Abstract: 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 profil

Abstract: 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.

Abstract: 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.

Abstract: 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 aircraft structure whic

Abstract: 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).

Abstract: 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).

Abstract: 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).

Abstract: 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.

Abstract: 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.

Abstract: 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.

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

Abstract: The 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.

Abstract: 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).

Abstract: 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 (5a) open and a back wall (5b), with the sealing means, comprising an elastic and flexible sealing 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 sealing body (3) and which push the latter against the surface of the second piece (1b) of the model.

Abstract: 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 (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.

Abstract: 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.

Abstract: 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.

Abstract: 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, 5a, 5b) immobilized on the reaction bar (4) and side arms (6, 7) which clasp the sides (1c, 1d) of the actuator (1), both having windows (8, 9) through which respective pins (10, 11) joined to these sides (1c, 1d) 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).

Abstract: 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.

Abstract: 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.