-curing, co-bonding or secondary bonding of several sub-components ( 7, 8 ), using fibre optic Bragg grating sensors attached or embedded to or between (in the bonding line) said sub-components, comprising a first step of measuring the wavelength spectra ( 31, 33, 40 ) of said Bragg grating sensors at the end of the manufacturing of the part, and in a known load condition, considered as reference and a second step of identifying the occurrence of a failure of the structure and the progress of said failure detecting the release of the residual stresses/strains stored during the curing process by measuring changes ( 31 to 32, 33 to 34, 40 to 41 ) with respect to said reference wavelength spectra.

Process for manufacturing composite material structures with collapsible tooling

A process for manufacturing stiffened structures (11) in composite materials formed by an outer coating (13) and a plurality of stiffeners (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) with an inert material inside them coated with membranes suitable for curing c) Providing stiffeners (15) in fresh or cured condition; d) Arranging the stiffeners (15) in the tool (31) and the auxiliary male tools (37) in their inner openings (23); e) Laminating the outer coating (13); f) Curing the stiffened structure (11) with high temperature and pressure; g) Removing the auxiliary male tools (37) after reducing their volume by withdrawing the inert material thereof; h) Separating the cured stiffened structure (11) from the tool (31). The invention also relates to the auxiliary male tools.

Mobile surfaces for aircraft with sealed slots

A lifting or stabilising component (11) for an aircraft comprising in the area of its trailing edge a rotary control element (13) rotating around an axis (61) with at least one slot (21) between the tip (15) of the component (11) and the control element (13), the edges of which are configured such that the distance therebetween, that is, the dimension of the slot (21), is constant for different angles of deflection of the control element (13), being located between them a seal (23) that assure the aerodynamic continuity of the component (11) when the control element (13) is at rest. In a preferred embodiment, the component (11) is a horizontal tail stabilizer and the control element (13) is a rudder.

Integrated composite skin for aircraft mid-box lifting surfaces

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).

System for sealing variable gaps for scale models tested in an aerodynamic tunnel

The System for sealing for variable spaces comprising sealing means ( 3 ) for at least partially packing a space ( 4 ) between two pieces ( 1, 2 ) of a model, characterised in that at least a first piece ( 1 ) of the model presents, on its side facing a second piece ( 2 ) of the model, at least one recess ( 5 ) with a front end ( 5 a) open and a back wall ( 5 b), with the sealing means, comprising an elastic and flexible sealing body ( 3 ), being located in the front end ( 5 a), and the back wall ( 5 b) containing a set of spring elements ( 6, 7 ) among which are to be found coil springs or elastic strips, at least partially compressed by the sealing body ( 3 ) and which push the latter against the surface of the second piece ( 1 b) of the model.

Piece made of composite material with areas of different thickness

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).

Mechanism for eliminating limit cycle oscillations on servocontrolled aerodynamic control surfaces

Servocontrolled aerodynamic control surfaces, such as the direction rudder ( 1 ) of an airplane, are equipped with servoactuators ( 5 ). The bearing supports ( 3 ) of the rotation axis ( 2 ) or hinged axis give rise to the formation of free play for the manufacturing tolerances, that will be more pronounced with the wear of the bearings or joint elements. The mechanism ( 8 ) in question includes an elastic element ( 10 ) fixed to the structure ( 4 ) wherein the control surface ( 1 ) is jointed, connected rigidly to an articulated connecting rod ( 12 ) having a following roller ( 13 ) of a cam ( 9 ) integral to the control surface ( 1 ). The cam ( 9 ) profile produces a continuous nonlinear, angular deformation, of the elastic element in order to determine a force torque on the control surface ( 1 ) that is maximum in the neutral position of the control surface and manages to be eliminated in the extreme side deviations.

Fitting with torsion box, of plastic material reinforced with carbon fibre, for coupling a drive motor / spindle unit for trimming of a horizontal stabiliser of an aircraft

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.

Process and tools for manufacturing composite ring frames

The present invention relates to a process for manufacturing composite ring frames for aeronautical fuselages by means of the application of the RTM technology to two preforms with C- and L-shaped sections manufactured using two tools (21, 55) in the following steps: providing the material; hot-forming planar rectangular laminates (41); hot-forming laminates of right angle section (51) on one part of right angle section of the first tool (21), placing an elastic membrane (55) and applying a temperature and vacuum cycle; hot-forming the preforms into a C shape (11) and L shape (13) on a second curved tool (55) by deforming said laminates of right angle section (51) thereon, and applying a temperature and vacuum cycle. The invention also relates to said tools (21, 55).

Actuator safety attachment device

Safety attachment device for an actuator ( 1 ) attached at its ends respectively to a first and a second part ( 2, 3 ) which are connected to one another also by means of a reaction bar ( 4 ) which is located on the top of the actuator ( 1 ), said device comprises an upper part ( 5, 5 a, 5 b) immobilized on the reaction bar ( 4 ) and side arms ( 6, 7 ) which clasp the sides ( 1 c, 1 d) of the actuator ( 1 ), both having windows ( 8, 9 ) through which respective pins ( 10, 11 ) joined to these sides ( 1 c, 1 d) pass, the pin ( 10, 11 ) thus being situated without making contact with the edges of the window when the actuator is attached to the parts ( 2, 3 ), and each window ( 8, 9 ) being of a horizontal extension such that its side edges do not touch the pin ( 10, 11 ) which passes through said window ( 8, 9 ) when the actuator ( 1 ) is disconnected from any of the parts ( 2, 3 ).

Quality control process for a structural bonded joint

The invention relates to a quality control process for an adhesive joint of two sub-components of a structure comprising the following steps: a) Providing the structure of composite material to be controlled; b) Providing at least one premanufactured testing device (1) representative of one of the sub-components of the structure; c) Bonding the at least one premanufactured testing device (1) to the other sub-component (7) of the structure in conditions similar to those of the real adhesive joint of the sub-components; d) Carrying out at least one mechanical test on the at least one premanufactured testing device (1) which allows assessing the quality of the adhesive joint.

Cover for an aircraft access opening

The invention relates to a cover for the access opening in a composite material aircraft structure, characterized in that it comprises the following elements: an outer cover (18) arranged on the outer part of the aircraft skin (28), the carbon fiber layers (14) forming the skin (28) being deformed towards the inner part of the aircraft structure, leaving enough space between the aircraft structure and the outer part of the skin (28) for the seating of the outer cover (18). a composite material plate (15) arranged in the outer part of the skin (28) and surrounding the area for installing the outer cover (18); an inner cover (27) arranged on the inner part of the aircraft skin (28).

Diffusion nozzle for an ultrasonic probe

A suitable diffusion nozzle ( 20 ) for an ultrasounds probe using water as coupling agent between a piece under inspection and an acoustic transductor ( 13 ) such that the water is driven towards the surface of the material through a propulsion channel ( 21 ) surrounded by a plurality of diffusion channels ( 23 ) through which air is injected under pressure, forming an insulating screen capable of repelling the drops of water generated on the jet of water striking the surface of the material.

Device for facilitating the lowering of a landing gear by gravity

Device for facilitating the lowering of a landing gear by gravity, comprising at least on roller device ( 4 ) positioned on at least one closing panel ( 2 ) of the hatch for the landing gear in a contact area with the tire ( 1 a) of the landing gear of an aircraft ( 7 ) on initiating its descent, the roller device ( 4 ) comprising a rigid roller device which can rotate freely about its axis ( 4 a), and being positioned such that its axis is perpendicular to the path of the tire ( 1 a) so that the tire will make the roller device ( 4 ) turn without sliding over it.

Method for evaluating the effect of an electric discharge on a composite material

The present invention relates to a method for evaluating the effect of an electric discharge on an aircraft structure through the evaluation of the damage caused by the effect of the thermal heating generated by the mentioned discharge on the structure, including the steps of: applying an electric discharge on the material of the aircraft structure using an intensity generator; distributing in an electric mesh with resistive electric elements the material of the aircraft structure; calculating the intensities running through the electric mesh of the material of the aircraft structure from the intensity introduced in the generators; calculating with the previous intensities the heat which is dissipated in each of the elements of the electric mesh of the material of the aircraft structure; calculating the distribution of temperatures in each of the elements of the electric mesh of the material of the aircraft structure; determining the elements of the electric mesh of the material of the ...

Method and system for a quick calculation of aerodynamic forces on an aircraft

A computer-aided method suitable for assisting in the design of an aircraft by providing the dimensioning aerodynamic forces and other relevant values comprising the following steps: a) Selecting a set of parameters of the aircraft, being said aerodynamic forces and other relevant values dependant of said parameters; b) Performing flow field CFD RANS computations for a number N1 of different combination of values of said parameters; c) Obtaining said aerodynamic forces and other relevant values for whatever combination of values of said parameters through a reduced-order model, generated by computing the POD modes of the flow variables, expanding the flow variables using said POD modes and obtaining the POD coefficients of said expanded flow variables using a genetic algorithm that minimizes the error associated to the expansion of the Navier-Stokes equations. The invention also refers to a system for carrying out said method.

Process of manufacturing composite panels with U-shaped stiffening members

The present invention relates to a process for manufacturing panels (9) for aeronautical structures with U-shaped stiffening members (13) and I-shaped stiffening members (17) between their webs comprising the following steps: providing laminates for shaping the skin (11) on the curing tool (31); providing planar laminates (23) for shaping the stiffening members (13, 17); shaping the U-shaped stiffening members (13) on individual shaping tools (33) and placing the I-shaped stiffening elements (17) in said tools; grouping said individual shaping tools (33) together on an assembly tool (41); placing the group of stiffening members (13, 17) on the skin (11); placing a vacuum bag (55) on the assembly with the aid of profiles (59); consolidating said assembly by means of a curing process under suitable pressure and temperature conditions using external tools (60) to assure verticality of the webs of the stiffening members (13, 17).

Fairing system for a horizontal stabiliser of an aircraft and process for installing said system

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 mentioned supports ( 30 ) and offer a good aerodynamic performance of the assembly. The invention also relates to a process for installing such a fairing system on the horizontal stabiliser ( 2 ) of an aircraft.

Integrative structure for aircraft fairing

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.

Rib-fitting

Rib-fitting as a component ( 9 ) of a torsion box ( 1, 3 ) of an aircraft lift surface made of composite material as a single piece, comprising: a substantially planar web ( 31 ) with a first part in the form of a lug ( 33 ) and a second part in the form of a rib web ( 35 ); two flanges ( 39 ) to attach the web ( 31 ) with the webs of each of the ends of the front or back spar ( 11, 13 ) to which it is joined; flanges ( 43 ) for being attached to the upper and lower skins ( 19, 21 ) of the torsion box; gaps ( 45 ) in the areas of intersection with the front or back spar caps ( 11, 13 ) to which the component ( 9 ) is connected and with the reinforcement stringers ( 25 ) of the skins ( 19, 21 ). The invention also comprises and assembly process for the rib-fitting in multispar torsion boxes.

