Abstract: Method of manufacturing “T” shaped stringers (1) with an angle different from 90° between the web (2) and the foot (3) whereby the “T” shaped stringers have a stringer web and a stringer foot. The method comprises, after placing together two hot-formed “L” shaped semi-stringers to form a “T” shaped stringer, placing the “T” shaped stringer inside an invar alloy angle leaving a gap between the stringer web and the invar alloy angle. Afterwards, a heating device (7) is moved over the stringer feet (3) surface and a roller (8) slides over said stringer feet (3) surface to adapt the geometry of the stringer feet to the geometry of the invar alloy angle. The resulting “T” shaped stringer is co-bonded on a cured skin with an adhesive line between the stringer and the cured skin, and finally the obtained “T” shaped stringer is cured.

Abstract: The method comprises the manufacture of a front section (2) and an rear section (3) with an assembly structure integrated into the rear section and the subsequent assembly of the two sections. The manufacture of the front section (2) comprises the manufacture of a main portion (9) of the front section separated from a coupling portion (10) of the front section. The assembly of the two sections comprises the assembly of the main portion (9) with the group comprising of the rear section assembly (3) and the coupling portion (10) of front section.

Abstract: Method of manufacture of one or more pairs of components in composite material (11, 13; 11?, 13?) intended to form part of aerofoils located on both sides of the axis of symmetry of an aircraft whose shape has a conical tendency in the longitudinal direction, the angle of conicity being less than 10°, that comprises the following stages: a) Lay-up of layers on a flat tape lay-up table forming the laminate of a hypothetical component formed by all of the components of said pairs (11, 13; 11?, 13?), the components of each pair being arranged adjacently and the first components (11, 11?) and the second components (13, 13?) of each pair being arranged alternately, and the whole being arranged symmetrically relative to a hypothetical horizontal axis X-X of the tape lay-up jig; b) Cutting of the flat laminates corresponding to each of said components); c) Forming and curing said components.

Abstract: It comprises applying a mould release agent on a compaction tool (1), carrying out its curing and applying a layer with adhesive properties (2) on the mould release agent. It is characterized in that it also comprises extending an aerator blanket (4) along the layer (2) with dimensions larger than the adhesive layer (2), extending a membrane with resilient properties (5) on the entire surface of the tool (1) by applying vacuum to the membrane (5) for compacting the adhesive layer. This method reduces the number of stages required to perform the compaction and avoids the generation of waste materials thus reducing the cost of the process.

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

Abstract: A fuselage section of an aircraft includes a skin, a plurality of frames positioned transversely to a longitudinal axis of a fuselage of the aircraft, and a plurality of longitudinal stringers at least one of which is configured with a closed transversal section including a stringer hat, two stringer webs, and two stringer feet joined to the skin. At least one of the frames is configured in at least one sector with a frame foot joined to the skin, a frame web having holes at crossing zones with the stringers, a frame cap, and a frame cap extension which does not interfere with the stringers. The frames are joined to, at least, the stringer hats at crossing zones of the frames and the stringer hats.

Abstract: The invention discloses a trailing edge (2) of an aircraft aerodynamic surface, which comprises an upper skin (6) and a lower skin (7), being the upper and lower skins (6, 7) made of composite material. In the trailing edge (2) a lightweight part (1) is assembled at an end section (13) of the upper and lower skins (6, 7). The lightweight part (1) is covered by a metallic sheet (8), and has a “V”-shaped cross section with arms and peak, where the peak of the “V”-shaped cross section has a rounded shape and is located opposite the end section (13) of the upper and lower skins (6, 7).

Abstract: Composite aircraft frame (1) comprising two segments (10, 20), each of these segments (10, 20) being single unitary element made in composite material and comprising at least one web (2), a stiffening structure and a fitting (4), the two segments (10, 20) being joint by the fittings (4) using rivets, forming the aircraft frame (1). The invention also refers to a method for fabricating such an aircraft frame (1).

Abstract: Rear fuselage of an aircraft comprising a tail cone end and a rest of the real fuselage whereby the tail cone end is attached to the rest of the rear fuselage by means of an attachment system comprising a balancer fitting, three lugs and a waiting link. When there is a failure in one of the lugs, the balancer fitting acts supporting loads in the transversal direction of the aircraft. The waiting link only acts if a failure in other link occurs. This waiting link is able to support loads in the longitudinal direction of the aircraft.

Abstract: Protection of elements (1) of composite material protected at its points of complex geometry, such as various edges and corners, against electric discharges to which they are potentially exposed, said elements (1) being obtained by applying a cycle of temperature and pressure to a composite material arranged in a mold (2) designed in such a way that, once the geometry of the element (1) is known, the mold (2) includes free spaces (6) intended for the accumulation of resin from the process of manufacture of said element (1), and said resin will form the dielectric protection of the various edges or corners of the element (1) as it builds up, forming layers (3) that are disposed around the free edges of the element (1), thus endowing these surfaces of various edges or corners with dielectric protection of electrical insulation that prevents possible discharges and expulsion of hot particles in the case of electric discharge on said element (1).

Abstract: A repair process for a composite material panel that forms part of a fuselage, wings, or stabilizers of an aircraft, with the panel having large dimensions and including irregular contours in some areas and edges prone to damages occurring from handling and assembly of the panel. The process a) locates the damage in the element of the composite material panel; b) sands the area that includes the damage in an area larger than the damage, makes a cut in the panel, having the cut of a same form as a piece that will serve to repair panel; c) places the piece in the cut, so that it is perfectly flush with the panel to be repaired; and d) attaches the piece to the composite material panel.