肋接头

作为飞机冲面的抗扭箱(1，3)的部件(9)的肋接头由复合材料制成为单个部件，包括：大致平坦的连接板(31)，具有突出部(33)形式的第一部分和肋形连接板(35)形式的第二部分；两个凸缘(39)，用于将连接板(31)连接至其所连接到的前加强杆或后加强杆(11，13)的每个端部的连接板；凸缘(43)，用于连接至抗扭箱的上外壳和下外壳(19，21)；间隙(45)，位于与部件(9)连接到的前加强杆盖或后加强杆盖(11，13)交叉并与外壳(19，21)的加强纵梁(25)交叉的区域中。本发明还包括用于将肋接头组装在多加强杆抗扭箱中的组装工艺。

surface treatment method of composite material structures with atmospheric plasma beams

MéTODO DE TRATAMENTO SUPERFICIAL DE ESTRUTURAS DE MATERIAL COMPOSTO COM FEIXES DE PLASMA ATMOSFéRICO. A presente invenção refere-se a um método de tratamento superficial de zonas predeterminadas de uma estrutura de material composto com, pelo menos, um feixe de plasma em pressão atmosférica, que é produzido por um gerador de plasma dotado de um bocal de emissão, para facilitar sua união adesiva a outra estrutura, no qual: a) o feixe de plasma emitido pelo bocal se projeta sobre a estrutura de material composto a uma distância a partir da mesma compreendida entre 0,2 e 10 cm; e b) o feixe de plasma se projeta sobre a estrutura de material composto com um ângulo de incidência compreendido entre 75° e 105°. O método é preferivelmente aplicado a estruturas de fibra de carbono ou fibra de vidro e resina epóxi ou resina de bismaleimida. Outras variáveis relevantes do procedimento são a potência do feixe de plasma e a velocidade de tratamento. SURFACE TREATMENT METHOD OF COMPOUND MATERIAL STRUCTURES WITH ATMOSPHERIC PLASMA BEAMS. The present invention relates to a method of surface treatment of predetermined zones of a composite material structure having at least one atmospheric pressure plasma beam which is produced by a plasma generator having an emission nozzle to facilitating its adhesive bonding to another structure in which: (a) the plasma beam emitted by the nozzle projects over the composite material structure at a distance from 0.2 to 10 cm from it; and (b) the plasma beam projects over the composite material structure having an angle of incidence of between 75 ° and 105 °. The method is preferably applied to carbon fiber or fiberglass structures and epoxy resin or bismaleimide resin. Other relevant procedural variables are plasma beam power and treatment rate.

Fairing system for an aircraft horizontal stabiliser and method for installing said system

本发明涉及一种用于航空器的水平稳定器的整流罩系统和用于安装所述整流罩系统的方法，使得整流罩系统密封航空器的水平稳定器(2)和机身(1)。所述整流罩系统包括配置在航空器的水平稳定器(2)上的主整流罩(50、51)，所述主整流罩(50、51)稳固连接到所述水平稳定器(2)，而所述水平稳定器(2)相对于机身(1)可调试，所述整流罩系统还包括辅助整流罩(80、81)，使得主整流罩(50、51)通过沿主整流罩(50、51)的翼弦的方向配置的支撑件(30)连接到水平稳定器(2)，所述支撑件(30)位于所述主整流罩(50、51)的相对于机身(1)的外部部分，所述辅助整流罩(80、81)配置在所述支撑件(30)的外部部分，所述辅助整流罩(80、81)的目的是用于覆盖所述支撑件(30)和为该组件提供优良空气动力性能。本发明还涉及将该整流罩系统安装在航空器的水平稳定器(2)上的方法。

MEAT SYSTEM FOR AIRCRAFT HORIZONTAL STABILIZER AND INSTALLATION PROCEDURE OF SUCH SYSTEM.

Sistema de carenas para estabilizador horizontal de aeronave y procedimiento de instalación de dicho sistema, tal que el sistema de carenas para el sellado del estabilizador horizontal (2) y el fuselaje (1) de la aeronave comprende una carena primaria (50, 51) dispuesta sobre el estabilizador horizontal (2) de la aeronave, siendo la citada carena primaria (50, 51) solidaria al citado estabilizador horizontal (2) que, a su vez, es trimable con respecto al fuselaje (1), comprendiendo el sistema de carenas a su vez una carena secundaria (80, 81), tal que la carena primaria (50, 51) se une al estabilizador horizontal (2) a través de unos soportes (30) dispuestos en la dirección de la cuerda de la carena primaria (50, 51), estando dichos soportes (30) situados en la parte exterior de dicha carena primaria (50, 51) con respecto al fuselaje (1), estando dispuesta la carena secundaria (80, 81) en la parte exterior de los soportes (30), siendo la misión de dicha carena secundaria (80, 81) la de cubrir los citados soportes (30) y ofrecer un buen comportamiento aerodinámico del conjunto. La invención se refiere también a un procedimiento de instalación de un sistema de carenas tal sobre el estabilizador horizontal (2) de una aeronave. Proper system for horizontal aircraft stabilizer and installation procedure for said system, such that the system of projections for sealing the horizontal stabilizer (2) and the fuselage (1) of the aircraft comprises a primary fairing (50, 51) arranged on the horizontal stabilizer (2) of the aircraft, said primary cage (50, 51) being integral with said horizontal stabilizer (2) which, in turn, is trimable with respect to the fuselage (1), the hunt system comprising in turn a secondary fairing (80, 81), such that the primary fairing (50, 51) joins the horizontal stabilizer (2) through supports (30) arranged in the direction of the primary fairing rope ( 50, 51), said supports (30) being located on the outside of said primary fairing (50, 51) with ...

Fairing system for an aircraft horizontal stabiliser and method for installing said system

本发明涉及一种用于航空器的水平稳定器的整流罩系统和用于安装所述整流罩系统的方法，使得整流罩系统密封航空器的水平稳定器(2)和机身(1)。所述整流罩系统包括配置在航空器的水平稳定器(2)上的主整流罩(50、51)，所述主整流罩(50、51)稳固连接到所述水平稳定器(2)，而所述水平稳定器(2)相对于机身(1)可调试，所述整流罩系统还包括辅助整流罩(80、81)，使得主整流罩(50、51)通过沿主整流罩(50、51)的翼弦的方向配置的支撑件(30)连接到水平稳定器(2)，所述支撑件(30)位于所述主整流罩(50、51)的相对于机身(1)的外部部分，所述辅助整流罩(80、81)配置在所述支撑件(30)的外部部分，所述辅助整流罩(80、81)的目的是用于覆盖所述支撑件(30)和为该组件提供优良空气动力性能。本发明还涉及将该整流罩系统安装在航空器的水平稳定器(2)上的方法。

EXISTING HOLE SEALING SYSTEM BETWEEN THE FUSELAGE AND HEIGHT TIMON OF THE ORIENTABLE HORIZONTAL STABILIZER OF AN AIRPLANE, EXTENDED WITH AN AERODYNAMIC SEALING SIDE OF THE EXISTING OPENING BETWEEN THE FUSELAGE AND ORIENTING STABILIZER.

Sistema de sellado del hueco existente entre el fuselaje y el timón de altura del estabilizador horizontal orientable de un avión, extendido con una carena aerodinámica de sellado de la abertura existente entre el fuselaje y el estabilizador horizontal orientable. Sistema de sellado del hueco (2) entre el fuselaje (3) y el timón de altura 4 del estabilizador horizontal orientable (5) de un avión, extendido con una carena aerodinámica (8) de sellado de la abertura (24) entre el fuselaje (3) y dicho estabilizador horizontal (5) que comprende un cuerpo principal (1) y una carena aerodinámica (8) que se extiende desde el cuerpo principal (1) en continuidad con éste, y una pluralidad de primeros perfiles de sellado elásticos (9) entre la primera superficie (11) del cuerpo principal (1) incluyendo a la carena aerodinámica (8) y la superficie externa del fuselaje (3) y contactando éstas, y una pluralidad de segundos perfiles de sellado elásticos (13) entre la segunda superficie (12) del cuerpo principal (1) y el primer extremo del timón de altura (15) y contactando éstas, de tal forma que se realiza el sellado del hueco (2) y la abertura (24) y se produce una continuidad aerodinámica entre el estabilizador horizontal orientable (5), el fuselaje (3) y el timón de altura (4) cuando el timón de altura (4) está en reposo en el plano del estabilizador horizontal orientable (5) para cualquiera de las posiciones de orientación del estabilizador horizontal orientable (5).

RIB STRUCTURE FOR TORSION DRAWERS OF A WING OR A STABILIZER OF AN AIRCRAFT.

Estructura de costilla para cajones de torsión de un ala o de un estabilizador de una aeronave. Rib structure for torsion drawers of a wing or a stabilizer of an aircraft. Elemento de rigidización vertical para una estructura de costilla que comprende un larguero frontal (1a), un larguero trasero (1b), un alma (2) que se extiende entre los largueros (1a, 1b), larguerillos superiores (5) e inferiores (6) respectivamente unidos internamente a un revestimiento superior (4a) y a un revestimiento inferior (4b), y dispuestos en sentido sustancialmente perpendicular al alma (2) de la costilla; comprendiendo el elemento de rigidización vertical (3) un cuerpo vertical (3a) unido al alma (2) de la costilla, un primer ala lateral (3d) que se extiende en dirección del larguero frontal (1a) y un segundo lateral (3e) que se extiende en dirección del larguero trasero (1b), siendo susceptible al menos una de las partes extremas (3b, 3c) del cuerpo vertical (3a) del elemento de rigidización vertical (3) de quedar unida directamente a un larguerillo (5, 6). Vertical stiffening element for a rib structure comprising a front stringer (1a), a rear beam (1b), a soul (2) that extends between the stringers (1a, 1b), upper (5) and lower (6) stringers respectively attached internally to a top lining (4a) and to a lower lining (4b), and arranged in the direction substantially perpendicular to the soul (2) of the rib; the vertical stiffening element (3) comprising a body vertical (3a) attached to the soul (2) of the rib, a first wing lateral (3d) extending in the direction of the front crossbar (1a) and a second side (3e) extending in the direction of the crossbar rear (1b), at least one of the end parts being susceptible (3b, 3c) of the vertical body (3a) of the stiffening element vertical (3) to be attached directly to a stringer (5, 6).

Fairing system for an aircraft horizontal stabiliser and method for installing said system

The invention relates to a fairing system for an aircraft horizontal stabiliser and to a method for installing said system, such that the fairing system for sealing the horizontal stabiliser (2) and the fuselage (1) of the aircraft includes a primary fairing (50, 51) arranged on the horizontal stabiliser (2) of the aircraft, said primary fairing (50, 51) being rigidly connected to the aforementioned horizontal stabiliser (2) which, in turn, is trimmable relative to the fuselage (1), said fairing system in turn including a secondary fairing (80, 81), such that the primary fairing (50, 51) is attached to the horizontal stabiliser (2) by means of supports (30) arranged in the direction of the chord of the primary fairing (50, 51), said supports (30) being located on the outside of said primary fairing (50, 51) relative to the fuselage (1), the secondary fairing (80, 81) being arranged on the outside of the supports (30), said secondary fairing (80, 81) being responsible for covering the aforementioned supports (30) and ensuring good aerodynamic performance of the whole. The invention also relates to a method for installing a fairing system on the horizontal stabiliser (2) of an aircraft.

Optimized aircraft manhole

The present invention relates to an aircraft manhole comprising a skin panel ( 1 ) of the aircraft structure, an outer door ( 6 ) and an inner door ( 5 ) of the aircraft manhole being positioned on the mentioned skin panel ( 1 ), further comprising stringers ( 3 ) installed on the skin panel ( 1 ) positioned on and attached to the skin ( 1 ), fixing edges ( 8 ) partially projecting through the manhole on the side for positioning the outer door ( 6 ) and inner door ( 5 ), further comprising reinforcements ( 4 ) arranged between the outer door ( 6 ) and the inner door ( 5 ) of the manhole that are attached to the skin ( 1 ), said reinforcements ( 4 ) forming the fixing edges ( 8 ) on one of its sides, and a cavity ( 7 ) arranged in the skin ( 1 ) to allow installing the outer door ( 6 ), thus maintaining the aerodynamic surface of the assembly as well as its tightness.

STRUCTURE FOR THE COUPLING OF THE FLAVOR OF AN AIRCRAFT

Estructura para el acoplamiento de la carena (7) de una aeronave, estando dicha carena (7) situada entre el estabilizador horizontal y el estabilizador vertical o sección de fuselaje (6) de dicha aeronave, comprendiendo dicha estructura unos herrajes (2) en forma de angulares, comprendiendo a su vez dichos herrajes una superficie vertical y una superficie horizontal, uniéndose los herrajes (2) a través de su superficie vertical a la carena (7) mediante unos elementos de fijación (12) desmontables, comprendiendo además dicha estructura entre la carena (7) y la superficie vertical del herraje (2) un elemento elástico (13) deformable, tal que dicho elemento elástico (13) es capaz de absorber el juego de deformaciones necesario y las tolerancias de fabricación y montaje de los herrajes (2), de la carena (7), del estabilizador horizontal y del estabilizador vertical o sección de fuselaje (6) mediante diferentes pares de apriete de los elementos anteriores. Structure for coupling the fairing (7) of an aircraft, said fairing (7) being located between the horizontal stabilizer and the vertical stabilizer or fuselage section (6) of said aircraft, said structure comprising fittings (2) in shape angular, said hardware in turn comprising a vertical surface and a horizontal surface, joining the hardware (2) through its vertical surface to the cowl (7) by means of detachable fixing elements (12), also comprising said structure between the fairing (7) and the vertical surface of the hardware (2) a deformable elastic element (13), such that said elastic element (13) is capable of absorbing the necessary deformation set and the manufacturing and assembly tolerances of the hardware ( 2), of the fairing (7), of the horizontal stabilizer and of the vertical stabilizer or fuselage section (6) by different tightening torques of the previous elements.

Method of producing a joining structure for an aircraft fairing

Estructura para el acoplamiento de la carena (7) de una aeronave, estando dicha carena (7) situada entre el estabilizador horizontal y el estabilizador vertical o sección de fuselaje (6) de dicha aeronave, comprendiendo dicha estructura unos herrajes (2) en forma de angulares, comprendiendo a su vez dichos herrajes una superficie vertical y una superficie horizontal, uniéndose los herrajes (2) a través de su superficie vertical a la carena (7) mediante unos elementos de fijación (12) desmontables, comprendiendo además dicha estructura entre la carena (7) y la superficie vertical del herraje (2) un elemento elástico (13) deformable, tal que dicho elemento elástico (13) es capaz de absorber el juego de deformaciones necesario y las tolerancias de fabricación y montaje de los herrajes (2), de la carena (7), del estabilizador horizontal y del estabilizador vertical o sección de fuselaje (6) mediante diferentes pares de apriete de los elementos anteriores.

Structure of the load introduction zone in the rear end of an aircraft

Load introduction zone in the rear end of an aircraft comprising receiving elements ( 3, 4 ) of the loads of the horizontal and vertical tail stabilizers connected to structural elements ( 2, 8, 9 ) wherein: a) said structural elements are the cladding ( 8 ), two contiguous load-bearing frames ( 2 ) with a Π or omega cross-section and, at least, a third frame ( 9 ); b) the receiving element of the horizontal tail stabilizer load is a first fitting ( 3 ) structured as a torque box, transversely arranged between said load-bearing frames ( 2 ) and connected to them, and the receiving elements of the vertical tail stabilizer loads are some second fittings ( 4 ) coupled to these frames ( 2, 9 ); c) said frames ( 2, 9 ) and said first and second fittings ( 3, 4 ) are entirely made of composite material.

Structure of the load-introduction area in the fuselage rear section of an aircraft

一种在飞行器的后端中的负载引导区域，包括：连接至结构元件(2，8，9)的水平和垂直尾部稳定装置的负载的接收元件(3，4)，其中：a)所述结构元件是覆层(8)、具有∏或Ω横截面的两个相邻的负载承受框架(2)和至少第三框架(9)；b)所述水平尾部稳定装置负载的接收元件为第一配件(3)，该第一配件(3)构造成抗扭箱，且横向地布置在所述负载承受框架(2)之间且连接至它们上，垂直尾部稳定装置的负载的接收元件是耦接至所述框架(2，9)的一些第二配件(4)；c)所述框架(2，9)和第一和第二配件(3，4)全部由复合材料制成。

CONFIGURATION OF THE PNEUMATIC INSTALLATION IN THE COLA ZONE OF AN AIRCRAFT.

Configuración de la instalación neumática en la zona de cola de un avión.#Que comprende una unidad auxiliar de potencia (1) de suministro de potencia neumática y eléctrica, cuya salida de potencia neumática está unida a una conducción (2) que convencionalmente incorpora codos (2) y elementos de compensación (6) de dilatación térmica y de compensación de las vibraciones que produce la APU (1) en su funcionamiento. Se caracteriza porque la salida de potencia de la APU (1) está dispuesta en su parte frontal en la que conecta la conducción que está constituida por un tubo (2a) recto inclinado descendente hacia la parte inferior del avión en el que se intercalan los elementos de compensación de dilatación térmica y de compensación (6) de las vibraciones, para reducir la longitud de la conducción (2) y eliminar los codos, de forma que se reducen las pérdidas de presión considerablemente. Configuration of the pneumatic installation in the tail area of an aircraft. # Which comprises an auxiliary power unit (1) for supplying pneumatic and electrical power, whose pneumatic power output is connected to a conduit (2) that conventionally incorporates elbows (2) and compensation elements (6) for thermal expansion and vibration compensation produced by the APU (1) during operation. It is characterized in that the power output of the APU (1) is arranged in its front part in which it connects the conduit which is constituted by a straight tube (2a) inclined downwardly towards the bottom of the plane in which the elements are intercalated of thermal expansion compensation and vibration compensation (6), to reduce the length of the conduit (2) and eliminate the elbows, so that pressure losses are considerably reduced.

INTEGRATION SYSTEM OF REMACHABLE NUTS IN PLASTIC PARTS.

Sistema de integración de tuercas remachables en piezas de plástico. Integration system of rivet nuts in plastic pieces. El sistema está previsto para fijar tuercas metálicas remachables (1) a piezas de plástico (2), realizándose la fijación mediante remachado por ultrasonidos o cualquier otro medio mecánico. Cada pieza de plástico (1) obtenida en el correspondiente proceso de fabricación, preferentemente por inyección, está dotada de unos tetones (4) que se alojan en orificios de la tuerca remachable (1), efectuándose seguidamente el remachado por ultrasonidos u otro medio mecánico sobre esos tetones (4), produciéndose la fusión de éstos y su correspondiente soldadura sobre la tuerca remachable (1) y, en definitiva, la fijación e integración de la misma sobre la pieza de plástico (2). El sistema de remachado referido para integración de tuercas remachables en piezas de plástico, permite unir entre sí piezas de plástico con otras piezas, que pueden ser metálicas o no metálicas, pertenecientes a una estructura. The system is intended to fix nuts rivetable metal (1) to plastic parts (2), the fixation by ultrasonic riveting or any other means mechanic. Each piece of plastic (1) obtained in the corresponding manufacturing process, preferably by injection, is equipped of some lugs (4) that are housed in nut holes rivet (1), followed by riveting by ultrasound or other mechanical means on those lugs (4), producing the fusion of these and their corresponding welding on the rivet nut (1) and, in short, the fixing e Integration of it on the plastic part (2). The system of riveted referred for integration of rivet nuts in plastic parts, allows to join together plastic parts with other parts, which may be metallic or nonmetallic, belonging to a structure.

Method for surface treatment of composite material structure by using atmospheric plasma beam

本发明涉及一种使用至少一种处于大气压下的等离子体束表面处理复合材料结构体的预定区域的方法，所述方法是为了有助于所述复合材料结构体对另一结构体的粘合剂粘接，所述处于大气压下的等离子束由安置有发射喷嘴的等离子体发生器产生，在所述方法中：a)从包括在0.2和10cm之间的距离，将穿过喷嘴发射的等离子体束投射在复合材料结构体上，所述等离子体束可以包括反应性气体；b)以包括在75°和105°之间的入射角将等离子体束投射在复合材料结构体上。优选将本方法应用于碳纤维或玻璃纤维和环氧树脂或双马来酰亚胺树脂结构体。本方法中的其它相关变量是等离子体束功率和处理速率。

Tool and process for manufacturing pieces of composite materials outside an autoclave

This invention relates to a tool ( 9 ) for manufacturing pieces of composite material outside an autoclave comprising a stacking table ( 11 ); a movable head ( 15 ) on said stacking table ( 11 ) provided with: automatic means for placing tapes ( 19 ) of composite material; compacting means for compacting the composite material and microwave emission means ( 25 ) to cure the composite material. The invention also relates to a process for manufacturing pieces of composite material outside an autoclave, comprising the following steps: a) Placing composite material in the form of tapes on a tool with the shape of the piece to be manufactured, compacting it and partially curing it after its placement until completing a layer of the piece; b) Repeating step a) until completing the stacking of the piece and c) Curing the last layer of the piece until the required degree of curing.

Method to monitor structural damage occurrence and progression in monolithic composite structures using fibre bragg grating sensors

A method of monitoring structural damage in a composite structure manufactured by co-curing, co-bonding or secondary bonding of several sub-components (7, 8), using fibre optic Bragg grating sensors attached or embedded to or between (in the bonding line) said sub-components, comprising a first step of measuring the wavelength spectra (31, 33, 40) of said Bragg grating sensors at the end of the manufacturing of the part, and in a known load condition, considered as reference and a second step of identifying the occurrence of a failure of the structure and the progress of said failure detecting the release of the residual stresses/strains stored during the curing process by measuring changes (31 to 32, 33 to 34, 40 to 41) with respect to said reference wavelength spectra.

Automatic dismoulding system

An automatic dismoulding system for moulding modules (7) located in a curing tool (12), comprising: a programmable robot (1) comprising an articulated arm (la) with one end free, actuation means (13, 14) for conferring a horizontal rotary movement and a vertical tilting movement to the articulated arm (la), and a head (2) located in the free end of the articulated arm (1a) with at least one vacuum suction pad (11) able to adhere itself to a surface (16, 17) of an object; a voiding system (15) for generating a vacuum; referencing means (3) for referencing a vertical and horizontal position of each object (6, 7); and programmable controlling means (18) which govern said actuation means (13, 14) so that the articulated arm (1a) can position the suction pad (11) on that surface (16, 17) of the object (6, 7) in an initial position and transport it to a final position.