Abstract: The invention relates to a method of assembling composite material parts (10, 12). The method makes provision for: mounting (S6) the parts for assembly under stress on a stand (14), which parts have been partially polymerized; heating (S7) these parts so as to polymerize them completely, thereby enabling the stresses in the parts to be relaxed and enabling them to match the shape of the sand; cooling (S8) the parts; and assembling (S9) in conventional manner the parts as cooled in this way.

Abstract: An aircraft tail surface, such as a horizontal tail plane or a vertical tail plane, including a leading edge having, in at least a section along the tail span, an undulated shape formed by a continuous series of smooth protrusions and recesses so that, in icing conditions, the ice accretion is produced only on the peaks of the protrusions and on the bottoms of the recesses, thereby creating a channeled airflow and an arrangement of air vortices which impart energy to the airflow in the aerofoil boundary layer which delay the airflow separation that causes the stall, thus reducing the detrimental effects of ice accretion in its aerodynamic performance.

Abstract: Weathertight fitting (41) for the mounting of the tail vertical stabilizer of an aircraft in a rear area of its fuselage, the construction being based on a casing (5) made with a composite material as one unitary piece and some frames (7) that are comprised of: one first piece (43) comprised of two bodies (45, 45?), that include the mounting lugs (47, 47?) for the tail vertical stabilizer and some vertical walls (49, 49?) for joining the fitting (43) to the frames (7) of the fuselage, and a weathertight central profile (51); two pairs of additional pieces (55, 55?; 65, 65?) of angular shape, that include some horizontal walls (57, 57?; 67, 67?) for joining to the casing (5) and some vertical walls (59, 59?; 69, 69?) for joining to the aforementioned bodies (45, 45?). All the weathertight fitting's (41) pieces are made with a composite material.

Abstract: A reinforcement structure for an opening in the primary structure of an aircraft, the structure comprising a skin, frame members which are transverse with respect to the flying direction of the aircraft, and stringers which are longitudinal with respect to the flying direction of the aircraft. The reinforcing structure includes a perimetral reinforcing element situated along the edge of the opening and reproducing the geometrical form thereof; at least one pair of transverse reinforcing elements arranged on both transverse sides of the opening; and at least one pair of longitudinal reinforcing elements arranged on both longitudinal sides of the opening.

Abstract: Its object is to allow measurement of the volume of fuel (2) using optic fiber cables (3); it is characterized in that, adhered to one or more of the inner faces of the tank (1) there is at least one sheet (4) in which a plurality of optic fiber cables (3) with different lengths are included, located between the inner face of the tank and the sheet (4). One set of the ends (10) of the optic fiber cables (3) are disposed so that they are in contact with the fuel (2) when it reaches their level, thereby allowing measurement of the fuel level at all times. The preferred embodiment distributes the optic fiber cables (3) on two opposing lateral faces of the tank (1) so that they cross over one another and in opposite directions to measure any fuel level. It is applied in aeronautics.

Abstract: The invention relates to a method for designing a composite material part with a curved surface (65) to which a plurality of stiffening elements will be joined, in which the design of a 3D fabric model comprises the following steps: generating a planar surface (63) developed in 2D of the surface (65) of the curved part in 3D as well as the same geometric references (71) in 2D as the references (51) of the surface in 3D; associating the surface (63) developed in 2D and the surface (65) in 3D as well as their geometric references (71, 51), such that any action in one of them is reflected in the other; initially computing the fabric model in the surface (63) developed in 2D; carrying out any subsequent modification of the fabric model either in the surface (63) developed in 2D or in the surface (65) in 3D.

Abstract: An aircraft pylon support system for mounting and supporting aircraft propulsion systems to a fuselage section of an aircraft that includes an internal, central support box disposed within and extending across the interior of an aircraft fuselage, and external support boxes extending laterally into pylons disposed at both sides of the aircraft, each of the internal and external support boxes being made of a composite material and structured as multi-spar boxes containing lateral spars and at least one central spar, wherein the central and lateral support boxes are connected together to provide a continuous interface therebetween with a full transfer of loads therebetween and between any interrupted, intermediate frame members.

Abstract: A device provides electrical continuity between aeronautical components including a first fixed component and a second component displaceable with respect to the first component. The first and second components include electrically conducting surfaces. A first conductive piece is fixed to the first component and a second conductive piece is fixed to the second component. The first and second pieces are joined together with the possibility of sliding one on the other, in such a way that the first piece is connected to the conductive surface of the first component and the second piece is connected to the conductive surface of the second component, so that, when displacements take place of the second component with respect to the first component, the first piece slides on the second piece maintaining the electrical conductivity between the first and second components.

Abstract: Joining arrangement for the lateral boxes (11, 11?) of a horizontal tail stabilizer with a central tubular box and manufacturing method for said box. The boxes (11, 11?) are joined using an intermediate joining part (31) that comprises a central box (33, 63, 73), upper (35, 35?) and lower (37, 37?) lateral flaps, fitting plates (A) for the trimming device and fitting plates (B) for the pivoting device, the join being made with rivets between the skins (15, 15?; 17, 17?) of the lateral boxes (11, 11?) and said flaps (35; 35?; 37; 37?). The method for manufacturing the intermediate joining part (31) comprises steps to: a) Provide preforms suitable for forming said part (31); b) Form and cure the part (31) made from said preforms using an RTM method.