Cover for an aircraft access opening

The invention relates to a cover for the access opening in a composite material aircraft structure, characterized in that it comprises the following elements: an outer cover ( 18 ) arranged on the outer part of the aircraft skin ( 28 ), the carbon fiber layers ( 14 ) forming the skin ( 28 ) being deformed towards the inner part of the aircraft structure, leaving enough space between the aircraft structure and the outer part of the skin ( 28 ) for the seating of the outer cover ( 18 ). a composite material plate ( 15 ) arranged in the outer part of the skin ( 28 ) and surrounding the area for installing the outer cover ( 18 ); an inner cover ( 27 ) arranged on the inner part of the aircraft skin ( 28 ).

Method for producing composite-material structures with collapsible tooling

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.

Aircraft having a lambda-box wing configuration

Aircraft having a lambda-box wing configuration, comprising a fuselage (1), a propulsion system (5), a first pair of swept-back airfoils (2), connected to the top forward portion of the fuselage (1), a second pair of swept-forward airfoils (3), connected to the lower rear portion of the fuselage (1) at a point of the said fuselage (1) aft of the connection of the swept-back airfoils (2), and a third pair of substantially vertical airfoils (4), the tips of the swept-forward airfoils (3) being connected to the lower side of the swept-back airfoils (2) at an intermediate point of the span of the said swept-back airfoils (2), by means of the substantially vertical airfoils (4), the swept-back airfoils (2) having a higher aspect ratio than that of the swept-forward airfoils (3), which makes the swept-back airfoils (2) have a reduced induced drag without penalizing their weight.

Device and method for the manufacture of three dimensional beam type elements in composite material

The invention relates to a device for the manufacture of three-dimensional beam type elements in composite material with reinforcing fibres preimpregnated with polymeric resins, starting with laminates (21, 22) layed and precut without polymerization, said device comprising a head (15) that comprises in its turn a roller train (17), it being possible for said head (15) to move longitudinally along a fixed bedplate, in such a way that, as said head (15) moves, the roller train acts on the laminates (21, 22) without polymerization, compacting and forming them, thus forming the laminates (21, 22) with their final geometry in one go, in such a way that said laminates (21, 22) are ready for their subsequent integration. The invention also relates to a method of manufacture of three-dimensional beam type elements in composite material.

Ventral fairing for an aircraft

Ventral fairing for an aircraft provided between a left wing ( 2 a ) and a right wing ( 2 b ), which comprises a lower central section ( 3 a ) of convex transverse cross-section located between the left lower surface ( 2 e ) of the left wing ( 2 a ) and the right lower surface ( 2 f ) of the right wing ( 2 b ), and with an angle (φ) between the cross-section of the fairing ( 3 ) and the line of cross-section corresponding to the lower surface ( 2 e, 2 f ) of each wing ( 2 a, 2 b ) of less than 90 degrees, separate left ( 3 c ) and right ( 3 b ) lower lateral sections of concave transverse cross-section which are respectively extended between the lower surface ( 2 e, 2 f ) of each wing ( 2 a, 2 b ) and the central section ( 3 a ); separate first lines of inflection ( 4 a, 4 b ) between the lower surface ( 2 e, 2 f ) of each wing ( 2 a, 2 b ) and the corresponding lateral section ( 3 c, 3 b ); separate second lines of inflection ( 4 c, 4 d ) between the lateral sections ( 3 c, 3 b ) and the central section ( 3 a ).

Optimized configuration of engines for aircraft

The invention relates to a configuration of engines (3) for aircraft located in the rear part of the fuselage (2) of said aircraft, the engines (3) being attached in a fixed manner by pylons (5) to the structure of the aircraft, said structure comprising a torsion box (14) which traverses the fuselage (2) and is used to attach the pylons (5), the fuselage (2) comprising an opening (4) allowing the passage of the suspension pylons (5) for the engines (3), said configuration further comprising a pivoting area (8), an actuator (7) and a fitting (6) through which the actuator (7) is attached to the suspension pylons (5) and to the torsion box (14) of the aircraft, such that the assembly formed by the actuator (7) and the fitting (6) allow balancing the pylon (5) and engine (3) assembly of the aircraft through the pivoting area (8), thus achieving controllable and optimal thrust vectoring of the aircraft for each flight phase.

Milling head in tricepts with electro-mechanical depth control

It comprises a nozzle (4) with an inlet (5) for introducing of a rivet's (2) extra stem (1); and in front of which a milling tool (6) is included to trim the stem (1) of the different rivets made on a piece. The milling tool is anchored by means of a trimming compensation spring (11), and the nozzle (4) is fixed with a separation height with respect to the milling tool (6) which is the same as the trimming height, in such a way that the inlet (5) determines a mechanical trimming depth limit when contacting the rivet (2) head surface (17). The nozzle (4) is adjustable in height in order to select the trimming height, being adaptable to the different rivets existing in the market. It comprises an electrical limit (16, 14) from which, after the mechanical limit has occurred, the retreat of the head of the rivet in order to allow making a new rivet,

Procedure for the preparation and cleaning of tools used for manufacturing composite material components, and the corresponding device

The procedure includes: a first phase of blowing and suction; a second phase of cleaning by laser where the tool (3) on a positioner (4) is held under a laser head (2); a third phase of applying a releasing agent by spray, and a fourth phase for drying off this releasing agent by means of the aforementioned head (2).

Electric insulation device for use in fuel lines

An electrical insulating device for a linear component forming part of a fuel system subjected to risks of possible external electrical discharges formed by a part (13) made of insulating material attached to two metal flanges (15, 17) that allow inserting the device (11) between two parts of said linear component, said part (13) being formed by a cylindrical body (21) with a castellated crown (23) on its outer side, formed by longitudinal (25) and transverse (27) segments, and said metal flanges (15, 17) having transverse projections (31, 33) facing the transverse segments (27) of the castellated crown (23) for facilitating their attachment by suitable means (39),

Optimized aircraft manhole

The present invention relates to an aircraft manhole comprising a skin panel (1) of the aircraft structure, an outer door (6) and an inner door (5) of the aircraft manhole being positioned on the mentioned skin panel (1), further comprising stringers (3) installed on the skin panel (1) positioned on and attached to the skin (1), fixing edges (8) partially projecting through the manhole on the side for positioning the outer door (6) and inner door (5), further comprising reinforcements (4) arranged between the outer door (6) and the inner door (5) of the manhole that are attached to the skin (1), said reinforcements (4) forming the fixing edges (8) on one of its sides, and a cavity (7) arranged in the skin (1) to allow installing the outer door (6), thus maintaining the aerodynamic surface of the assembly as well as its tightness.

A REMOVAL AND / OR FIXED TRACTION-COMPRESSION UNION PROCEDURE FOR APPLICATION TO COMPOSITE MATERIALS.

Un procedimiento de unión a tracción-compresión de carácter desmontable y/o fijo para aplicar a materiales compuestos, en el que se pliega al menos una vez un laminado básico (1) de material compuesto para obtener al menos una faldilla (2) que forma un ángulo próximo a 90º con el resto del laminado, utilizando un radio de transición entre la faldilla y el resto del laminado, se aplica sobre el laminado básico un conjunto de placas de distribución (5, 6) para estabilizar la faldilla y producir una distribución de cargas de carácter uniforme, se practican taladros en cada conjunto de faldilla-placas de distribución y se introducen en dichos taladros unos bulones de tracción (4) con una pretensión inicial, produciéndose así la conexión entre los elementos que deben ser unidos. La invención es aplicable a estructuras aerodinámicas de aviones.

System for sealing variable gaps for scale models tested in an aerodynamic tunnel

The system for sealing for variable spaces comprising packing means (3) for at least partially packing a space (4) between two pieces (1, 2) of a model, characterised in that at least a first piece (1) of the model presents, on its side facing a second piece (2) of the model, at least one recess (5) with a front end (5a) open and a back wall (5b), with the packing means, comprising an elastic and flexible packing body (3), being located in the front end (5a), and the back wall (5b) containing a set of spring elements (6, 7) among which are to be found coil springs or elastic strips, at least partially compressed by the packing body (3) and which push the latter against the surface of the second piece (1b) of the model.

Aircraft directional control and stabilizing surface

Stabilizing and directional-control surface of an aircraft, said surface comprising a vertical stabilizer (2) and a rudder (3), it being possible for said rudder (3) to be deflected relative to the vertical stabilizer (2), and moreover the rudder (3) comprises an internal profile (10) that can be extended and retracted by means of an actuating system (40) relative to the rest of the structure of the rudder (3), so that the stabilizing and control surface, in the retracted position of the internal profile (10) of the rudder (3), is an excellent aerodynamic surface in normal flying conditions, and at the same time increase of the aerodynamic control surface of the vertical stabilizer (2) is achieved for requirements of controllability of the aircraft at low speeds of said aircraft and against strong yawing moments acting thereon, in the position in which the internal profile (10) of the rudder (3) is extended, and moreover the structure of the rudder (3) can be opened, to permit extension of the internal profile (10), and said structure closes once the internal profile (10) is fully extended, so that the rudder (3) forms, with the extended internal profile (10), a structure that preserves its properties of aerodynamic surface and continuity of flow.

Front trimming fitting and related assembly

This relates to the assembly between a front trimming fitting and the tension joint of the two lateral boxes of the horizontal stabiliser of an aeroplane that comprises a front trimming fitting (1, 1a and 1b), a front shear plate (2), a top shear plate (4), a bottom shear plate (5), and a "rib post" joining part (3), such that it minimises the loads transferred to the lateral boxes, while enabling the horizontal stabiliser to be properly trimmed. Furthermore, it describes a compact front trimming fitting, made from a single part and of composite material.

USEFUL HYBRID FOR THE CURING OF COMPOSITE MATERIAL PARTS.

Util (1) híbrido para el curado de piezas de material compuesto que comprende una estructura de soporte (3) realizada en un material metálico de alta difusividad térmica, una placa base (31) de idéntica geometría superficial a la de la pieza a curar, realizada en un material metálico termoestable, una estructura intermedia (21) entre la estructura de soporte (3) y la placa base (31), de configuración celular, con las caras de cada celda suficientemente aligeradas como para facilitar el flujo de aire bajo la placa base (31) durante el ciclo de curado, realizada en un material metálico termoestable y unos medios de fijación de la estructura intermedia (21) a la estructura de soporte (3) que permitan su libre dilatación. Useful (1) hybrid for curing composite parts comprising a support structure (3) made of a metallic material of high thermal diffusivity, a base plate (31) of identical surface geometry to that of the piece to be cured, made of a thermostable metal material, an intermediate structure (21) between the support structure (3) and the base plate (31), of cellular configuration, with the faces of each cell sufficiently lightened to facilitate the flow of air under the base plate (31) during the curing cycle, made of a thermoset metal material and fixing means of the intermediate structure (21) to the support structure (3) that allow its free expansion.

Protection for elements made from composite material

对于由复合材料制成的元件(1)的保护，所述由复合材料制成的元件(1)在其复杂的几何结构之处例如边界、隆起、边缘和拐角被保护，以免受到这些元件可能暴露到的放电的影响，所述元件(1)是通过如下方法获得的：对安置在模具(2)上的复合材料施加温度和压力循环，所述模具(2)被设计为使得一旦知道所述元件(1)的几何结构，所述模具(2)就包括打算用于积聚在所述元件(1)的制造工艺中的树脂的空腔(6)，从而树脂将形成用于元件(1)的边界、隆起、边缘或拐角的介电保护，所述树脂积聚并形成保持在元件(1)的自由边界周围之处的层(3)，因此对这样元件的边界、隆起、边缘或拐角赋予电绝缘的介电保护，所述介电保护在对所述元件(1)的放电的情况下防止可能的放电和热粒子的喷出。本发明还涉及制造这种复合材料元件(1)的方法。

Process for the manufacturing of parts made of composite materials with two curing cycles

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 T 1 , comprised between the resin gelling temperature GT and the resin curing temperature CT, applied for a predetermined time PT 1 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 T 2 comprised between 90% and 100% of the resin curing temperature CT, applied for a predetermined time PT 2.

Hybrid tooling for the curing of composite pieces

本发明涉及一种用于固化复合工件的混合式工具(1)，包括：由具有高热扩散性的金属材料制成的支撑结构(3)；基板(31)，所述基板与要固化的工件具有相同表面几何形状，由INVAR 36或NILO 36制成；位于支撑结构(3)和基板(31)之间的中间结构(21)，该中间结构具有单元结构，每个单元的各个面都足够轻，以在固化周期期间使基板(31)下部的空气易于流动；以及把中间结构(21)固定到支撑结构(3)上而允许所述中间结构和所述支撑结构自由膨胀的装置。

STRUCTURAL CONFIGURATION FOR THE ASSEMBLY OF AN ENGINE ON THE STRUCTURE OF AN AIRCRAFT.

Configuración estructural para el montaje de un motor sobre la estructura de una aeronave. Structural configuration for mounting a engine on the structure of an aircraft. Configuración para el montaje de un sistema de propulsión de un motor 1 sobre el fuselaje 12 de una aeronave mediante una estructura de soporte 4 externa que conecta el motor 1 con el fuselaje 12, comprendiendo la citada estructura de soporte 4 una pluralidad de largueros 8 que se introducen en el fuselaje 12 y medios 7, 9 y 11 para unir dichos largueros 8 a la estructura interna 10 del fuselaje 12, pudiendo ser dichos medios 7, 9 y 11 amortiguadores, amortiguadores de choque o actuadores activos cuyas características se escogen de manera que los citados medios 7, 9 y 11 estén adaptados para la reducción eficiente de las cargas dinámicas internas y del acoplamiento, y para el amortiguamiento de las vibraciones continuadas en el conjunto sistema propulsor-aeronave. Configuration for mounting a system propulsion of an engine 1 on the fuselage 12 of an aircraft by means of an external support structure 4 that connects the motor 1 with the fuselage 12, said support structure 4 comprising a plurality of stringers 8 that are introduced into the fuselage 12 and means 7, 9 and 11 for joining said stringers 8 to the structure internal 10 of the fuselage 12, said means 7, 9 and 11 being able to be shock absorbers, shock absorbers or active actuators whose characteristics are chosen so that the mentioned means 7, 9 and 11 are adapted for efficient load reduction internal and coupling dynamics, and for damping the vibrations continued in the whole system propeller-aircraft.

PROCEDURE FOR MANUFACTURING OF PREPARED PARTS IN COMPOSITE MATERIAL WITH RIGIDIZERS APPLIED IN FRESH STATE.

Procedimiento de fabricación de piezas precuradas en material compuesto con rigidizadores aplicados en estado fresco, que comprende laminar capas superpuestas de material compuesto preimpregnado para obtener una pieza base (1) y una segunda pieza (2) destinada a unirse con ella; curar la pieza base (1); conformar en caliente el laminado destinado a formar la segunda pieza (2); retirar la segunda pieza (2) del útil de conformado de la misma y depositarla sobre la pieza base (1), con intercalación de una lámina de adhesivo estructural entre ambas; encerrar el conjunto de ambas piezas dentro de una bolsa de vacío; y cargar dicha bolsa en un autoclave, realizando en éste un ciclo de curado de la segunda pieza (2) bajo presión y temperatura, con lo cual ésta queda fuertemente adherida a la pieza base (1). La invención es aplicable al campo de la aeronáutica. Process for manufacturing pre-cured pieces in composite material with stiffeners applied in the fresh state, which comprises laminating superimposed layers of pre-impregnated composite material to obtain a base piece (1) and a second piece (2) intended to join with it; cure the base piece (1); hot forming the laminate intended to form the second piece (2); remove the second piece (2) of the forming tool thereof and deposit it on the base piece (1), with interleaving a sheet of structural adhesive between them; enclose the set of both pieces in a vacuum bag; and loading said bag into an autoclave, carrying out a curing cycle of the second part (2) under pressure and temperature, whereby it is strongly adhered to the base part (1). The invention is applicable to the field of aeronautics.

Pressure bulkhead made of composite material for an aircraft

The present invention relates to a curved rear pressure bulkhead ( 1 ) made of a composite material, comprising three layers: an outer layer ( 2 ) and an inner layer ( 3 ) symmetrical with one another, which are made of a fiber laminate, and the intermediate layer or core ( 4 ) made of a lightweight material. With such a sandwich-type shape, the resulting bulkhead ( 1 ) is able to withstand the bending loads coming from the strain and pressurization of the fuselage without having to increase its stiffness by means of using some type of stiffener. The bulkhead ( 1 ) comprises a ring ( 6 ) that is attached to the ends of the sandwich shape. This ring ( 6 ) is split into several parts ( 8 ) to facilitate the assembly of the bulkhead ( 1 ) made of composite material to the fuselage skin ( 5 ).

System for incorporating half-sections of aerofoils in aircraft

The aerofoil consists of a torsion box, capable of withstanding aerodynamic and inertial loads, and the half-sections (1, 1') which form it are connected together by joining means which conventionally consist of a rib forming one piece in its plane of symmetry, the final assembly of these aerofoils requiring access holes or openings for allowing access to the corresponding joints. The invention provides further joining means which avoid this rib and which consist of a plurality of bars (2) arranged in modular fashion in the manner of latticework, so that the required number of said access holes is reduced and sequential assembly is facilitated by connecting the top and bottom ends of the bars (2).

IMPACT RESISTANT AIRCRAFT FUSELAGE.

Fuselaje de aeronave resistente a impactos. Aircraft fuselage resistant to impacts. Configuración estructural del fuselaje trasero (4) de una aeronave con motores de hélice (1) que comprenden hélices (23) formadas a su vez por palas (3), estando los citados motores de hélice (1) ubicados en la parte trasera de la aeronave y estando a su vez el empenaje (5) de la aeronave situado detrás del plano de las hélices (23), comprendiendo la configuración estructural del fuselaje trasero (4) un revestimiento exterior (6) y un revestimiento interior (7), estando ambos revestimientos (6, 7) unidos por medio de unos elementos radiales (13) que configuran unas celdillas (14), de tal forma que la configuración estructural obtenida maximiza la resistencia a torsión del fuselaje trasero (4) de la aeronave en caso de daño del citado fuselaje trasero (4) por desprendimiento de una de las palas (3) de los motores de hélice (1). Structural configuration of the rear fuselage (4) of an aircraft with propeller engines (1) comprising propellers (23) formed in turn by blades (3), said engines being propeller (1) located at the rear of the aircraft and being at turn the instep (5) of the aircraft located behind the plane of the propellers (23), comprising the structural configuration of the rear fuselage (4) an outer shell (6) and a inner lining (7), both linings being (6, 7) joined by means of radial elements (13) that configure cells (14), so that the structural configuration obtained maximizes the torsion resistance of the rear fuselage (4) of the aircraft in case of damage of the aforementioned rear fuselage (4) by detachment of one of the blades (3) from the propeller motors (one).

REINFORCED COMPOSITE MATERIAL AIRCRAFT ATTACK EDGE.

Borde de ataque de aeronave de material compuesto reforzado. Material aircraft attack edge reinforced compound. La invención se refiere a un borde ataque (1) para aeronave fabricado en material compuesto caracterizado porque comprende en su cara interior un refuerzo (2) de tipo metálico adherido firmemente a la citada cara interior del borde de ataque (1) dispuesto de tal forma que confiere al borde de ataque (1) una mejor capacidad de adaptación a la deformabilidad y una mayor absorción de energía frente a impactos. Asimismo, la invención se refiere a un método de fabricación de un borde de ataque (1) tal con un refuerzo metálico (2) para aeronave realizado en material compuesto. The invention relates to a leading edge (1) for aircraft made of composite material characterized in that comprises on its inner face a reinforcement (2) of metallic type firmly adhered to the said inner face of the leading edge (1) arranged in such a way that it gives the leading edge (1) a better adaptability to deformability and greater energy absorption against impacts. Also, the invention is refers to a method of manufacturing a leading edge (1) such with a metal reinforcement (2) for aircraft made of material compound.

Sealing device for a flap track in a belly fairing of an aircraft for a flap-actuating shaft

Sealing device for the flap track for the flap-actuating shaft, in the belly fairing of an aircraft, with a top hollow body 1 and a bottom hollow body 7 which are flexibly compressible, each comprising at least one interior chamber 2, 8 with open sides 2 a, 2 b , a wall 3, 9 with respective external contact surfaces 3 a which are in flexible contact with each other, the outer surfaces 4 a, 10 a of the front walls 4, 10 of the hollow bodies 1, 7 jointly forming an outer sealing surface 11 which covers the flap track 19 , said hollow bodies 1,7 having jointly a vertical compression flexibility such that the contact surfaces 3 a, 9 a retract around the actuating shaft 20 , the interior chambers 2, 8 extending between the front walls 4, 10 and rear walls 5, 11 of the hollow bodies 2, 7.

Aircraft stabilizer surface trailing edge

Trailing edge ( 3 ) of an aircraft stabilizer surface ( 1 ), where this surface ( 1 ) is manufactured of a composite material and comprises an outer cladding ( 40 ) and an inner cladding ( 41 ) that are connected by a connecting clip type element ( 20 ) on this trailing edge ( 3 ), the connecting clip type element ( 20 ) comprising at its ends some recesses ( 23 ) used for coupling to the inner zone of the upper and lower claddings ( 40, 41 ) of the stabilizer surface ( 1 ), such that the connecting clip type element ( 20 ) is flexible enough to be pinched so that its ends will be housed, by means of these recesses ( 23 ), between the outer and inner claddings ( 40, 41 ) of the aircraft stabilizer surface ( 1 ), whereby the outer zone of the stabilizer surface ( 1 ), on its trailing edge ( 3 ), is formed by a continuous aerodynamic surface without changes of gradient.

Tubular components for aeronautical fuselages and processes and jigs for its manufacturing

The present invention relates to a tubular-shaped component ( 19 ) for an aeronautical fuselage made of composite materials comprising an outer skin ( 1 ) without joints having a cylindrical or cylindrical-conical shape, a plurality of stiffeners or stringers ( 2 ) longitudinally arranged in said outer skin ( 1 ) and a plurality of anchoring elements ( 3 ) for anchoring other elements to subsequently be incorporated transversally arranged between said stiffeners or stringers ( 2 ). The invention also comprises a jig ( 18 ) comprising a support body ( 12 ) and a plurality of panels ( 9 ) that can shift between an extended position ( 10 ) and a retracted position ( 10 ′) in both directions and a process for manufacturing said component ( 19 ) in the last step of which once the tubular component ( 19 ) is formed and cured, it is separated from the jig ( 18 ) by means of a longitudinal movement after the shifting of those elements of the jig ( 18 ) preventing it, and particularly the panels ( 9 ) that can be shifted.

Method and tool for manufacturing composite annular frame

本发明公开了一种用于制造用于航空机身的复合环形框架的方法，所述方法借助将树脂传递模塑技术应用于具有C型和L型截面的两个预成型件实现，采用两个工具(21，55)以如下步骤制造所述预成型件：制备材料；对矩形平面层板(41)进行热成形；对在第一工具(21)的直角截面的一部分上的直角截面(51)的层板进行热成形，放置弹性膜(55)，并采用温度和真空循环处理；通过使在第二弯曲工具(55)上的直角截面(51)的所述层板进行变形并采用温度和真空循环处理，将所述预成型件在所述第二弯曲工具(55)上热成形为C型(11)和L型(13)。本发明还涉及所述工具(21，55)。

Part of composite material having a wedge between two zones

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 ).

Sealing system for the gap existing between the fuselage and the elevator of the orientable horizontal stabiliser of an aircraft, extended with an aerodynamic fairing for sealing of the opening existing between the fuselage and the orientable horizontal stabiliser

Sealing system for the gap (2) between the fuselage (3) and the elevator (4) of the orientable horizontal stabiliser (5) of an aircraft, extended with an aerodynamic fairing (8) for sealing of the opening (24) between the fuselage (3) and said orientable horizontal stabiliser (5) which comprises a main body (1) and an aerodynamic fairing (8) which is extended from the main body (1) in continuity with the latter, and a plurality of first elastic sealing profiles (9) between the first surface (11) of the main body (1) including the aerodynamic fairing (8) and the outer surface of the fuselage (3) and making contact with them, and a plurality of second elastic sealing profiles (13) between the second surface (12) of the main body (1) and the first end of the elevator (15) and making contact with them, in such a way that the sealing takes place of the gap (2) and the opening (24) and an aerodynamic continuity is produced between the orientable horizontal stabiliser (5), the fuselage (3) and the elevator (4) when the elevator (4) is at rest in the plane of the orientable horizontal stabiliser (5) for any of the orientation positions of the orientable horizontal stabiliser (5).

cleaning device for a pre-impregnated carbon fiber wick positioning machine

DISPOSITIVO DE LIMPEZA PARA UMA MáQUINA POSICIONADORA DE MECHAS DE FIBRAS DE CARBONO PREVIAMENTE IMPREGNADAS. A presente invenção refere-se a um dispositivo de limpeza (5) para ser usado em uma máquina posicionadora de mechas de fibras de carbono previamente impregnadas, o qual inclui pelo menos um cilindro guia (47) de um feixe (21) de mechas até um cabeçote de aplicação (25), que compreende dois elementos filtrantes (11, 13) dispostos de maneira que entre eles é circulada o feixe (21) de mechas que são suportadas por duas placas (15, 17) unidas com meios reguláveis de retenção (29) dos ditos elementos filtrantes (11, 13) e meios de sujeição/separação para /da dita máquina. A invenção também refere-se a uma máquina posicionadora de mechas de fibras de carbono previamente impregnadas com pelo menos um dispositivo de limpeza (5). CLEANING DEVICE FOR A CARBON FIBER MACHINERY POSITIONING MACHINE PREVIOUSLY IMPREGNATED. The present invention relates to a cleaning device (5) to be used in a machine for positioning previously impregnated carbon fiber wicks, which includes at least one guide cylinder (47) of a wick bundle (21) up to an application head (25), comprising two filter elements (11, 13) arranged in such a way that the bundle (21) of wicks is circulated between them and are supported by two plates (15, 17) joined with adjustable retention means (29) of said filter elements (11, 13) and holding / separating means for / from said machine. The invention also relates to a machine for positioning carbon fiber wicks previously impregnated with at least one cleaning device (5).

Computer-Aided Method for a Cost-Optimized Calculation of Aerodynamic Forces on an Aircraft

A computer-aided method suitable for assisting in the design of an aircraft by providing the values of dimensional variables, dependant of a predefined set of parameters, for the complete aircraft or an aircraft component, comprising the following steps: a) Defining a parametric space grid; b) Obtaining a suitable Reduced Order Model (ROM) model, particularly a Proper Orthogonal Decomposition (POD) model, for calculating said variables for whatever point over the parametric space through an iterative process. Computer Fluid Dynamics (CFD) is used to calculate said variables for an appropriately selected set of points over the parametric space, which are used to approximate, via ROM and ad hoc interpolation, the variables in any other point over the parametric space. The method minimizes the required number of CFD calculations (to minimize the computational cost, which dramatically depends on this number) for a given level of error.

Aircraft tail cone with folding movable fairing

The invention comprises a skin provided with a movable fairing (1) which is supported on the skin by means of a swing hinge mechanism (7), so as to allow the opening/closing of the movable fairing (1) and access to the tail cone interior. The invention comprises actuators (9) for actuating the swing hinge mechanism (7) to allow the automatic opening and closing of the movable fairing (1).

System for integrating rivet nuts into plastic components

The system is intended for fixing metal rivet nuts (1) to plastic components (2), achieving the fixing by means of ultrasound riveting or any other mechanical method. Each plastic component (1), made through the relevant manufacturing process, preferably by injection moulding, has a number of nipples (4) that are placed in the rivet nut holes (1). Riveting is subsequently carried out by ultrasound or other mechanical method on these nipples (4), fusing them and their corresponding union with the rivet nut (1) and in short, fixing the nut onto and integrating it into the plastic component (2). The riveting system referred to for integrating rivet nuts into plastic components creates a union between the rivet nuts and other components that are part of a structure. These other components can be metallic or non-metallic.

Automatic dismoulding system

一种用于处在固化工具(12)中的模压模件(7)的自动脱模系统，包括：一个可编程的机械手(1)包括一端自由的铰接臂(1a)，能使该铰接臂(1a)具有水平旋转运动和垂直倾斜运动的致动装置(13、14)，以及一个位于铰接臂(1a)的自由端的头部(2)，它具有至少一个能使之粘附到目的物的表面(16、17)的真空吸附垫(11)；为产生真空的真空系统(15)；用于定位每个目的物(6、7)的垂直和水平位置的定位装置(3)；以及可编程的控制装置(18)，该装置控制所述致动装置(13、14)，因此铰接臂(1a)可以将吸附垫(11)定位到处在初始位置的目的物(6、7)的表面(16、17)上并将它转送到最终位置。

A system for transmitting pressures between a detachable part and a fixed part of a scale model placed on aerodynamic tunnel

System for transmitting pressures between a detachable part (2a,2b) and a fixed part (1) of a model in a wind tunnel, which presents pressure tapping holes (3a-3h) and pressure outlets (5a-5h) in the detachable part connected together by means of internal conduits (4a-4h), and pressure inlets (7a-7h) in the fixed part (1) connected to a pressure measuring module via pressure transmission ducts (8a-8h) and pressure transmission conduits (9a-9h), in such a way that when the two pieces are coupled by coupling means (6), the pressure outlets (5a-5h) are facing the pressure inlets (7a-7h), with a sealing joint 10 existing between them with passage holes (10a-10h), the entire unit permitting the transmission of pressure from the detachable part (2a,2b) to the fixed part (1) without any need for pipes or connectors.

AIRCRAFT NOTEBOOK DESIGN

Cuaderna (1) de refuerzo del fuselaje (2) de una aeronave, comprendiendo dicho fuselaje (2) un revestimiento (3) y unos larguerillos (20), comprendiendo la citada cuaderna (1) un elemento estructural (4) y un elemento alveolar (5) cerrado, hueco en su interior, comprendiendo el elemento estructural (4) una base inferior (6) que va dispuesta sobre la cara interior del revestimiento (3) del fuselaje (2), y a través de la cual pasan los larguerillos (20) que unen las cuadernas (1) al resto de la estructura del fuselaje (2), y unos elementos de refuerzo superiores (7) sobre los que van dispuestos los citados elementos alveolares (5), tal que los elementos alveolares (5) aumentan la inercia y la rigidez de la cuaderna (1) sin añadir peso a la misma, teniendo la cuaderna (1) elevada rigidez en dirección transversal, además de ser estable intrínsecamente a pandeo local.

Method for optimising the structural design of a composite stiffened panel

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.

Process and jig for manufacturing composite material structures

The invention relates to a process for manufacturing stiffened structures ( 1 ) in composite materials formed by an outer skin ( 2 ) and a plurality of stiffeners ( 3 ) the cross-section of which has a closed shape delimiting an inner opening ( 6 ), comprising the following steps: a) providing a shaping jig ( 34 ); b) providing auxiliary male jigs ( 36 ) with inner cavities ( 32 ) reinforced ( 33 ) along their length, coated by membranes ( 37 ) suitable for curing; c) providing stiffeners ( 3 ); d) arranging the stiffeners ( 3 ) in the jig ( 34 ) and in their inner openings ( 6 ) the auxiliary male jigs ( 36 ); e) laminating the outer skin ( 2 ); f) curing the stiffened structure ( 1 ) with high temperature and pressure; g) removing the auxiliary male jigs ( 36 ) after reducing their section ( 41 ); h) separating the cured stiffened structure ( 1 ) from the jig ( 34 ), the design of the inner cavities ( 32 ) including their reinforcement ( 33 ) and the capacity to maintain or reduce their section by the application of overpressure or underpressure respectively.

WIND TUNNEL WITH A CLOSED SECTION FOR AEROACUSTIC MEASUREMENTS WITH AN ANECOIC COATING.

Túnel de viento con una sección cerrada para mediciones aeroacústicas con un revestimiento anecoico. Wind tunnel with a closed section for aeroacoustic measurements with an anechoic coating. Túnel de viento con una sección cerrada para mediciones aeroacústicas con un revestimiento anecoico con la siguiente estructura: a) una primera cavidad (11) de espesor D _{1} en el rango 20 \leq D _{1} \leq 50 mm rellena de un material fibroso con una resistividad al flujo R en el rango 10 \leq R \leq 50 krayl/m; b) un primer panel microperforado (13) de espesor t_{1} en el rango 0.25 \leq t_{1} \leq 0.75 mm y con un porcentaje de perforación p _{1} en el rango 15% \leq p _{1} \leq 30%; c) una segunda cavidad (15) de aire de espesor D _{2} en el rango 10 \leq D _{2} \leq 30 mm; d) un segundo panel microperforado (17) de espesor t_{2} en el rango 0.25 \leq t_{2} \leq 0.75 mm y con un porcentaje de perforación p _{2} en el rango 15% \leq p _{2} \leq 30%. Las perforaciones de los paneles pueden tener forma circular o de hendidura longitudinal de diámetro d o anchura a en el rango 0.2 \leq d, a \leq 0.5 mm. Wind tunnel with a closed section for aeroacoustic measurements with an anechoic coating with the following structure: a) a first cavity (11) of thickness D 1 in the range 20 ≤ D 1 ≤ 50 mm filled of a fibrous material with a flow resistance R in the range 10 ≤ R ≤ 50 krayl / m; b) a first microperforated panel (13) of thickness t_ {1} in the range 0.25 ≤ t_ {1} ≤ 0.75 mm and with a perforation percentage p 1 in the range 15% ≤ p _ { 1}? 30%; c) a second air cavity (15) of thickness D 2 in the range 10 ≤ D 2 ≤ 30 mm; d) a second perforated panel (17) tR thickness {2} in the range 0.25 \ leq t_ {2} \ leq 0.75 mm and with a perforation percentage p _ {2} in the range 15% \? p _ { 2}? 30%. The perforations of the panels may have a circular or longitudinal groove with a diameter of width a in the range 0.2 d d, a ≤ 0.5 mm.

RIB-HARDWARE.

Costilla-herraje como componente (9) de un cajón de torsión (1, 3) de una superficie sustentadora de una aeronave que está realizado en material compuesto como una pieza unitaria y cuya configuración comprende: un alma (31) sustancialmente plano con una primera parte en forma de orejeta (33) y una segunda parte en forma de alma de costilla (35); dos alas (39) para unir el alma (31) con las almas de cada uno de los extremos del larguero (11, 13) del cajón al que se conecta; alas (43) de unión a los revestimientos superior e inferior (19, 21) del cajón; vanos (45) en las zonas de intersección con cordones del larguero (11, 13) al que se conecta el componente (9) y con los larguerillos de refuerzo (25) de los revestimientos (19, 21). La invención también comprende procedimientos de montaje de la costilla-herraje en cajones de torsión multilarguero. Rib-ironwork as a component (9) of a torsion box (1, 3) of a supporting surface of an aircraft that is made of composite material as a unitary piece and whose configuration comprises: a substantially flat (31) core with a first lug-shaped part (33) and a second rib-shaped part (35); two wings (39) to join the soul (31) with the souls of each of the ends of the crossbar (11, 13) of the drawer to which it is connected; wings (43) joining the upper and lower liners (19, 21) of the drawer; spans (45) in the intersection zones with stringers of the stringer (11, 13) to which the component (9) is connected and with the reinforcing stringers (25) of the linings (19, 21). The invention also comprises methods of mounting the rib-ironwork in multi-beam torsion drawers.

Aircraft loading frame by Composite Preparation

本发明涉及由复合材料制造的飞行器负载框架(1)，其特点是，它包括两个侧元件(2，3)和基底元件(7)，通过基底元件(7)将侧元件(2，3)连接在框架(1)的内部，每个侧元件(2，3)包括将框架(1)连接至飞行器机身外壳的腿(4)、连接板(5)以及连接所述连接板(5)和基底元件(7)的下凸缘(6)。

System for compensating aerodynamic measurements in a wind tunnel

Sistema para compensar mediciones aerodinámicas en un túnel de viento. Sistema para compensar mediciones aerodinámicas en un túnel de viento realizados con una disposición de ensayo que comprende un estilete principal 1 conectado a la parte trasera de una maqueta que emula al menos el empenaje (1a) y un ala 1b de un avión, cuyo sistema comprende un estilete gemelo simulado 3 que se extiende coaxialmente al estilete principal 2, con un extremo libre 3a enfrentado en configuración de contorno aerodinámico a un área de medición 1c en la parte trasera de dicho ala 1b; un intersticio 4 entre dicho extremo libre 3a del estilete gemelo simulado 3 y dicho área de medición (1c); y medios sensores 5 en dicho extremo libre 3a del estilete gemelo simulado 3 que detectan desplazamientos entre dicho área de medición 1c y dicho extremo libre 3a cuando la maqueta 1 se somete a una corriente de aire.

The composite structure and the manufacture method thereof that are used for aircraft fuselage

本发明提供用于在阳夹具上成形的飞行器机身的封闭复合材料结构，所述封闭复合材料结构可沿特定方向从阳夹具分离，所述结构包括单个外面板和集成在所述面板中的多个内纵向支肋，使得所述阳夹具的膨胀系数大于所述结构的复合材料的膨胀系数，从而能够在单个操作中移除由面板和集成的内支肋形成的已经制造的结构。本发明进一步提供用于制造这样的封闭结构的过程。

Tail cone for aircraft with movable fairing and support structure for auxiliary power unit and the ancillary elements thereof

Tail cone (2) for aircraft (1) with a cover (8) housing an auxiliary power unit (3) and the ancillary systems thereof (6,7), and the cover (8) has a fixed forward section (11) by means of which the tail cone (2) is secured to the rest of the fuselage (4), and a movable fairing (9), housing a support structure (10) secured to the fixed front section (11) which supports the auxiliary power unit (3) and the ancillary systems thereof (6,7). The movable fairing (9) moves by means of folding processes or rather longitudinally by means of telescopic rails (16) secured to the fixed front section (11), from a closed position in which the cover (8) remains closed housing on the inside thereof the support structure (10) with the auxiliary power unit (3) and the ancillary elements thereof (6, 7), and an open position in which the cover (8) remains open leaving the support structure (10) with the auxiliary power unit (3) and the ancillary elements thereof (6, 7) exposed, affording the possibility of access thereto for purposes of maintenance, repair or replacement.

Reinforced cover for gaps in an aerodynamic contour

Reinforced cover for gaps in an aerodynamic contour of a vehicle, with a first attachment section 1 a attachable to a structural element 2, 5 of the vehicle 11 , on a first axial plane I; a second elastic section 1 c 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 1 e , an outer surface 1 k with a fiberglass layer 7 a ; a main internal body 9 of an elastic material, a flexible transition section 1 g between the first and the second section, 1 a , 1 c ; the outer surface 1 k of the second section 1 c flushing with the outer surface 5 a and extending towards the free end 1 d as a continuation of the outer surface 5 a of the fixed part 5.

Forming of double curvature laminated composite sheets comprises uniaxial bending of one layer, for controlled deformation via a tool and plates

The forming of double-curvature laminated composite sheets has uniaxial bending of one layer of the laminate on a tool leaving a border which slides between two forming plates which then fold the bent layer via the tool. The upper plate is spaced relative to the vertical surface of the tool to the extent of the size of the bent layer.

CONTROLLABILITY INCREASE SYSTEM FOR AN AIRCRAFT.

Sistema de aumento de controlabilidad para una aeronave.#El sistema incorpora medios (16) de transmisión de deflexión para proporcionar una posición de deflexión de un dispositivo (8) hipersustentador móvil, instalado en el HTP (2) móvil de una aeronave (1), para cada posición del ángulo de calado del HTP (2).#Preferiblemente, los medios (16) de transmisión de deflexión (18, 20, 23) consisten esencialmente en una ligadura mecánica conectada entre el dispositivo (8) hipersustentador y la aeronave (1). Increase of controllability system for an aircraft. # The system incorporates deflection transmission means (16) to provide a deflection position of a mobile hypersustainer device (8), installed in the mobile HTP (2) of an aircraft (1) , for each position of the draft angle of the HTP (2). # Preferably, the deflection transmission means (16) (18, 20, 23) consist essentially of a mechanical ligature connected between the hyper-sustaining device (8) and the aircraft (one).

Tool and method for producing wide parts of composite material

本发明公开了一种通过铺放、切割和热成型操作制造多片复合材料的设备，所述设备包括：两个相同的工作台(11，13)，所述工作台(11，13)构造有平坦上表面(41)或构造有将被制造的片材形状的上表面(43)，所述工作台(11，13)还包括一装置，用于在所述装置的封闭空间内产生真空；自动铺放头(15)，所述自动铺放头(15)由允许所述自动铺放头(15)在所述工作台(11，13)上移动的装置支撑，且切割头(17)连接到所述自动铺放头(15)；以及工具(21)，所述工具(21)也可以在所述工作台(11，13)上移动，所述工具(21)包括成型薄膜(25)和加热装置(27)。本发明还包括通过在同一设备上实施的铺放、切割和热成型操作制造复合材料片材的方法。

Opening and secure-closing system for doors

Opening and secure-closing system for doors comprising two doors (1,2) for accessing the interior of a chamber, which open towards the outside being hinged around two hinging axes (19,20), both doors (1,2) comprising a set of bolts (17,18) and the opening and closing system comprising additional means of opening/closing operation (26) and an overpressure door (3) located in one of the doors (1,2) which opens towards the outside, in such a way that the means of opening/closing operation (26) comprise a blocking bar (4), a blocking fitting (7) and means of retention (10) between the two doors (1,2) in order to provide a controlled and sequential opening/closing operation of the chamber, and the overpressure door (3) is opened automatically in the event of emergency due to overpressure, and it also participates in the controlled and sequential opening/closing operation of the chamber.

Monolithic composite panel with multicell box shape

This invention relates to monolithic composite panel, the structure of which comprises a skin (1), a plurality of omega-shaped stiffeners (3), an also omega-shaped element (5) covering said stiffeners (3).

Lightning strike protection system for aircraft fuel tanks made of low electrical conductivity composite material

A system of structural protection against electrical discharges comprises electrical conductive thin wire mesh (1) covering external surface of tank outer skin made of low electrical conductivity composite; and thick wire metallic mesh (2) overlapping mesh at minimum distance of 50 mm to both sides of row of fasteners joining outer skin to internal part of composite or metallic material. Both metallic meshes maintain electrical contact by their installation/assembly and by means of metallic countersunk head washers (3) connected to bonding points and set to the gap existing between the fastener and the outer skin. The system also includes an isolating ply (F.V.) of fibre-glass material or any other isolating material if the internal part is metallic. An independent claim is included for manufacturing the structural protection system.

Metallic local reinforcement for heavy loaded joints of composite components

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 fibres 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.

Aircraft having a lambda-box wing configuration

Aircraft having a lambda-box wing configuration, comprising a fuselage ( 1 ), a propulsion system ( 5 ), a first pair of swept-back airfoils ( 2 ), connected to the top forward portion of the fuselage ( 1 ), a second pair of swept-forward airfoils ( 3 ), connected to the lower rear portion of the fuselage ( 1 ) at a point of the said fuselage ( 1 ) aft of the connection of the swept-back airfoils ( 2 ), and a third pair of substantially vertical airfoils ( 4 ), the tips of the swept-forward airfoils ( 3 ) being connected to the lower side of the swept-back airfoils ( 2 ) at an intermediate point of the span of the said swept-back airfoils ( 2 ), by means of the substantially vertical airfoils ( 4 ), the swept-back airfoils ( 2 ) having a higher aspect ratio than that of the swept-forward airfoils ( 3 ), which makes the swept-back airfoils ( 2 ) have a reduced induced drag without penalizing their weight.

Process and tooling for lining molds of curved pieces

本发明涉及一种对用于弯曲件的模具的加衬的工艺和工具，该工艺包括：以连续几遍的形式施加宽度为A2的衬里(15)在模具(3)上，在其上施加压力以使其适合所述弯曲件的曲率，所述遍的次数是最小需求的次数，由此带(15)的宽度A2允许实现这种适应，工具包括施加衬里(15)的鼓(25)，在每一遍的开始时将它的末端连接到模具(3)上的装置，在衬里(15)上沿着它的整个宽度均匀施加压力的圆锥形辊(25)，和通过安置在模具(3)上的边侧引导件(27、28)而使其可以运动的部件(26、29)，所述模具(3)平行于划定所述弯曲件的曲线(5、7)。

ULTRASOUND HEAD FOR MULTICHANNEL INSPECTION BY PULSE-ECO.

Cabezal de ultrasonidos para inspección multicanal por pulso-eco aplicado a superficies planas o curvas, que se une a un sistema de desplazamiento (3), y que comprende un bastidor (1) en el que está dispuesto un marco basculante (4) unido a dicho bastidor a través de un primer punto de giro (1a) y un segundo punto de giro (1b) los cuales definen un eje principal de basculamiento (1c) alrededor del cual gira el marco basculante (4), presentando dicho marco basculante (4) en su abertura central (4h) un conjunto de porta-palpadores (2) que portan una pluralidad de palpadores (12) de ultrasonidos, pudiendo girar dicho conjunto de porta-palpadores (2) alrededor de al menos un eje de giro (5,5a,5b) transversal al eje principal de basculamiento (1c) en un movimiento independiente al que realiza el marco basculante (4), pudiendo estar dividido el marco basculante (4) en dos marcos secundarios (6,8) que giran de forma independiente al giro de marco basculante (4). Ultrasonic inspection head multichannel by pulse-echo applied to surfaces flat or curved, which joins a displacement system (3), and comprising a frame (1) in which a frame is arranged swingarm (4) attached to said frame through 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), said tilting frame (4) presenting in its central opening (4h) a set of probe holder (2) bearing a plurality of ultrasonic probes (12), said set of being able to rotate probe holder (2) around at least one axis of turn (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) can be divided into two frames secondary (6.8) that rotate independently of the frame rotation tilting (4).

Complex geometries made from composite material and method for forming said geometries

本发明涉及一种用于成型由复合材料形成的复杂几何叠层的成型方法，其可以获得非常复杂的弯折几何体，减少成型区域的影响范围并使得其影响局部化，在待成型区域附近处理叠层的几何形状，用弯折几何体代替成型几何体，由此迫使成型的影响区域局部化，以使得受所述成型影响的增强纤维的长度最小。本发明还涉及通过上述成型方法获得的由复合材料制造的复杂叠层几何体。

Tail cone for aircraft with movable fairing and support structure for auxiliary power unit and the ancillary elements thereof

Tail cone (2) for aircraft (1) with a cover (8) housing an auxiliary power unit (3) and the ancillary systems thereof (6,7), and the cover (8) has a fixed forward section (11) by means of which the tail cone (2) is secured to the rest of the fuselage (4), and a movable fairing (9), housing a support structure (10) secured to the fixed front section (11) which supports the auxiliary power unit (3) and the ancillary systems thereof (6,7). The movable fairing (9) moves by means of folding processes or rather longitudinally by means of telescopic rails (16) secured to the fixed front section (11), from a closed position in which the cover (8) remains closed housing on the inside thereof the support structure (10) with the auxiliary power unit (3) and the ancillary elements thereof (6,7), and an open position in which the cover (8) remains open leaving the support structure (10) with the auxiliary power unit (3) and the ancillary elements thereof (6,7) exposed, affording the possibility of access thereto for purposes.of maintenance, repair or replacement.

Hatch-actuating mechanism, particularly for an aircraft landing gear

A hatch-actuating mechanism, particularly for aircraft landing gear with a lateral closing panel ( 1 ) and a secondary closing panel ( 4 ) articulated to one another, a lateral pivoting fitting ( 5 ) joined at the bottom to the lateral panel ( 1 ), couplable by lateral means of articulation ( 6 ) to a fixed structure ( 7 ), a secondary pivoting fitting ( 8 ) joined at the bottom to the secondary panel ( 4 ), a lateral actuating cylinder ( 10 ) articulated on the fixed structure and on the lateral pivoting fitting ( 5 ), an intermediate actuating cylinder ( 13 ) coupled to a first coupling point ( 14 ) between the lateral means of articulation ( 6 ) and the intermediate means of articulation ( 9 ) on the lateral pivoting fitting ( 5 ), as well as on a second coupling point ( 15 ) on the secondary pivoting fitting ( 8 ), the intermediate actuating cylinder ( 13 ) being actuatable independently from the lateral actuating cylinder ( 10 ).

Be used to simulate the method for characteristic of the adhint of two parts

本发明涉及一种用于模拟具有粘合材料层(13)的两个复合材料部件(11，15)的粘合接头的特性的方法，包括下述步骤：a)提供具有用于分析结构接头的所有相关信息的有限元模型；b)提供用于在粘合材料层(13)经受拉伸/剥离和剪切应力时计算粘合材料层(13)的变形的计算模型，包括处于拉伸/剥离状态下的线性类型和处于剪切状态下的非线性类型的粘合剂的塑料特性；c)通过应用计算模型，模拟粘合接头的特性，以获得每个故障模式的故障指数。本发明还涉及一种用于采用所述结构接头的由计算机执行的有限元模型辅助设计所述结构接头的系统。

Method and tool for manufacturing composite annular frame

本发明公开了一种用于制造用于航空机身的复合环形框架的方法，所述方法借助将树脂传递模塑技术应用于具有C型和L型截面的两个预成型件实现，采用两个工具(21，55)以如下步骤制造所述预成型件：制备材料；对矩形平面层板(41)进行热成形；对在第一工具(21)的直角截面的一部分上的直角截面(51)的层板进行热成形，放置弹性膜，并采用温度和真空循环处理；通过使在第二弯曲工具(55)上的直角截面(51)的所述层板进行变形并采用温度和真空循环处理，将所述预成型件在所述第二弯曲工具(55)上热成形为C型(11)和L型(13)。本发明还涉及所述工具(21，55)。

Device to control the ventilation and air conditioning in an aircraft on the ground

New ecological operating mode that limits ventilation and air conditioning in the cabin (3) of the aircraft when the ecological operating mode has been selected, the aircraft is stationed on the ground, the engines are off, the engines' start up system has not been activated, the pneumatic or electrical energy supply system that typically controls ventilation and air conditioning in the aircraft has been activated and is supplying air, to provide a new method of operating at an established minimum limit independent of the difference between the selected temperature and the temperature measured, saving energy, and reducing pollutant emissions and noise.

Device for facilitating the lowering of a landing gear by gravity

Device for facilitating the lowering of a landing gear by gravity, comprising at least on roller device (4) positioned an at least one closing panel (2) of the hatch for the landing gear in a contact area with the tire (1a) of the landing gear of an aircraft (7) on initiating its descent, the roller device (4) comprising a rigid roller device which can rotate freely about its axis (4a), and being positioned such that its axis is perpendicular to the path of the tire (1a) so that the tire will make the roller device (4) turn without sliding over it.

Integrated aircraft structure in composite material

Integrated aircraft structure in composite material that comprises a skin and stringers (1), with the skin comprising a skin part (5) and a basic skin (3), and said structure further comprises U-shaped elements (15) each of which comprises in its turn two L-shaped sections (4a+5a and 4b+5b) together with the skin part (5), in such a way that these U-shaped elements (15) fulfil two structural functions in said structure at the same time, acting as frame feet and as skin, providing an integrated fuselage structure without rivets or joints. The invention also relates to a manufacturing process of an integrated aircraft structure in composite material.

SEALING PROFILE AND USE OF THE SAME.

Perfil de sellado y uso del mismo. Sealing profile and use thereof. Constituido en material deformable para permitir realizar la unión y sellado entre un primer (2) y un segundo elemento (3); y que tiene por objeto evitar la formación de pinzamientos del perfil de sellado al realizar el montaje y sellado entre el primer (2) y segundo elemento (3). Para ello el perfil de sellado (6 y 7) presenta una configuración en cuña que sobresale respecto del primer elemento (2), y sobre cuya superficie inclinada presiona el segundo elemento (3) para facilitar su deformación y el montaje y sellado del primer (2) y segundo elemento (3), posición en la que la superficie de la cuña que sobresale respecto del primer elemento (2) presiona contra el segundo elemento (3). Made of deformable material to allow perform the bonding and sealing between a first (2) and a second element (3); and which aims to prevent the formation of stamping of the sealing profile when assembling and sealing between the first (2) and second element (3). For this, the profile of sealed (6 and 7) features a wedge configuration that protrudes with respect to the first element (2), and on whose inclined surface press the second element (3) to facilitate its deformation and the assembly and sealing of the first (2) and second element (3), position in the surface of the wedge that protrudes from the first element (2) presses against the second element (3). La invención es preferentemente aplicable a estructuras aeronáuticas, y más concretamente en aquellas en las que el primer y segundo elemento forman parte de la superficie aerodinámica. The invention is preferably applicable to aeronautical structures, and more specifically in those in which the first and second elements are part of the surface aerodynamics.

Fitting for trimming a horizontal stabilizer of an aircraft

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

Tool and process for manufacturing long pieces of composite material

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.

Stringers made of a composite material with a bulb

The invention relates to a stringer ( 11, 21, 31 ) made of a composite material for stiffening panels made of a composite material, particularly those used in the aeronautics industry, formed by a foot ( 15 ) for joining with the panel and a structural member ( 13, 23, 33 ) with a structural reinforcement ( 17, 27, 37 ) embedded at the end opposite to the foot ( 15 ) formed by high-modulus unidirectional fibers of the same material as the stringer or another material compatible with it. The invention also relates to a process of manufacturing the stringer.

Sealing device for a flap track in a belly fairing of an aircraft for a flap-actuating shaft

Sealing device for the flap track for the flap-actuating shaft, in the belly fairing of an aircraft, with a top hollow body 1 and a bottom hollow body 7 which are flexibly compressible, each comprising at least one interior chamber 2, 8 with open sides 2a, 2b, a wall 3, 9 with respective external contact surfaces 3a which are in flexible contact with each other, the outer surfaces 4a, 10a of the front walls 4, 10 of the hollow bodies 1, 7 jointly forming an outer sealing surface 11 which covers the flap track 19, said hollow bodies 1,7 having jointly a vertical compression flexibility such that the contact surfaces 3a, 9a retract around the actuating shaft 20, the interior chambers 2, 8 extending between the front walls 4, 10 and rear walls 5, 11 of the hollow bodies 2, 7.

Dynamic verification method for a riveting process with blind rivets carried out with an automatic riveting apparatus, and verifying device for carrying out the verification

A method and device are provided for verification of the quality of placing of blind rivets. The verification method involves successively carrying out, during the application of traction force to a stem, initial measurements in an initial state of the application of the traction force to obtain an initial traction force and an initial displacement of the rivet stem relative to the rivet body, intermediate measurements in an intermediate stage of the application of the traction force to obtain an intermediate traction force and an intermediate displacement, and terminal measurements in a terminal stage of the application of the traction force to obtain a terminal traction force and a terminal displacement. The obtained values are compared to acceptable values in an assignment table.

Process for manufacturing composite material structures with collapsible tooling

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.

Method of protecting fuel tanks manufactured with composites against electrical discharges

A method of protecting a group of metal devices ( 13, 15 ) installed inside a fuel tank ( 11 ), manufactured either completely or partially with composites, such as a tank located in an aircraft, against electrical discharges, according to which the following steps are carried out: a) an insulating member ( 21 ) is included at the points where the metal devices ( 13, 15 ) are fixed/attached to the fuel tank ( 11 ) to assure their electrical insulation with respect to the fuel tank ( 11 ); b) insulating inserts ( 23 ) are included in the linear metal devices ( 13 ), such as pipes, such that they are subdivided into parts ( 31, 33, 35 ) insulated from one another; c) a connection with a metal sub-structure ( 25 ) with a very low resistance is provided in each one of the insulated parts ( 31, 33, 35 ).